The e.Deorbit CDF Study - Space...

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The e.Deorbit CDF Study

A design study for the safe removal of a large space debris

Robin Biesbroek and the e.Deorbit CDF [email protected]

6th IAASS ConferenceMontréal, Canada, 22/05/2013

mailto:[email protected]

e.Deorbit CDF Study | Robin Biesbroek | 6th IAASS Conference | Montreal, Canada, 22/05/2013 | Slide 2


e.Deorbit Mission Objectives

1. The mission is to perform active space debris removal

a. uncooperative target (large satellite or upper

stage) with heavy mass

b. in the 800-1000 km SSO/polar region

2. A study to design this mission was launched in the frame

of ESA Cleanspace initiative

3. The e.Deorbit study was performed in ESA’s Concurrent

Design Facility (CDF) between June and September

2012.

ESTEC CDF is a state-of-the-art facility equipped with a

network of computers, multimedia devices and software

tools, allowing a team of experts from several disciplines

to apply the concurrent engineering method to the design

of future space missions (www.esa.int/cdf)



e.Deorbit Mission Objectives

The e.Deorbit CDF study is intended to produce a preliminary system design for

the most promising option(s), identify the required technology roadmap(s), and

investigate its (their) applicability to relevant ESA missions. In details:

1. Assess the feasibility of a mission for the controlled de-orbiting and re-entry

of a large target in SSO, using technologies analyzed in previous CDF

studies (e.g. tentacles, robotic arm, harpoon, net, …)

2. Carry out a system level conceptual design of the spacecraft with the

participation of all discipline specialists

3. Trade-off different mission scenarios

4. Assess programmatics, risk and cost aspects of the various alternatives

5. Consolidate the technology road maps in line with the programmatic aspects

of the mission

6. Evaluate the applicability of the technologies to different categories of

satellites and debris remediation mission



Target assumptions

1. The target is in the ±8t mass range

2. The target is inactive, not cooperative, and not passivated

3. The orbit is SSO, ±800 km, near dawn-dusk

4. It is assumed that the target LV-PL adapter zone is not accessible (deployed solar

panel may obstruct it)

5. Long term attitude assumption: gravity-gradient stabilized (various attitudes possible)

6. Possibility of oscillation around equilibrium with angular rates comparable with ω0



Will the target break up?

1. Analysis of Fragmentation Events in

DISCOS (257 total, 58 relevant)

2. Analysis concentrated on SSO missions

only and excluded (Russian)

reconnaissance: this leaves 15 events

3. Statistics with low confidence (only 15

events for EO satellites) however results

allow to assume that the target will likely

be still intact after 10 years in orbit



Mission timeline



Main system trade-off options

1. Disposal:

a. Re-orbit to >2000 km

b. De-orbit to <600 km

c. Controlled re-entry

2. Propulsion:

a. Chemical (CP)

b. Electrical (EP)

3. Capture technique:

a. Robotic arm

b. Tentacles

c. Net

d. Ion-beam shepherd



System trade-off (re-orbit & de-orbit to 600 km)

Re-o

rbit

De-o

rbit t

o 6

00 k

m



System trade-off (controlled re-entry)Contr

olled r

e-e

ntr

y



System trade-off results

Two options received the best scores in terms of cost, programmatics,

safety and risk evaluation criteria:

Tentacles option

• Controlled re-entry

• Chemical propulsion

• Capture with tentacles

Net option

• Controlled re-entry

• Chemical propulsion

• Capture with net

For these options a conceptual design was performed by the CDF team,

which is shown in the following slides



Tentacles Option



Tentacles: launch & transfer

1. S/C main dimension of 1.2 x 1.2m x 2.92m

2. Fits in Vega LV fairing

3. 4x 425N bi-prop thrusters mounted on the side panel (w.r.t.

launch configuration)

4. Injection orbit:

• The lower the altitude, the higher the S/C dry mass

• Injection altitude: 300 km

• circular orbit

• Inclination: 98.2 deg

• Daily launch opportunity

• Launcher performance: 1660 kg

5. Launcher dispersions correction: 7m/s (also covers the

phasing worst case)

6. Hohmann transfer from injection orbit to target orbit (260

m/s)

