$6,:RUNVKRS3LFFROH0LVVLRQL6FLHQWLILFKH 5RPD1RYHPEUH'LFHPEUH

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1 $6,:RUNVKRS3LFFROH0LVVLRQL6FLHQWLILFKH 5RPD1RYHPEUH'LFHPEUH

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17 GG Phase A Study: the Spacecraft Requirements and Design Criteria Satellite Configuration, Structure & Mechanisms Thermal Control Electrical Power & Avionics Attitude and Drag Free Control Budgets Development Approach & Plan Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

18 REQUIREMENTS AND DESIGN CRITERIA Design drivers: Circular equatorial orbit, 520 km, short life (six months after commissioning and set-up/calibration) One experiment running uninterrupted : few operational modes, small telemetry rate, no attitude manoeuvres Undemanding resources (250 kg, 110W) Suppression of external and internal disturbing accelerations and thermal effects constraints on mass distribution & area-to-mass ratio drag free control, spin control, thermal stability Design approach: Build satellite around experiment ad-hoc configuration Decouple satellite service functions (easy) from experiment (complex) standard satellite functions (OBDH, TT&C, Power, Propulsion, AOCS) easily adapted to PRIMA designs dedicated, highly integrated experiment control (experiment produces its own feedback signals, dedicated processor generates test mass position, drag free & spin rate control commands and drives actuators) Design for compatibility with wide range of small launch vehicles design exercise focused on Pegasus (small payload mass, small fairing, high launch loads) Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

19 SATELLITE CONFIGURATION Configuration drivers: fit reference launch vehicle fairing envelope outside shape must be a cylinder, diameter = 1 m low area-to-mass ratio easy integration of PGB spin axis must be a principal axis of inertia β= (J spin - J trans )/J trans 0.2 Proposed solution: ad-hoc spinning top structure supporting the PGB and equipment plus cylindrical solar panel in two pieces sensors and electric thrusters mounted to central belt, two S-band antennas aligned with spin axis, one deployable β = 0.23 cross sectional area = 1.3 m² area-to-mass ratio = m 2 /kg Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

20 Structural configuration with solar panels removed Configuration under the Pegasus fairing Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

21 Structure: Carbon-fibre primary structure for minimum thermal distortions, designed to: Pegasus specifications, with large margins (first frequency > 40 Hz) current mass budget (250 kg) plus 20% system margin 500-node finite element model built; dynamic analysis performed; local panel instabilities addressed Compact and stiff design. Normal modes at launch: 49 Hz (axial), 66 Hz (lateral) Total structural mass: 57 kg including 20% margin No problem from static circumferential (spin) load, even for solar cells STRUCTURE & MECHANISMS Mechanisms: PGB Launch-lock, one-shot S-band antenna deployment, one-shot Passive compensation of thermal expansion/contraction First axial mode eigenvector plot (longitudinal bounce ) Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

22 PASSIVE COMPENSATION OF THERMAL EXPANSION/CONTRACTION High CTE rod Thermal distortion of the spacecraft structure changes moment of inertia about spin axis, hence, at constant angular momentum, the spin rate. Effect can be compensated by suspending balance masses on rods that expand & contract in "counterphase" to spacecraft a) ω s Very low CTE rod Compensation mass Spacecraft structure Spacecraft spin axis System can consist of high-cte rod cantilevered (and thermally anchored) on the outer shell and connected to a parallel, low-cte rod supporting a balance mass on its free end (upper panel). Configuration constraints lead to short rod lengths and large balance masses; moreover, large spin centrifugal force suggests to modify the scheme into that shown in middle panel b) High CTE rod Very low CTE rod Compensation mass Spacecraft structure High CTE rod Yet another configuration consists in two rods in triangular configuration supporting the balance mass (lower panel). Angle α is constrained by considerations of friction angle and linearity. The rod s l is amplified by a factor of 1/sinα 6, allowing use of conventional Al alloy rods and small balance masses (on 4 positions around the cylinder) c) α Compensation mass Hinge (typ.) Spacecraft structure Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

23 SATELLITE MASS BUDGET ITEM MARGIN MASS (KG) Test masses 0 % 20.0 Pico Gravity Box (PGB) 3 % 62.0 Capacitors, Inch worms, Rods, etc. 30 % 5.0 PGB electronics 30 % 9.0 Payload Data Processor 20 % 9.4 TOTAL PAYLOAD Integrated Control System 20 % 12.0 TOTAL SATELLITE kg Radiofrequency 20 % % SYSTEM MARGIN 50.3 kg Electrical Power & Solar Array 20 % 25.8 GRAND TOTAL kg Attitude sensors 20 % 2.0 FEEP propulsion 20 % 15.4 Nitrogen propulsion 20 % 6.6 Harness 20 % 6.0 Thermal Control 20 % 9.0 Structure & mechanisms 20 % 63.5 TOTAL SVM Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

