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NGGM ASSESSMENT STUDY Progress Meeting 2 TUD, Delft, 24-25 March 2010

NGGM ASSESSMENT STUDY Progress Meeting 2 TUD, Delft, 24-25 March 2010. Agenda. 10.15 Introduction, agenda 10.30 Task 2: Observing techniques - part 1 WP 2120 Instrument Concepts TAS-I WP 2121 Measurement Technologies ONERA

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NGGM ASSESSMENT STUDY Progress Meeting 2 TUD, Delft, 24-25 March 2010

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  1. NGGM ASSESSMENT STUDYProgress Meeting 2TUD, Delft, 24-25 March 2010

  2. Agenda • 10.15 Introduction, agenda • 10.30 Task 2: Observing techniques - part 1 • WP 2120 Instrument Concepts TAS-I • WP 2121 Measurement Technologies ONERA • 11.30 Task 3: Mission analysis / attitude and orbit control concepts - part 1 • WP 2220 Attitude and Orbit Control Concepts TAS-I • 12.00 Lunch break • 13.00 Task 3: Mission analysis / attitude and orbit control concepts - part 2 • WP 2210 Mission Analysis DEIMOS • 13.45 Task 5: Mission Architecture Outlines • WP 2420: Mission Architecture GIS • WP 2410: Architecture Definition and Trade-Off TAS-I • 15.15 Task 1: Mission Requirements • WP 1100 Requirements Analysis (update) Uni Lux. • 15.45 Task 2: Observing techniques - part 2 • WP 2110 Observing Techniques IAPG • 16.45 Task 4: Simulation Tool • WP 2310 End-to-End Simulator Design and Implementation TAS-I • WP 2320 Variable Gravity Model IAPG • WP 2330 Backward Module DEOS • Discussion • 18.30 End

  3. WP 2120 Instrument Concepts (TAS-I)

  4. Contents • Outline of an optical metrology concept option without beam steering mechanism. • Update of the top-level requirements and lower level specifications on satellite-to-satellite distance measurement and non-gravitational acceleration measurement for a reference scenario with inter-satellite distance = 75 km and laser pointing/tracking function performed by the satellite. • Outline of P/L configuration options for the optical metrology without BSM and different arrangements of the accelerometers. • Assessment of the implications on the laser interferometer of cartwheel and pendulum formations with inter-satellite mean distance = 75 km.

  5. Optical metrology concept w/o BSM • Main differences with respect to the original metrology concept: Retro-reflector (RR) • Laser interferometer concept: unchanged. • Beam steering mechanism on Satellite 1: removed (laser beam pointing and control in charge of the attitude control of Satellite 1). • Angle/lateral displacement metrology on Satellite 2: maintained, but moved externally to the satellite for increasing the baseline between the 3 telescopes (to keep nearly the same baseline/laser beam size ratio at 75 km distance). • Angular metrology on satellite1: removed.

  6. Optical metrology concept w/o BSM Measurement of the “scientific observables”. • Satellite to satellite (RR-to-RR) distance variation: laser interferometer. • Non-gravitational accelerations of Satellite 1, Satellite 2: accelerometer set. • Orientation of Satellite 2 in Satellite-to-Satellite Reference Frame (needed for transferring the laser interferometer measurements from the RR to the COM and for projecting the acceleration measurements along the satellite-to-satellite line): angular-lateral metrology on Satellite 2. • Orientation of Satellite 1 SSRF: angular-lateral metrology on Satellite 2 (through the measurement of the lateral displacement of the Satellite 2 from the laser beam axis, aligned to X1).

  7. Top-level requirements update • Requirements established on the basis of the best performance on the satellite-to-satellite distance and non-gravitational acceleration measurement ( maximisation of the scientific return) that is believed achievable in orbit with state-of-the art technology either flight proven or laboratory proven ( minimization of the development risks and costs).

  8. Error tree for Satellite-Satellite distance • The limit of 20 nm/Hz is set by the laser frequency stability. • The suppression of the Beam Steering Mechanism leads to the suppression of two important contributors to the error budget  The re-distribution of their error portions allows to relax in particular the requirements on the laser beam pointing and on the satellite orientation relative to the satellite-satellite line. • The requirements linked to the thermo-structural stability are kept unchanged.

  9. Error tree for non-gravitational acceleration • The requirements on the satellite orientation relative to the satellite-satellite line have been aligned to those coming from the satellite to satellite distance measurement. The other requirements have been kept unchanged from d = 10 km.

  10. Requirements on optical metrology • Laser frequency stability requirement [Hz/Hz] • Interferometer measurement noise requirement [m/Hz] • Optical power on the interferometer sensor (pd2) required for 10 nm/Hz: ~ 1pW • Optical power on pd2, for = 26 mm RR and 0.5 W laser emitted power: margin 10 up to 100 km.

