© gmv, 2015 property of gmv all rights reserved formation flying guidance for space debris...

© GMV, 2015 Property of GMV

All rights reserved

FORMATION FLYING GUIDANCE FOR SPACE DEBRIS OBSERVATION, MANIPULATION AND CAPTURE

ASTRONET II INTERNATIONAL FINAL CONFERENCE

T. V. Peters (GMV)

© GMV, 2015FORMATION FLYING GUIDANCE FOR DEBRIS OBSERVATION, MANIPULATION AND CAPTURE

OVERVIEW

Presentation loosely organized around mission phases, taking examples from different projects– Engineering issues– GNC aspects, with emphasis on guidance

Material from following projects:– Detumbling: detumbling space debris after capture– Patender: net capture tests– COBRa: influencing debris (orbit and) attitude by plume impingement– Android: demonstrate robotic and net capture of space debris– eDeorbit: de-orbit Envisat

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INTRODUCTION

Introduction– Space debris distribution– Space debris dynamics– Debris capture options

Mission phases for robotic capture– Mid-range rendezvous– Inspection from spiral orbit– Attitude synchronization– Capture and detumbling

Conclusion

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All rights reserved

INTRODUCTION



DEBRIS CLASSIFICATION

Removal options– Removal of tiny & small debris not practical– Removal of large objects removes potential sources of fragments in case of

collision• Target selection is based on debris generating potential

Conclusion: remove large objects

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Type Characteristics HazardTiny Not tracked, <1 cm Shielding exists,

damage to satellites may occur

Small Not tracked, diameter 1 – 10 cm, 98% of lethal objects, ~400.000 objects in LEO

Too small to track and avoid, too heavy to shield against

Medium Tracked, diameter >10 cm, <2 kg, 2% of lethal objects, ~24.000 objects in LEO, > 99% of mass (incl. large objects)

Avoidance manoeuvres performed most often for this category

Large Tracked, >2 kg, <1% of lethal objects, > 99% of mass (incl. medium objects)

Primary source of new small debris, 99% of collision area and mass


DEBRIS DISTRIBUTION

Debris population– Total mass estimated at 6300

tons– High concentration at 82-83°

inclination• COSMOS 3M

SSO particularly important for and Earth observation and science– SSO inclination-paired with 82-

83° inclination orbit– Heightens collision probability

• Orbit planes may align, leading to head-on collisions during entire orbit instead of only at nodes

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DEBRIS ATTITUDE DYNAMICS

No systematic survey for attitude Several sources of data are available

– The spin rate of upper stages tends to slow down (1 & 2)• Spin-ups have been observed, likely due to outgassing events (2)

– Envisat (3 & 4)• Spin-up event occurred some time between april 2012 and november 2013• Rotation rate has been slowing down since then

– Rocket upper stages have generally been observed in a flat spin (i.e., non-axial) (5)• Initial spin state of rocket bodies tends to be axial• Therefore it is expected that a transition to a major axis spin occurs at some

point due to energy damping

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1. Boehnhardt, H., Koehnhke, H. and Seidel, A. 1989, The acceleration and the deceleration of the tumbling period of Rocket Intercosmos 11 during the first two years after launch, Astrophysics and Space Science, vol. 162, no. 2, p. 297-313.

2. Williams, V., Meadows, A.J., 1978, “Eddy current torques, air torques and the spin decay of cylindrical rocket bodies in orbit”, Planetary and Space Science, vol. 26, 1978, p.721-726

3. Bastida Virgili, B., Lemmens, S., Krag, H., 2014, Investigation on Envisat attitude motion, e.Deorbit Workshop4. Kucharski, D., Kirchner, G., Koidl, F., Fan, C., Carman, R., Moore, C., Feng, Q., 2014, “Attitude and Spin Period of Space Debris

Envisat Measured by Satellite Laser Ranging”, IEEE Transactions on Geoscience and Remote Sensing, Vol. 52 , Issue 12, pp. 7651 – 7657, DOI 10.1109/TGRS.2014.2316138

