1 / 58

NGGM ASSESSMENT STUDY Progress Meeting 1 TAS-I, Torino, 27-28 January 2010

NGGM ASSESSMENT STUDY Progress Meeting 1 TAS-I, Torino, 27-28 January 2010. Agenda – Day 1. Day 1 - 27-Jan-2010 14.00 Introduction, agenda 14.15 Task 5: Mission Architectures and Task 2: Observing techniques WP 2420 Mission Architectures DEOS [GIS] WP 2110 Observing Techniques DEOS [IAPG]

gerryk
Télécharger la présentation

NGGM ASSESSMENT STUDY Progress Meeting 1 TAS-I, Torino, 27-28 January 2010

An Image/Link below is provided (as is) to download presentation Download Policy: Content on the Website is provided to you AS IS for your information and personal use and may not be sold / licensed / shared on other websites without getting consent from its author. Content is provided to you AS IS for your information and personal use only. Download presentation by click this link. While downloading, if for some reason you are not able to download a presentation, the publisher may have deleted the file from their server. During download, if you can't get a presentation, the file might be deleted by the publisher.

E N D

Presentation Transcript

  1. NGGM ASSESSMENT STUDYProgress Meeting 1TAS-I, Torino, 27-28 January 2010

  2. Agenda – Day 1 Day 1 - 27-Jan-2010 14.00 Introduction, agenda 14.15 Task 5: Mission Architectures and Task 2: Observing techniques WP 2420 Mission Architectures DEOS [GIS] WP 2110 Observing Techniques DEOS [IAPG] WP 2120 Instrument Concepts TAS-I WP 2121 Measurement Technologies ONERA 17.00 Discussion 17.30 End

  3. Agenda – Day 2 • Day 2 - 28-Jan-2010 • 9.00 Task 3: Mission analysis and attitude and orbit control concepts • WP 2210 Mission Analysis DEIMOS • WP 2220 Attitude and Orbit Control Concepts TAS-I • 10.30 Task 4: Simulation Tool • WP 2310 End-to-END Simulator Design and Implementation TAS-I • WP 2320 Variable Gravity Model DEOS [IAPG] • WP 2330 Backward Module DEOS • 12.00 Task 5: Mission Architecture Outlines / Discussion and Conclusions • WP 2410 Architecture Definition and Trade-Off TAS-I • Discussion and Conclusions • 13:00 End

  4. WP 2120 Instrument Concepts (TAS-I)

  5. Distance metrology for increased satellite separation • Top-level requirement for satellite-to-satellite distance d = 100 km (preferred distance resulting from the parametric analyses with the quick-look tool). • The distance measurement relative error is maintained as for the 10-km case: d/d = 510-13 1/Hz. • The limit is set by the achievable laser frequency stability.

  6. Distance metrology for increased satellite separation • Top-level requirement breakdown for satellite-to-satellite distance d = 100 km. • Same relative weight of the various error terms maintained as for the 10-km case. All requirements are relaxed by a factor 10, but less optical power is received by the sensorsdegraded distance, angle measurement performance.

  7. Distance metrology for increased satellite separation • Optical power on the metrology sensors (computed with a laser source output = 0.75 W)vs requirements. S1 S2 • The original design of the laser interferometer and of the angle/lateral metrology on S2 is still applicable to d = 100 km. For the S1 orientation measurement relative to laser beam, a different concept has to be devised.

  8. Distance metrology for increased satellite separation • Possible concepts for measuring the S1 orientation relative to laser beam (S1-S2 line), to be assessed. • Concept 1: The position of S2 is measured in the S1 LORF by the GPS (’1). The orientation of S1 is measured in the S1 LORF by the STR (”1)  1 = ’1+”1. • Concept 2: Utilization of a high-resolution camera pointing in the direction of S2 and equipped with CCD or APS detector (less optical power required, but higher sensitivity to straylight and background). • Concept 3: Utilization of a quadrant photodiode (qpd) in the laser interferometer. Applicable only in case theBSM is removed (not impossible with 10x relaxed requirements).

  9. Acceleration measurement with increased sat separation • Top-level requirement for non-gravitational acceleration measurement kept unchanged at any satellite-to-satellite distance in the parametric analyses with the quick-look tool (noise floor = 10-11 m/s2/Hz). A slight relaxation (1.110-11 m/s2/Hz) is however necessary for d = 100 km in order to relax the requirements on the S/C orientation knowledge relative to the S1-S2 line to the same level of those derived for the distance measurement (1.510-6 rad/Hz). Relaxed top-level requirement Original top-level requirement

  10. Acceleration measurement with increased sat separation • Top-level requirement breakdown for satellite-to-satellite distance d = 100 km. Alternatively, the top-level requirement can remain unchanged by slightly tightening the drag-free requirements.