  11. Requirements on accelerometer Applicable to an accelerometer pair 1/f2 • The noise floor of 310-12 m/s2 along the X-axis (nominally aligned to the satellite-to-satellite line) is consistent with the predicted noise of the ultra-sensitive axis of the GOCE accelerometer. • The noise floor of 10-10 m/s2 along the Y, Z-axis is close (factor ~2) to the predicted noise of the less-sensitive axis of the GOCE accelerometer. • The 1/f2 limit to the accelerometer noise a low frequency has been inferred from the spectral density of trace of the gravity gradient measured by GOCE over a time period of 60 days.

  12. Requirements on accelerometer Applicable to an accelerometer pair • Accelerometer bias: 210-7 m/s2 (all axes) • Common and differential scale factors knowledge (from calibration): 210-4 for X axis  110-3 for Y, Z axes The requirement of the common SF is more stringent than in GOCE (210-3) where the differential accelerations and not the common-mode ones are the main observables. • Accelerometer scale factor stability (all axes):

  13. Requirements on satellite pointing control • Satellite X-axis alignment to the satellite-to-satellite direction: Satellite 2: 1 (thanks to the use of a retro-reflector) Satellite 1:  210-5 rad (driven by the laser beam pointing) • Satellite X-axis pointing stability relative to the satellite-to-satellite direction: Merging of the needs of the satellite-to-satellite distance measurement and of the non-gravitational acceleration measurement Driven by the laser beam pointing

  14. Requirements on satellite attitude measurement • Satellite X-axis alignment to the satellite-to-satellite direction (knowledge): Satellite 2: 10-4 rad Satellite 1: <210-5 rad (driven by laser beam pointing control) • Satellite X-axis pointing measurement error spectral density: • Optical power on the angle metrology sensor (PSD) required for 2 µrad/Hz: ~ 70 nW. Always fulfilled up to 100 km with  = 10 mm telescopes. Measured by the angle metrology for the Satellite 2 by the lateral displacement metrology for the Satellite 1.

  15. Requirements on accelerations control Requirement on linear acceleration control, all axes (drag-free requirement) Requirement on angular acceleration and angular rate control, all axes.

  16. Payload configuration concepts Ultra sensitive axis (US) Retro-reflector • Original configuration concept 1: Two accelerometers. Linear accelerations along X, Y and angular acceleration around Z measured with the US. Linear acceleration along Z and angular accelerations around X, Y measured with LS axes only. Measurement of the YY component (cross track) of the gravity gradient. Less sensitive axis (~0.01US) 367 mm Interferometer core Z BSM Angular metrology (operative on Satellite 1) 772 mm 540 mm Angular-lateral metrology (operative on Satellite 2) Interferometer telescope X Y

  17. Payload configuration concepts Ultra sensitive axis (US) Less sensitive axis (~0.01US) • Configuration concept 1: Linear and angular accelerations about X, Y, Z measured with the ultrasensitive axes. No ultra sensitive axis aligned to the accelerometer-to-accelerometer baseline  no gradiometry service provided. Z 674 mm 300 mm Y X

  18. Payload configuration concepts Z • Configuration concept 2: Linear acceleration along Z measured with less sensitive axes only. Measurement of the YY component (cross track) of the gravity gradient. 505 mm Y 670 mm X

  19. Payload configuration concepts 670 mm • Configuration concept 3: Angular acceleration around X measured with less sensitive axes only. Measurement of the YY, ZZ components (cross track, radial) of the gravity gradient. Z 505 mm Y X

  20. Payload configuration concepts 670 mm Z • Configuration concept 4: Angular acceleration around Y measured with less sensitive axes only. Measurement of the YY, ZZ components (cross track, radial) of the gravity gradient. Y X

  21. Payload configuration concepts • Payload configuration concept selection logic Need of gravity gradient measurement? YES NO Trade-off between concepts 2, 3, 4 (criteria: performance on linear/angular acceleration measurement, failure tolerance) Concept 1 optimized for size and mass (trade-off 4 vs 3 accelerometers)

  22. Cartwheel formation impact on metrology Relative distance = [50 100] km (75 mean) Maximum angular variation of the line joining the two satellites: 19.47° Maximum Doppler shift of the laser frequency: 108.12 MHz

  23. Pendulum formation impact on metrology Relative distance = [62 88] km (75 mean) Maximum angular variation of the line joining the two satellites: 45° Maximum Doppler shift of the laser frequency: 55.65 MHz

  24. WP 2220 Attitude and Orbit Control Concepts (TAS-I)

  25. Activities since PM1 • Performed activities since PM1 • Formation flying control algorithm re-tuning and update (described in the following) • Assessment on the possibility to substitute the laser beam pointing mechanism with the satellite attitude tracking (described in the following). • Generation of additional attitude noise as for on-ground attitude reconstitution (data fusion between star-tracker and accelerometers’ data). • Support for system level analysis (e.g. fuel budget, presented elsewhere)

  26. FF algorithm update Extension of formation flying control algorithm for in-line formation (named also GRACE, trailing) with longer baseline. • As anticipated in PM1, the already available algorithms have been considered for a longer baseline (75km instead of 10km considered in previous study). • Problems have been found for what concerns the residual acceleration spectral density and magnitude (parameter re-tuning was not enough). • It was decided to update the algorithm in order to support the performance analysis. The new control algorithms (to be used after an acquisition phase, that may be based on what has been developed in previous study phase) do not introduce any control action on across and radial direction. The along-track control action is at very low frequency only. • Preliminary results for in-line formation are provided in the following.