5. Santoni, F., Cordelli, E., Piergentili, F., 2013, "Determination of Disposed-Upper-Stage Attitude Motion by Ground-Based Optical Observations", Journal of Spacecraft and Rockets, Vol. 50, No. 3, pp. 701-708, doi: 10.2514/1.A32372


DEBRIS ATTITUDE DYNAMICS

Envisat Rotation axis known Characteristic decay time ~4.5 years

COSMOS Rotation around major axis Characteristic decay time between

100 and 470 days, with a mean of 161 days and median of 129 days

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CAPTURE METHODSCapture method Sensitivity to rotation rate Rotation related issues Structural issues

Net low fast de-spin required to avoid tether wind-up around target

may require measures to avoid breaking off pieces of target

Grappling high synchronization required requires structural hard point

Docking with nozzle high synchronization required requires non-steerable nozzle

Tentacles high synchronization requiredmay require structure not covered by MLI for firm grip

Harpoon (Rigid) high synchronization required

requires strong structure for contact (e.g., honeycomb panels) and avoidance of propellant tanks

Harpoon (Non-rigid) lowfast de-spin required to avoid tether wind-up around target

requires strong structure for contact (e.g., honeycomb panels) and avoidance of propellant tanks

Pushing sock air-bag high requires pre-capture de-spinmay require measures to avoid breaking off pieces of target

Foam projection highcentrifugal forces may disrupt foam; requires pre-capture de-spin

may require structure not covered by MLI for firm grip (i.e., MLI may tear off)

Ion-beam Shepherd lowlow sensitivity to spin rate; method may be used to control rotation

none

Electrostatic tractor (only for GEOs)

lowlow sensitivity to spin rate; method may be used to control rotation

none

Magnetic tractor lowlow sensitivity to spin rate; method may be used to control rotation

none

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ROBOTIC ARM CAPTURE

Precursor activities dealt with cooperative targets (attitude controlled, visual markers, grappling interfaces) – ETVS-VII– Orbital Express (DARPA program)

FREND (DARPA) performed on-ground demonstration of capture of uncooperative target debris

Other missions/concepts being investigated:– DEOS (passive v.s. active chaser AOCS

investigated)– eDeorbit (several robotic arm and

tentacles configurations proposed, as well as net-based capture)

– ANDROID (double demonstration of robotic arm and net)

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ETS-VI I (NASDA/JAXA) Orbital Express (DARPA)

DEOS eDeorbit concept from ESA CDF study


NET CAPTURE

Net Several studies to mature net

capture technology (net design, net deployment strategy and mechanisms)– Patender

• Scalable to debris mass and size • Composed of a pyramidal, conical or

plane net stowed in a canister with four masses (bullets) attached to net vertices

• Pneumatic or spring-driven ejection of bullets

• Tether connection after capture Studies to investigate

controllability– AGADiR

• Controllability remains a difficult problem

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Rbar


All rights reserved

MISSION PHASES



ACTUATORS AND SENSORS

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Actuators Depends on size of object Separate thrusters for orbit raising / lowering & rendezvous

– Orbit raising / lowering: • (2 x) 2 x 22 N (Android; stack mass 425 kg)• 2 x 500 N (eDeorbit; stack mass 9500 kg)• Acceleration per thruster 0.05 m/s2

– Rendezvous• (2 x) 8 x 1 N (Android; mass during rendezvous 298 kg)• 28 x 22 N (eDeorbit; mass during rendezvous 1700 kg)• Acceleration per thruster 0.01 m/s2 (eDeorbit) / 0.003 m/s2 (Android)

Sensors

Sensor ModelMass [kg]

Power [W]

Range [km]

Performance Source Comments

GPS Phoenix 0.02 0.85- up to 2 m DLRWAC DVS 2.4 130.02-150 1" TSD

LIDAR RVS 13.8 350.001-2

0.01 m short range; min range < 1m0.5 long range

JENA OPTRONIK scanning


MID-RANGE RENDEZVOUS

Far-range rendezvous performed using TLE and GPS

TLE accuracy after 1 week of propagation:– Radial: maximum error < 1.5

km => drift of 14 km per orbit• Bias .25 km + 1σ of .1 km =>

drift of 3.3 km per orbit– Cross-track: maximum error <

1.5 km– Along-track: maximum error <

30 km Is a detection & handover to

WAC possible?