  11. Acceleration measurement with increased sat separation • Modifications of the current accelerometer layout will be implemented if: • The analyses/simulations will proof the benefit of adding to the satellite-to-satellite tracking the measurement of one or more gravity gradient components. • It is beneficial for the measurement of the linear and angular accelerations.

  12. STR noise from GOCE in-flight results • Star-tracker noise from GOCE in-flight results. • In the frame of the study, a preliminary processing of GOCE star-tracker measurements has been performed to assess the actual noise spectral density. • As it will be seen, the low-frequency noise spectral density is higher than the high-frequency one. • The spectral densities have been estimated considering: • uncorrelated the error on each axis; • equal the power spectral density on X and Y axes (Z axis is the boresight).

  13. STR noise from GOCE in-flight results • The attitude error power spectral densities SXY and SZ have been estimated considering two sets of data: • STR1 and STR2 (172799 samples at 2Hz); • STR1 and STR3 (79199 samples at 2Hz). • SXY and SZ densities have been determined comparing the inertial attitude measured by each star-tracker, after removing the mean values of the relative attitude: i, j index of the star-tracker

  14. STR noise from GOCE in-flight results • One-sided spectral density from star-tracker comparison • STR1-STR2 (Nov 3rd, 2009) STR1-STR3 (Nov 10th, 2009)

  15. STR noise from GOCE in-flight results • The results obtained considering the two sets of data are consistent. • The low-frequency noise spectral density is about 10 times greater than the high-frequency one. It is due to FOV errors like: • residual from focal plane calibration; • uncertainty on star colour; • residual from differential aberration compensation; • S-shape; • star-tracker thermo-elastic deformation. • Above errors become random bias for inertial pointing. If during the in-orbit movement the star-tracker images the same sky regions, they become harmonic errors. • For orbit plane GOCE like (about 16 orbit/day and 1 deg/day precession), the above errors may be considered as harmonic on time span of several orbit periods. • The correct noise shape shall be considered for attitude control design, post-facto attitude and angular rate reconstitution.

  16. WP 2210 Mission Analysis (DEIMOS)

  17. Mission Analysis Presentation Contents General Considerations about Repeating Orbits Repeat Cycle Temporal/Spatial Sampling Analysis tools: coverage matrix, gap evolution graph Low resonances avoidance

  18. Mission Analysis Repeating Orbits “M orbits in D days” The satellite flies M orbits Earth makes D revolutions w.r.t. the orbital plane But the line of nodes drifts: Polar : no drift D revolutions = D sidereal days Sun Synchronous D revolutions = D days Otherinclinations no precise correspondence I+N/D formulation M = I*D+N: number of orbits in 1 RC D: repeat cycle duration (in “days”) I+N/D: number of orbits per “day” N/D: useful for quick identification of sub-cycles

  19. Mission Analysis Repeating Orbits [250 km – 400 km] altitude range RC < 30 days Sun Synchronous

  20. Mission Analysis Repeating Orbits [250 km – 400 km] altitude range RC < 30 days Polar

  21. Mission Analysis Repeating Orbits [250 km – 400 km] altitude range RC < 30 days i = 62.7°

  22. Mission Analysis Repeating Orbits Spatial Sampling After 1 day: I+N/D orbits Fundamental interval: S = 360°/(I+N/D) The lower the orbit, the thinner S Typical value at the Equator for alt = 300km, S ≈ 2500 km After 1 repeat cycle: M = I*D+N orbits Sampling interval: Si = 360°/M = S/D The longer the RC, the thinner Si Typical value at the Equator for alt = 300 km, 30-d RC, Si ≈ 80 km

  23. Mission Analysis Repeating Orbits Spatial/Temporal Sampling The coverage pattern depends on N/D ratio 2 representations: Coverage matrix Day-by-day filling pattern of the fundamental interval E.g. 15+16/17 FundamentalInterval

  24. Mission Analysis Repeating Orbits Spatial/Temporal Sampling The coverage pattern depends on N/D ratio 2 representations: Coverage matrix Gap evolution graph Max/min gap at Equator after n days E.g. 15+16/17

  25. Repeating Orbits Spatial/Temporal Sampling The coverage pattern depends on N/D ratio 2 representations: Coverage matrix Gap evolution graph When N = 1 or N = D-1 Drifting orbit Progressive filling Slow max. gap evolution E.g. 15+16/17 Mission Analysis