7. RDV with target followed by attitude manoeuvre to match

rotation



Tentacles: capture

1. Strategy :

a. Capture a holding point using a robotic arm

b. Embrace the target with clamping mechanism so that it

cannot escape

c. Positively lock the clamping mechanism with a force higher

than the de-orbiting force (using pushing rods based on

IBDM actuator)

d. Composite (chaser + target) is stiff during de-orbit operation

e. The method is generic: it can be applied to the LV I/F

adapter

2. Is the robotic arm necessary?

a. If no robotic arm is used, bouncing may occur

b. Current simulations make the assumption that there is no relative

motion between chaser & target during the clamping mechanism closing

time (after AOCS cut-off and before achieving capture)

c. Detailed analysis should conclude if this assumption is valid or not

d. If not, a robotic arm is used to achieve the first capture; allows to guide

the clamping mechanism to a position suitable for capture

e. The robotic arm also helps for contingency reasons

3. For this reason a robotic arm is baseline, at the price of an extra cost



Tentacles: de-orbit strategy

1. Impulsive burn: 202 m/s

2. Initial mass: 9350 kg

3. Engine Isp: 321 s

4. Different set of engines: 1x425 N, 2x425 N = 850 N, 4x425 N = 1700 N, 8x425 N = 3400 N, 16x425 N = 6800 N

5. Different scenarios:a. Immediate de-orbit

b. Perigee reduction to 200 km, then re-entry

c. One intermediate perigee altitude to 200 km, then re-entry

d. Two intermediate perigee altitude to 200 km, then re-entry

e. Three intermediate perigee altitude to 200 km, then re-entry

6. In case of intermediate perigee altitudes, they are optimised to minimise the gravity losses

7. Final burn should lower the perigee to 40 km in order to obtain <3000 km footprint

If one burn, 4x425 N is

recommended to get

reasonable gravity losses

If 3+1 = 4 burns, 2x425 N or even

1x425 N is enough

Baseline for tentacles option: 2x425N



Tentacles: Baseline design overview (1)

Spacecraft Characteristics

Mass

Dry mass 784.2 kg

Propellant

mass810 kg

Total mass 1648 kg incl. adapter

AOCS/GNC

3-axis stabilized

Sensors

3 x Star Tracker

2 x Sun Sensor

2 x IMU

2 x GPS receiver

2 x LiQuaRD LiDAR

2 x Far Field Camera

2 x Near Field Camera

Actuators 4 x Reaction Wheels

RCS 24 x 22 N Thrusters





Propulsion

Bi-propellant system

Main Engine: 2 x 425N LAE (+ 2 redundant)

Tanks: 4 polar mounted based on Eurostar 2000, 2

pressurant tanks

Power

SA

1 wing

Solar cells: 30% 3J GaAs

Area: 2.8 m2

785 W (EOL)

Battery1 x Lithium Ion ( -HC cells)

BoL energy: 562 Wh

Bus 28V MPPT dual reg/unregulated bus





Communications

TDRS – S-

band

2 x TRSP

3 x Antenna

2 x HPA

DTE – X-

band

2 x TRSP

3 x Omniantenna

Thermal MLI, OSR, black paint, heaters

DHSdynamically reconfigurable payload processor

(DRPM)

Mechanism

Clamping mechanisms & drive mechanism

Pushing rods

SADM & SADE

Robotics Robotic arm incl. gripper

StructureBox-shaped structure with central shear panel and

stiffeners for the tank panels



Net Option



Net: launch & transfer

1. Use Ø 937mm VEGA I/F

2. Ø 937mm central cylinder (2.4m high)

3. 2x Ø 430mm pressurant tanks (inside the central cylinder)

4. Use 4x425N biprop thrusters + 4x ATV 220N thrusters

5. Solar Panel of 2m x 1m x 1.5m

6. Injection orbit same as Tentacles option:• The lower the altitude, the higher the S/C dry mass

• Injection altitude: 300 km

• circular orbit

• Inclination: 98.2 deg

• Daily launch opportunity

• Launcher performance: 1660 kg

7. Launcher dispersions correction: 7m/s (also covers the phasing worst case)

8. Hohmann transfer from injection orbit to target orbit

(260 m/s)