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Debris TLE

Abs NGPS based

-

Trans G

Relativestate

NORAD TLE

GPS DATA

ΔV

Att GTGT TRACK

Att. Cmd.

Rel NWAC based

WAC DATA

Rough estimateof relative state

Translation

Attitude

1. Legendre P., Deguine B, Garmier R., Revelin B., 2006, Two Line Element Accuracy Assessment Based On A Mixture of Gaussian Laws, AIAA 2006-6518, AIAA/AAS Astrodynamics Specialist Conference and Exhibit, 21 – 24 August 2006, Keystone, Colorado



Search phase Detection limit 0.25

pixels– Camera FOV 28°

Detection can occur between 33 - 151 km distance– Depending on size of target– Uncertainty cone of .5° –

2.5°• Well within WAC FoV

– Handover to WAC is possible

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range [m]number of pixels covered

/ object sizecomments

2 4 9 151000 0.06 0.11 0.25 max range WAC

67000 0.13 0.25 0.56 max range WAC33500 0.25 0.50 1.13 max range WAC

2000 4 8 19 handover distance

20 419 835 185590% of WAC FOV filled

9 927 1833 388690% of WAC FOV filled

4.5 1833 3505 658390% of WAC FOV filled

2 3886 6583 9660 working distance

target

chaser

Detection range 50 - 150 km

1.5 km

2ϕ

z

x

±5 km



Phase 1– Difference in SMA larger than

radial uncertainty in TLE• ±2 km difference in SMA

– Vision-based navigation requires some radial motion for faster convergence

– Small target may lead to late detection

Phase 2– Terminal orbit may require

specific relative geometry • Lighting conditions• Earth in background• Ground contact

– Modulate drift to accommodate terminal conditions

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100 m

50 m500 m

z

x

S1

S3aS3b

S4S5

~ 4000 m

S2a

S2b

4 km

.5 km2 km

z

x

S0S1

33 – 150 km

# of orbits

<1 km

# of orbits available for detectiondetection distance [km]

33 67 150δS.M.A. [km]

0.5 5.94 13.16 30.772 1.49 3.29 7.69

3.5 0.85 1.88 4.40



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Plan generation

Mean orbit Kepler orbittarget orbit

chaser LVLH state

Table of times, referencestates & ΔV’s

reference trajectorycorrectiontime check

time

ΔV computation

Reference trajectory

ΔV

Guidance function

Plan database

Situationassessment

mode manager commands

Guidance expert function

Plan generation

Mean orbit Kepler orbittarget orbit

chaser LVLH state

Table of times, referencestates & ΔV’s

reference trajectorycorrectiontime check

time

ΔV computation

Reference trajectory

ΔV

Guidance function

Full guidance Guidance as implemented


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Initial errors are quite large After first 5 hours (3 orbits) errors decrease

– Decrease occurs when chaser enters drift orbit – Large errors in position and velocity occur at

large distances Errors in position and velocity show a slight

increase over time– due to the fact that a unperturbed Keplerian

propagator is used to propagate relative trajectories• Keplerian orbit is initialized at start of simulation • starts diverging from true orbit over time

Causes of errors are known, and could be improved. – J2-based relative propagator could be used to

improve the reference trajectory– Guidance could be made to operate on

linearized differential orbital elements• Suffer less from linearization errors

– Guidance could periodically update its plan• Re-initializing Keplerian orbit used to generate

reference trajectory• Reference orbit will be closer to true orbit and

reference trajectory will be closer to truth

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time [h]

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INSPECTION FROM SPIRAL ORBIT

Camera in target pointing when constraints are met

Several constraints shall be taken into account:– Eclipse times (no operation)– Sun exclusion angle (50deg)– Earth in the field of view (IP

problems)– Illumination conditions, angle

Sun Picard Mango below 90 deg (IP problems)