  26. Repeating Orbits Spatial/Temporal Sampling The coverage pattern depends on N/D ratio 2 representations: Coverage matrix Gap evolution graph Other cases Skipping orbit Possible sub-cycles E.g. 15+14/17 Mission Analysis

  27. Repeating Orbits Spatial/Temporal Sampling The coverage pattern depends on N/D ratio Sub-cycles N/D ≈ N’/D’ where D’<D E.g. 15+20/29 20/29 ≈ 9/13 ≈ 2/3 13-day and 3-daysub-cycles Sub-cycles allow combining Large RCs: fine sampling Homogeneous coverageafter few days like withlow RCs Mission Analysis Sub-cycles

  28. Low RC Resonances Avoidance GRACE lesson learnt: M > 2 L where L is the maximum degree of L x L solution For L = 120, D must be greater than 15 Distance between ≤ 15-dayrepeating orbits Drives the altitude controlmargin Mission Analysis

  29. Low RC Resonances Avoidance GRACE lesson learnt: M > 2 L where L is the maximum degree of L x L solution For L = 120, D must be greater than 15 Distance between ≤ 15-dayrepeating orbits Drives the altitude controlmargin Example with all 29-dayrepeating SSOs Mission Analysis

  30. Low RC Resonances Avoidance GRACE lesson learnt: M > 2 L where L is the maximum degree of L x L solution For L = 120, D must be greater than 15 Distance between ≤ 15-dayrepeating orbits Drives the altitude controlmargin Example of preliminarycandidate 15+4/5, polar 11-d orbit at -5 km 14-d orbit at +4 km 29-d orbit at +2 km 26-d orbit at -2 km Mission Analysis

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

  32. Activities since RRM • Performed and on-going activities since Requirements Review Meeting • Re-tuning of the drag-free control algorithms to take into account the anti-aliasing filter on accelerometer’s outputs. • Check on the possibility to extend the current control design for in-line formation (named also GRACE, trailing) with longer baseline. • Check on the feasibility to perform the laser beam pointing by satellite, instead of using the beam steering mechanism. • Update of the requirements for the accelerometer measurement noise profile. • Requirement specification for stochastic noise/error and harmonic noise/error.

  33. Drag-free tuning update Re-tuning of the drag-free control algorithms to take into account the anti-aliasing filter on accelerometer’s outputs. • This activities has been necessary since the previous drag-free control tuning has not taken into account of the anti-aliasing filter present on the DFAC channel of each accelerometer. • The anti-aliasing filter is the same considered for the DFAC channel present on the GOCE electrostatic gradiometer, i.e. 3rd order Butterworthlow-pass digital filter, 3.5 Hz cutoff frequency. • The re-tuning has been done taking into account all the requirements, providing enough robustness toward plant uncertainty, and still improvement capabilities. • For now, the activity is considered closed.

  34. Drag-free tuning update • Spectral density of the linear X-axis acceleration before tuning • Damping too low!

  35. Drag-free tuning update • Linear X-axis acceleration after tuning (drag-free, no FF) • Good damping!

  36. Extension to longer baseline Check on the possibility to extend the current control design for in-line formation (named also GRACE, trailing) with longer baseline. • It has been verified that the actual formation flying control algorithms are able to manage in-line formation with longer baseline, by simple set-point update and re-tuning. • 100km baseline has been checked. Results show the necessity to re-tune the control parameter to cope with the higher perturbation (gravity gradient acceleration). • The requirement R on the relative position shall be updated according to the baseline B since they come from the laser beam, i.e.R(B)=R(10km)*B/10, B in km • The activity for the formation controlre-tuning is in progress (for 100km baseline).

  37. Extension to longer baseline • Results from 100km baseline without any re-tuning. • To be reduced • Earth gravity field has been limited to order 20 to speed up the simulation.

  38. Laser beam pointing by satellite AOCS Check on the feasibility to perform the laser beam pointing by satellite, instead of using the beam steering mechanism. • A laser beam pointing controller has been designed during previous study phases. It has been inserted mainly to overcome the reaction-wheel mechanical noise effects on laser beam pointing stability. • The attitude control shall constrain the angular accelerations and the angular rates to be very stable (< 10-8 rad/s²/Hz and < 10-6 rad/s/Hz respectively in [0.001,0.01]Hz band). The coupling of angular acceleration and the angular rate with the accelerometer displacement from the satellite COM, produces a linear acceleration which contributes to the measurement error of the non-gravitational accelerations of the satellite COM. • Above requirements impact on the attitude control closed loop and on the reference attitude trajectory to be tracked.