  27. FF algorithm update • Preliminary results for in-line formation (1/3) • Satellites’ relative position.

  28. FF algorithm update • Preliminary results for in-line formation (2/3) • Residual accelerations on Satellite 1

  29. FF algorithm update • Preliminary results for in-line formation (3/3) • One-sided spectral density of the residual acceleration.

  30. Laser beam pointing by satellite AOCS • Requirement for attitude control on YZ axes (requirement for S1)

  31. Laser beam pointing by satellite AOCS • From above analysis it is possible to observe: • the effect of satellite relative movement (S2-S1) is negligible with respect to the attitude jitter naturally induced by environment; • the requirement on S1 attitude error due to the laser pointing stability is more stringent (black bold line) than the requirement on attitude error due to LORF tracking (black line). • At the end, it is possible to conclude that the tracking of S2 by S1 is compatible with the angular accelerations/angular rate/attitude requirements of LORF tracking. • Impacts on thruster’s parameters: • in-line formation: the already provided micro-thruster requirements are still enough. • pendulum formation: specific analyses have not been yet done. In any case large impacts are expected related to the angular variation joining the two satellites, and the lateral cross-section. Impacts will be present on micro-thruster and mini-thruster requirements.

  32. Laser beam pointing by satellite AOCS Preliminary performance requirements applicable to the ion thrusters for an NGGM

  33. On-ground attitude reconstitution emulator • In order to support the refined analysis based on E2E simulator, a model of the on-ground attitude reconstitution errors has been developed. • This function, named on-ground attitude reconstitution emulator, builds the attitude errors as results of the data fusion between star-tracker and accelerometers measurements. • To provide a realistic scenario, a new mathematical model of star-tracker errors has been developed starting from GOCE in-flight data. The model takes into account bias, random and harmonic errors.

  34. On-ground attitude reconstitution emulator • Comparison of star-tracker error spectral-densities. • From flight data From developed model

  35. On-ground attitude reconstitution emulator • Example of the attitude reconstitution error spectral density

  36. WP 2410 Architecture Definition and Trade-Off (TAS-I)

  37. Spacecraft volume limits from launcher envelope Assumption: Rockot

  38. Spacecraft volume limits from launcher envelope

  39. Spacecraft volume limits from launcher envelope

  40. SSO Body mounted & wing mounted solar panels (8.5 m2 total GaAs cell area) Power budget - GOCE

  41. GOCE ion thruster power consumption • NGGM reference for orbit control requirements

  42. Mini-RIT thruster power consumption • NGGM reference for attitude control requirements (max thrust level  500 μN)

  43. Simulation of 316 km circular polar orbit Configuration as currently implemented in the simulator MSIS density model Very high solar flux F10b=200, F10b(mean)=200, Ap=20.2 Thruster mass and power requirements

  44. Mini-RIT thruster thrust and power

  45. Mini-RIT specific impulse and mass flow rate

  46. Orbit control thrust and power

  47. Ion thruster mass and power requirements • Propellant mass needs • Estimated from average thrust • Xenon mass AOCS: 3.0 kg/year • Xenon mass orbit: 4.2 kg/year • 72 kg for 10 years • Electrical power needs - AOCS • 8 thrusters, 400 μN peak thrust demand per thruster • 300 W peak • Electrical power needs - orbit • 1 GOCE thruster @ 6.5 mN orbital peak : 320 W (MSIS, F10.7 = 200) • 1 GOCE thruster @ 3.5 mN peak : 220 W (J-B) • Spacecraft power budget •  500W platform + 300W (AOCS) + 200W÷300W (orbit)  1 kW

  48. NGGM solar array configuration model 5 1 2 3 7 6 4

  49. In polar orbit, the sun rotates around the orbit plane “summer”: sun normal to orbit plane “winter”: sun in orbit plane Option 1, seasonal change of attitude: pos. 1 in “summer”, pos. 2 in “winter” Large power excursion – factor of 2.5 - from “summer” to “winter” Worst case is the driver anyway  waste of power in “summer” Position relative to Earth of star sensors, radiators, GPS, changes too Option 2: keep position 2 throughout the year Smaller power excursion – factor of 1.2 Power always near worst case Power provision can be improved by wings with seasonally variable attitude Solar array configuration in polar orbit Position 1: sun normal to orbit plane Position 2: sun in orbit plane

  50. Effective solar array area – option 1 Velocity Earth Winter position Velocity Earth Summer position

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