When all constraints taken into account only about 25% of the orbit is useful, +ZY in LVLH

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INSPECTION FROM SPIRAL ORBIT

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Effect of perturbations (SRP and Drag) lead to non-constant drift rate– Needs to be taken into account in manoeuvre definition for

spiral orbit insertion if in drift-free orbit– Leads to more correction manoeuvres– Note: size of spiral orbit fairly small; 10 m vs. 20 – 50 m for

COSMOS - Envisat Possible input to spacecraft design to ensure small

differences in ballistic coefficient


SYNCHRONIZATION PHASE

Elements Target attitude propagation Reference frame transformations Quaternion spline curve for fly-around

to limit accelerations Straight-line approach using ramp-

constant-ramp velocity profile

Guidance plan consists of:1. Perform station keeping on Vbar2. Transfer to target ω-vector (h-vector is an

alternative)3. Station-keeping at target ω-vector4. Rotate to target co-rotating5. Perform transfer closer to target6. Station-keeping at w-vector7. Transfer to target body fixed frame position8. Station-keeping in body fixed frame position9. Transfer closer to target

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XLVLH

ZLVLH

S1

S2

S3

S4

S5

A

1. Myoung-Jun Kim, Myung-Soo Kim, and Sung Yong Shin. 1995. A general construction scheme for unit quaternion curves with simple high order derivatives. In Proceedings of the 22nd annual conference on Computer graphics and interactive techniques (SIGGRAPH '95), Susan G. Mair and Robert Cook (Eds.). ACM, New York, NY, USA, 369-376. DOI=10.1145/218380.218486 http://doi.acm.org/10.1145/218380.218486



Synchronization simulated in simplified simulator– Propagation models

• Trajectory propagator contains J2 perturbation

• Attitude propagation propagates torque-free tumbling motion (no perturbations)

– Sensors and actuators• 1 N thrusters with thruster

management function• “Perfect” sensors

– GNC• Synchronization guidance• LQR controller• “Perfect” navigation

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Camera view LVLH view

ω = 0.5 °/s


SYNCHRONIZATION

Guidance trajectory is precisely followed– Centimetre level accuracy in

position– Millimetre per second level

accuracy in position– Pointing error smaller than

0.1°– Better accuracy possible

with more aggressive controller

But, no navigation is included

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SYNCHRONIZATION

ΔV required for synchronization– guidance ΔV is red– true ΔV is black

Four phases can be distinguished– transfer to the angular velocity vector

• fairly expensive; sharp increase in ΔV right at start

– station-keeping at the angular velocity vector• comparatively cheap; gradual increase in ΔV

– transfer to the body fixed frame – station-keeping in body fixed frame

• Station-keeping in target reference frame is cheaper than transfer

• But considerably more expensive than station-keeping at angular velocity vector

• Especially considering that station-keeping at angular velocity vector is performed at 10 m, while station-keeping in target body frame is performed at between 2 and 5 m

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CAPTURE AND DETUMBLING

Capture and detumbling currently under investigation– Inverse kinematics dependent on

design of arm; full implementation provides low added value compared to cost • Some simplification (e.g. convenient

arm/joints configuration and allow small joint angles w.r.t. rigid)

– Contact model between robot hand and target approximated by translational and rotational spring damper Kc Dc system • no gripper is attached to end point of

manipulator, contact occurs just between two points

– state of spring damper could supply a metric on what is happening at contact

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1. Dimitrov, D. N., Kazuya Y., 2004, "Momentum distribution in a space manipulator for facilitating the post-impact control," Intelligent Robots and Systems, 2004.(IROS 2004). Proceedings. 2004 IEEE/RSJ International Conference on, vol. 4, pp. 3345-3350. IEEE, 2004.


CONCLUSION

Brief outline of mission phases for a debris capture mission has been presented– Overview of results of several related projects

Envisat most likely candidate for a debris removal mission– Exceptional for high rotation rate– Special measures to be taken for attitude synchronization

Smaller ADR demonstration mission with a smaller target should be implemented

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© gmv, 2015 property of gmv all rights reserved formation flying guidance for space debris...

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