24 THERMAL CONTROL Driving requirement is temperature stability of the test masses: dt/dt < 0.2 C/day Requirement is met by high efficiency thermal insulation, and minimisation of thermal conduction and power dissipation in the payload compartment Multi stage thermal insulation is realised by covering with MLI blankets (effective ε = 0.01) the external surfaces and the solar array backside, the inner side of the main structure, and the outer and inner side of the PGB Because of the high insulation and thermal inertia, the time to steady state is very long if the initial temperature is far from the equilibrium temperature. For the design assessment, T = 10 C was assumed Spacecraft thermal configuration Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

25 THERMAL ANALYSIS Temperature drifts and stability were assessed by a detailed thermal mathematical model. Worst case thermal gradients occur at zero declination of the sun (equinox) T eq turns out to be -2.7 C, and temperature drift becomes < 0.2 C/day after about 25 days, if T 0 = T eq + 10 C Temperature oscillations at orbit frequency, due to eclipse transits, are < 0.01 C (PGB) and < C (test masses) Excellent temperature uniformity and stability is achieved Model confirms that spacecraft does not affect radiatively the PBG environment; thermal control of the payload and spacecraft modules is effectively decoupled Thermal control of the spacecraft elements is conventional; all equipment comply with design temperature limits. Maximum heater power 10 W (batteries). Estimated radiator area 0.32 m² + 20% (additional 50% margin available) Internal Test Mass PGB 4 3 Ext. Test Mass Time (days) Temperature drift to equilibrium Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

26 ELECTRICAL POWER Power demand: constant through sunlight (59') and eclipse (36') payload : 6 W FEEP: 60 W spacecraft : 46 W 185 W provided by GaAs solar array, sufficient for power provision & battery recharge sunlight-regulated bus is sufficient (28 V in sunlight, V in eclipse) SOLAR ARRAY SOLAR ARRAY S3R main S3R battery PCU BATTERY BDR 32 cells VR 3.5 DE telemetry lines to RTU telecommand lines from RTU electronic switch + overcurrent limiter + telemetry signal Standby Payload (early phases) RELEASE MECHANISM PPDU Transmitter Receiver IAC Load 4 Load N small Ni-Cd battery (32 cells, 3.5 Ah) All regulation, protection and distribution functions can be unified in a single box (PPDU, Power Protection and Distribution Unit) Power subsystem block diagram Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

27 Highly integrated payload data processing tasks are executed in a dedicated payload processor. Spacecraft data management is perfectly suited for the Integrated Control System architecture planned for the PRIMA bus The Spacecraft Control Unit handles the ground link and performs data management, fault detection and recovery, and pre-operational attitude control. Data are collected and commands distributed through a 1553 bus, and a Remote Unit is used for interfacing with attitude and other equipment ON BOARD DATA HANDLING AND TELECOMMUNICATIONS Payload Processor 1553 bus Total data collection rate is 16 Mbit/orbit and an on-board mass memory sized for 24h autonomy would amount to 240 Mbit With an equatorial ground station (Malindi), data can be transmitted to ground once per orbit at 26 kbps, or even once per day at 0.5 Mbps The TT&C system is a standard architecture with one coherent transponder, two S-band antennas and a RDFU switch. The RF transmit power is < 1 W MMU RU to thrusters to PPDU Transponder Formatter Decoder Sun & Earth sensors SCU to antennas Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

28 Pre-operational requirements : remove initial precession & nutation point the spin axis within 1 of orbit normal spin the satellite up to 120 rpm ATTITUDE MEASUREMENT AND CONTROL Pre-operational actuators consist of four 20mN Nitrogen thrusters; 1.8 kg propellant required for rate damping (0.8 h) and spin-up (5.5 hours) Operational requirements: control phase lag between S/C and PGB to < 10-2 rad measure spin rate with RMS( ω/ω) 10-4, equivalent to a phase lag of (drag free control) measure absolute orientation of spin axis and Earth direction with θ < rad (0.35 ) Momentum bias of 400 Nms provides gyroscopic stability against torques Earth sensor provides sufficient measurement accuracy (bias 0.05, 3σ random error 0.01 ) Differential spin rate control based on passive compensation mechanism (90%) and FEEP (residual 10%). Spin controller works on light-emitting diode signals (10 µm resolution over 30 cm = 43 arcseconds) Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

29 DRAG FREE CONTROL Layout of the pre-operational thrusters The GG Drag Free Control (DFC) has to reduce the common mode drag force acting on the satellite by a factor of about 50,000 in a narrow bandwidth centred at the orbital frequency Drag model used in the simulations (in the inertial frame): amplitude F = N plus noisy force with amplitude equal to 10% of the mean force, at the orbital frequency f = 1/ s = Hz no second harmonic component for simplicity (it does not add any more physics) Simplified model of the drag used in the simulations: Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