9. RDV with target (up to 50 m)



Net: capture

1. Approximately 250 (~6000 DOF per simulation) multi-body simulations performed:

a. Included capture of the target, tensioning, and de-orbitation burn.b. For a variety of relative rotation rates, velocity, tether parameters and

controllers2. Mesh: 2m (to limit computation time), 0.25m shall be analysed

The following parameters were varied in the 250

simulations:

a. Relative position (3 axis)

b. Relative rotation (3 angles)

c. Relative rotation rates (3 axis)

d. Relative motion (3 axis)

e. Tether length

f. Tether stiffness

g. Tether material damping



Net: de-orbit strategy

1. Impulsive burn: 202 m/s

2. Initial mass: 9350 kg

3. Engine Isp: 321 s

4. Different set of engines: 1x425 N, 2x425 N = 850 N, 4x425 N = 1700 N, 8x425 N = 3400 N, 16x425 N = 6800 N

5. Different scenarios:a. Immediate de-orbit

b. Perigee reduction to 200 km, then re-entry

c. One intermediate perigee altitude to 200 km, then re-entry

d. Two intermediate perigee altitude to 200 km, then re-entry

e. Three intermediate perigee altitude to 200 km, then re-entry

6. In case of intermediate perigee altitudes, they are optimised to minimise the gravity losses

7. Multiple burns lead to complicate AOCS control of the stack

If one burn, 4x425 N is

recommended to get

reasonable gravity losses

If 3+1 = 4 burns, 2x425 N or even

1x425 N is enough

Baseline for net: 2x425N+ 2x220N (pulse mode / off modulation)

option for net



Net: Baseline design overview (1)


Mass

Dry mass 708.9 kg

Propellant

mass878.2 kg

Total mass 1622 kg incl. adapter

AOCS/GNC

3-axis stabilized

Sensors

3 x Star Tracker

2 x Sun Sensor

2 x Gyrometer

2 x GPS receiver

2 x LiQuaRD LiDAR

2 x Far Field Camera

RCS 24 x 22 N Thrusters





Propulsion

Bi-propellant system

Main Engine: 4 x 425N LAE (+ 2 redundant)

4 x 220 N thruster

Tanks: 4 polar mounted based on Eurostar 2000,

3 pressurant tanks

Power

SA

2 wings

Solar cells: 30% 3J GaAs

Area: 2.8 m2

785 W (EOL)

Battery1 x Lithium Ion (- HC cells)

BoL energy: 562 Wh

Bus 28V MPPT unregulated bus





Communications

TDRS – S-

band

2 x TRSP

2 x Antenna

2 x HPA

DTE – X-band2 x TRSP

2 x Omniantenna

Thermal MLI, OSR, black paint, heaters

DHSdynamically reconfigurable payload processor

(DRPM)

Mechanism

2 x Net & ejector

2 x Tether & Cable Cutter

2 x SADM & SADE

StructureCentral cylinder with box-shaped structure around,

stabilized with shear panels



e.Deorbit CDF study: conclusions

• Two options were designed showing feasible first assessments

a. Both are similar in mass & dimensions and appear suitable for a VEGA launch within one decade

b. Net option shows fewer severe risk items, is less sensitive to target shape, 6-7% cheaper and 6 months lower development time

• Open points tentacles:

a. Further analysis of the rendezvous and mating operations for the option using only the tentacles (need for extra LIDAR, tentacles closing parameters)

b. Good models of the target and sensor models are required, namely for the points being used for relative navigation in the close range (SAR antenna)

c. Need for ground in the loop for the capture operations has to be re-assessed

d. Structural integrity of the target holding points should be further analyzed

• Open points net:

a. Plume impingement and chemical reactions with the tether material.

b. ATD analysis for optimization of the thrusters configuration of both the multi and single burn designs (number of thrusters, tilting angle, TPS required)

c. Tether cutting scenario, behavior of the tether after cutting, interaction with following capture

d. Optimization of the tether control and tensioning ΔV can have an important impact on the propellant mass

e. Detailed analysis and definition of the operation concept for the de-orbit phase and tether dynamics damping

f. For single-burn option, out-gassing of the TPS after orbit raise burn should be assessed (even though it is a shorter burn)



Life after the e.Deobit CDF study

1. Three system activities were initiated:

a. Service oriented approach towards the procurement/development of an

ADR mission (Astrium SAS, Kayser Threde, SSTL)

2. Activities in support of ADR capturing (ITT already issued):

a. Advanced GNC for ADR (net/tether control during burns)

b. Assessment and simulation of a tentacles based capture mechanism for ADR

3. GSP proposals (detumbling, etc.)

4. GSTP compendium published on EMITS

a. GNC activities (nav-cam, LIDAR, image recognition)

b. Net activities (characterisation and 0g testing, throwing mechanism,

locking)

c. Clamping mechanism activities (breadboard)

d. Other sub-systems (harpoon, ATD, reconfigurable payload processor)

e. Debris attitude modelling

5. E.Deorbit Phase A ITT release planned for Summer 2013

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