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The L1 Orbit Used for Servicing (LOTUS):
Enabling Human/Robotic Servicing Missions
in the Earth-Moon System
Brent Wm. Barbee
Future In-Space Operations (FISO) Telecon Colloquium
June 16th, 2010
Background
 NASA-GSFC is currently studying a suite of
notional missions to inform a forthcoming
congressional report on spacecraft servicing
capabilities and concepts
 The 5th Notional Mission involves
human/robotic servicing of a large Sun-Earth
L2 (SEL2) telescope in the Earth-Moon
system
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Mission Profile
 A large telescope stationed at SEL2 returns to the
Earth-Moon system and rendezvouses with a robotic
servicing vehicle in a Lyapunov orbit about EarthMoon L1 (EML1)
 A crew vehicle carrying astronauts launches to
rendezvous with the servicer/telescope stack
 After servicing is complete, the crew vehicle returns
to Earth and the telescope returns to SEL2
 The robotic servicer spacecraft remains in orbit for
25 years
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Telescope Considerations
 Minimize telescope maneuver magnitudes

Conserve telescope propellant

Avoid large thruster-induced accelerations
 Minimize telescope down-time

Avoid excessive travel time to/from SEL2
 Telescope is assumed to be a cooperative
rendezvous target for the robotic servicer
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Robotic Servicer Orbit
 Robotic servicing vehicle has a 25 year
lifetime
 The orbit it inhabits in the Earth-Moon
system must:



Be easily accessed by both the crew vehicle and
the telescope
Require minimal station-keeping ΔV
Remain well clear of the Van Allen Belts and
GEO
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Crew Vehicle Objectives
 Crew vehicle trajectory should:


Maximize available time for servicing
Provide a total round-trip flight time (launch to
landing) of at most 21 days

Offer a free return from launch if possible

Stay clear of the Van Allen Belts and GEO

Provide safe atmospheric re-entry

Maximum atmospheric re-entry velocity of 11 km/s, as per
Apollo 10
 Notional Orion was assumed for crew vehicle
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Robotic Servicer Trajectories
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Telescope Trajectories
 Telescope can travel relatively easily between
its SEL2 halo orbit and the EML1 Lyapunov
orbit via low-energy transfers

ΔV from SEL2 to EML1 = 45 – 50 m/s

ΔV from EML1 to SEL2 < 1 m/s

Flight time between EML1/SEL2 = 50 – 130 days


Faster transfers are possible but require
considerable ΔV
Some telescope downtime will have to be tolerated
in exchange increased lifetime from servicing
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Rendezvous at EML1
 The robotic servicer can rendezvous with and
capture the telescope on the EML1 Lyapunov
orbit relatively quickly for modest ΔV costs
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Example EML1 Rendezvous

The relative motion
dynamics between
spacecraft on a
libration point orbit
are completely
different from the
familiar relative
motion dynamics
between
spacecraft in Earth
orbit (LEO, GEO,
etc.)
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Crew Trajectory Alternatives

The first option considered was to
send the crew directly to the EML1
orbit and perform servicing there


However, this only offered ~ 11 days
for servicing, which was insufficient
for the planned activities


Crew has a free return from launch if
necessary
Outbound and inbound times are not
selectable
Additionally, the EML1 orbit
experiences eclipses that can be 9 to
12 hours in length
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Crew Trajectory Alternatives
 The second option considered was to place
the crew onto a large 21 day long Highly
Elliptical Orbit (HEO) about Earth

Completely free return for the crew
 However, bringing the robotic servicer and
telescope to this orbit within 1 – 3 days of
launch and having them depart within 1 – 3
days of re-entry required > 2,000 m/s of ΔV
from the robotic servicer and telescope,
which is not permissible
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Zero-Velocity Curve Analysis

The next approach
was to study the
restricted three-body
dynamics

I noticed that there
was a large volume
of space around
Earth that should be
accessible from the
EML1 orbit for very
little ΔV …
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The LOTUS

Ultra-low departure ΔV
from EML1 orbit is
easily achieved by the
servicer/telescope
stack
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The LOTUS in the Inertial Frame

The LOTUS is a
“HEO” with a high
perigee

Eccentricity of 0.54

Period of ~ 10 days

The LOTUS perigee
is 83,777 km, well
above the Van Allen
Belts and GEO
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Crew Launch to a LOTUS Apogee
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Crew Free Return From Launch
 The crew always
has a free return
from launch
available from in
case LOTUS
insertion must
be aborted
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Servicing on the LOTUS

The crew spends 16
days on the LOTUS

1 day is for AR&B with
the servicer/telescope
stack

15 days for servicing

Meets requirements
for the notional
mission under
consideration
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Crew Return to Earth
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Return to EML1

The LOTUS naturally
returns to EML1 ~ 98
days after initial
departure

The servicer can easily
reinsert into the EML1
Lyapunov orbit

The telescope can
easily continue past
EML1 and transfer
back to SEL2
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Mission Summary

Total crew vehicle ΔV
(including 100 m/s for
AR&B) is 2120 m/s


Quite reasonable considering
crew vehicles for lunar
missions historically had a
2800 m/s capability
Well within the notional Orion,
Ares I, Ares V capability

Total servicing time of 15
days

Total round-trip time of
20.54 days
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Trajectories in the Inertial Frame
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LOTUS Eclipse Analysis

Eclipses on the LOTUS are much reduced compared to the
EML1 Lyapunov orbit

Judicious choice of start date completely avoids eclipses during the LOTUS

Worst case eclipse duration on the LOTUS is about 4 hours
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Servicing Time Flexibility

The 15 day servicing case shown here is only one possibility of
many

The advantage of the LOTUS is that the crew can arrive at / depart
from any points on the LOTUS, making the servicing time
selectable

With a 21 day round-trip flight time limit, the maximum servicing
time available is 19 days

Launch into /depart from a LOTUS perigee

0.6 day flight time from launch to LOTUS insertion, same for de-orbit

Launch C3 is -7.98 km2/s2

Insertion / De-Orbit ΔV is 1800 m/s

Total crew vehicle ΔV of 3700 m/s, not unreasonable for such a low launch
C3 and in light of historical and notional future mission capabilities
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Summary and Conclusions

The LOTUS offers key advantages for human/robotic
servicing missions in the Earth-Moon system

Selectable time on-orbit for servicing

Low ΔV access to/from EML1 and therefore to/from SEL2

Avoidance of eclipses reduces battery size requirements, saving
considerable spacecraft mass

No orbit maintenance/station-keeping maneuvers required on LOTUS

Launch C3 and ΔV requirements consistent with anticipated capabilities

Crew always has a free return from launch if necessary

LOTUS perigee is well above the Van Allen Belts and GEO

Atmospheric re-entry velocity when de-orbiting from any point on the
LOTUS is always ≤ 11 km/s
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