30 drag force along the X axis DRAG FREE CONTROL (2) The DFC designed for GG is a notch filter at the orbital frequency, with narrow bandwidth, so that it acts in a frequency band where there is no passive attenuation due to the PGB Theoretical relative displacement between the PGB and the spacecraft, without (continuous black curve) and with (dashed blue curve) the application of drag free control. m S/C = kg, m PGB = 63.6 kg, f S = 2 Hz, Q SPRING = 90 Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

31 DRAG FREE CONTROL (3) Reducing the drag by a factor of 50,000 at the orbit frequency means reducing by the same factor the amplitude of oscillation at the orbit frequency of the spacecraft + PGB In 7 orbits only the control is able to reduce drag effect from 1 mm to less than 10-8 m. RMS of the reading capacitance sensors: 10-6 cm, Bias reading capacitance sensors: 10 micron, Angular bias of the reading capacitance sensors: 1 deg. Added white noise to the Drag Force = 10% of its amplitude Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

32 FIELD EMISSION ELECTRIC PROPULSION (FEEP) FEEP is an electrostatic propulsion based on field ionisation of a liquid metal (usually Cesium) and acceleration of the ions by a strong electric field FEEP has unique features, ideal for dragfree control applications: thrust range 1 µn to 1 mn near instantaneous switch on / switch off capability high resolution throttleability (better than 1 part in 10 4 ) accurate thrust modulation in both continuous and pulsed modes Because of the high Isp (4000 to 10,000 s), propellant flow rate is very low and the propellant reservoir is integrated with the thruster FEEP thrusters are developed by CENTROSPAZIO and the power control electronics by LABEN ω s φ=60 ο )5217Ã 9,(: φ=60 ο 723Ã 9,(: ω s φ=60 ο φ=60 ο /$7(5$/Ã9,(: ω s Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

33 TECHNOLOGY ASSESSMENT & EXPERIMENT VERIFICATION Stringent scientific requirements can be met by existing technologies Payload: Experimental test and breadboarding needed for electrical / signal wire routing, lock-unlock mechanisms, capacitance read-out and active damper control electronics Inch-worm mechanisms exist for use in ultra-vacuum and high-radiation environments; they need to be space-qualified Suspensions need to be characterised and tested for space-like environmental conditions GGG ground experiment relevant to all above subjects FEEP: thruster prototypes developed since many years at CENTROSPAZIO thruster control electronics under development at LABEN thruster endurance test and characterisation shortly to be started at ESTEC flight demonstration planned in early 2000 on a Get Away Special canister on the Space Shuttle (Payload No. G-752) Drag free control: real-life system cannot be tested on the ground. Verification by software simulator, incorporating test data of key elements (sensors, dampers, FEEP) Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

34 SATELLITE MODEL PHILOSOPHY & VERIFICATION APPROACH Equipment and Subsystems: AMCS test bench for checking system / payload interfaces (attitude and drag free control) Standard subsystems verified as part of the system test flow EQUIPMENT AMCS DM on PRIMA EM (testbench) Commercial components Functional/ electrical testing FM or PFM Flight standard Acceptance testing Payload Module : Development Model for functional and performance verification Structural / Thermal Model Protoflight Model PAYLOAD DM (Breadboard) Commercial components Functional/ electrical testing STM Thermal/ mechanical testing PFM Flight standard Acceptance and protoflight testing System: Verification completed on Protoflight Model (PFM), the only model manufactured and assembled at flight standard SYSTEM PFM Protoflight testing Ground Support Equipment one set reused from bottom to system level LEGENDA: EM : Engineering Model DM : Development Model STM : Structural/Thermal Model PFM : Proto-Flight Model PROGRAMME SCHEDULE Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

35 ID Task Name Duration Start Finish 1 System Phase B 191d Qtr 1 Qtr 2 Qtr 3 Qtr 4 Qtr 1 Qtr 2 Qtr 3 Qtr 4 Qtr 1 Qtr 2 Qtr 3 Qtr 4 Qtr 1 Qtr 2 Qtr 3 2 Kick Off 1d Preliminary Design Review (PDR) 1d GG Laben Phase B 191d System Phase C/D 580d Kick Off 1d Critical Design Review (CDR) 1d Qualification Results Review 1d FAR 1d PFM Equipments Manufacturing & Test 201d S/C PFM Structure Mn. & Int. 132d PFM Units Delivery To Alenia 0d S/C PFM Equipm.s Int. & Tests 50d PGB PFM Delivery To Alenia 1d PGB PFM Int. & Tests 5d S/L PFM Funct. & Environm. Tests 149d Workshop Nazionale ASI Piccole Missioni Scientifiche - Roma 30 Nov. - 1 Dic

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