  39. Laser beam pointing by satellite AOCS • Is the tracking of S2 by S1 compatible with the angular accelerations/angular rate/attitude requirements? • The answer has been provided according to the following steps: • derivation of the real driver requirement starting from the provided angular requirements. This is necessary because the angular requirements have been derived without taking into account the kinematics constraint. • comparison of driver requirement with what it is induced by relative satellite movement (across-track and radial). • The tracking of S2 induces on S1 an angular jitter reduced by a factor equal to the baseline. S2 attitude jitter does not have any impact. S1 disturbing torque (environment, thrusters, etc.) shall be reduced to the level of laser pointing stability requirement.

  40. Laser beam pointing by satellite AOCS • Driver requirement for attitude control on YZ axes (requirement for S2) • For frequency higher than 5mHz, the driver requirement is the angular acceleration. For frequency below 5mHz, the driver requirement is the attitude requirement.

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

  42. 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 maximum force and force slew-rate is expected due to higher control bandwidth and lower noise level. The work is in progress.

  43. Accelerometer noise requirement • Preliminary definition of the accelerometer noise profile requirement for DFAC channel. • The preliminary acceleration configuration is • In principle, two output channels may be considered for each accelerometer (as for GOCE): a) science channel and b) DFAC channel (considered below). • Linear accelerations are directly measured by accelerometers. • X and Z axes angular accelerations are measured as differences of linear accelerations (Z and X respectively) and scaled by accelerometers distance (L=0.1835m). • Y axis angular acceleration is provided directly by two accelerometers.

  44. Accelerometer noise requirement • Requirement on linear axes is < 10-8 m/s^2/Hz. Considering preliminarily the control engineer rule of thumb, the measurement noise spectral density shall be < 10-9 m/s^2/Hz. • From the following figures, the angular acceleration noise shall be <10-10 rad/s^2/Hz on Y and Z axes, and 10-9 rad/s^2/Hzon X axis.

  45. Accelerometer noise requirement • Taking into account the following composition rules: • the following constraints result (angular acceleration requirements drive the results): • The same consideration for accelerometers on S2 (indices 3 and 4).

  46. Accelerometer noise requirement • DFAC channel – Requirement on the accelerometer noise. • The noise spectral density for frequency lower than 1 mHz is very important for accelerometer/star-tracker data fusion (real-time and post-facto).

  47. Requirements on harmonics errors Definition of requirements for harmonic errors • As addressed in previous slides, there are errors that may not be considered as the results of a stochastic random process. Those errors are better described as harmonic components. • It seems that requirements shall be provided for two classes of errors. • What do the scientists think about such topic?

  48. Reply to ESA comments on TN2 • Reply to ESA comments on TAS-I TN2 “SYSTEM DRIVERS” , concerning FF and AOCS. • ESA Q2) Pag. 25 – Section 3.3.7.3 “If the beam pointing device is removed…” & Pag. 42ESA comment/question: Let’s not forget that we have introduced and tested the Beam Steering Mechanism in order to eventually compensate the HF noise of the RWs. This brings an additional argument in removing it now, but of course not in the relaxations of the requirements (to be verified along the NGGM study). • TAS-I reply: at the distance of 10km, the requirement remains the same.

  49. Reply to ESA comments on TN2 • ESA Q3) Table 3.3.4 “Telemetry generation rate” – “3 linear + 3 angular accelerations per accelerometer” • ESA comment/question: From each accelerometer is intended to have 6 measurements? • TAS-I reply: It should be better to provide different requirement for accelerometers’ scientific channel, DFAC channel and housekeeping. • Scientific channel: 8 electrodes voltages + validity data @1Hz • DFAC channel: 3 linear + 1 angular accelerations + validity data @10Hz. • Housekeeping: TBD. The estimate will be updated taking into account the GOCE EGG data.

  50. Reply to ESA comments on TN2 • ESA Q5) Pag 43 “..the formation control must operate below the science BW of 1 mHz”ESA comment/question: This is very interesting, as proposed at the beginning of the last TAS study. In principle It can lead to a better performance and convergence of the formation controller w.r.t. the steps chosen in the last versions (from every 0.1s to 40s roughly) • TAS-I reply: the above sentence shall be considered as for the control bandwidth, not for the sampling frequency. It is necessary to take into account that the lines due to sampling frequency are present at frequency 2fs, 3fs, 4fs, etc.

More Related