staging from earth-moon l-2 orbits - gateway or...
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Staging from Earth-Moon L-2 Orbits -
Gateway or Tollbooth?
David W. Dunham, KinetX Aerospace, Inc.
Kjell Stakkestad & James McAdams, KinetX Aerospace
Jerry Horsewood, SpaceFlightSolutions, Inc.
Anthony Genova, NASA-Ames & Florida Institute of Technology
FISO Telecon, 2019 March 13 1
Robert Farquhar Envisioned an “International
Exploration Station” in a high-energy
libration-point orbit in 1969
• His early idea was to use a Sun-Earth L1 halo
• Later, Bob realized that an EM-L2 Halo was
a better staging location than SE-L1 and
realized that EM-L2 could support Lunar
exploration as well
Published in: Farquhar, R. W., “Future Missions
for Libration-Point Satellites,” Astronautics &
Aeronautics, Vol. 7, No. 5, pp. 52-56, May 1969.
Robert W.
Farquhar
1932 – 2015
Master Celestial
Mechanician,
Father of Halo Orbits
and Asteroid
Exploration (NEAR-
Shoemaker to Eros)
2
Human Exploration of the Moon, Near-Earth
Asteroids, and Mars using Staging from Earth-
Moon L-2 Orbits and Phasing Orbit Rendezvous
David W. Dunham, KinetX Aerospace, Inc.
Kjell Stakkestad, Peter Vedder, & James McAdams, KinetX Aerospace
Jerry Horsewood, SpaceFlightSolutions, Inc.
Anthony Genova, NASA-Ames & Florida Institute of Technology
Roberto Furfaro and John Kidd, Jr., Univ. of Arizona, Tucson
IAC-18-A5.2.2 (x45050)
69th IAC, Bremen, Germany, 2018 October 3
My presentation is
largely taken from
this one.
3
Introduction - 1• Creation of a Sustainable Reusable Infrastructure for Human
Missions to the Moon, NEOs, Mars, and beyond.
• Adoption of a “Pathways Approach”
to Human Space Exploration as
recommended by the NRC
Committee on Human Spaceflight.
• Our Pathway is from an Earth-Moon
L2 Halo Orbit to Earth Phasing Orbits,
and an Earth-Perigee Injection
Maneuver sending Humans to a variety
of Interplanetary Locations.
• R. Farquhar had basic ideas in 1968
• Earth-Moon L2 = EM-L2 4
Introduction - 2• Participation by International Partners is essential
• Our past work used impulsive burn trajectories
• Now, Xenon low-thrust solar electric propulsion (SEP) systems are
planned for key elements due to the higher Isp of SEP, so our newest
trajectories emphasize hybrid systems that would use SEP most of the
time, but chemical high-thrust would be used for some maneuvers to
avoid gravity losses
• Our work has used a 7000-km Z-amplitude EM-L2 halo, but now a
very large-amplitude halo, called a Nearly-Rectilinear Halo Orbit
(NRHO) is favored by many
• With a substantial lunar infrastructure, we favor 3 comm sats
spaced around a large-amplitude EM-L2 halo orbit, which with Earth,
would provide continuous coverage of all of the Moon & its
environment; then, the proposed Lunar Orbital Platform-Gateway
could optimize its lunar orbit for the current exploration goal
NRHO
5
Cargo Mission Possibility to EM-L2 Halo:Outward Lunar Swingby & SE-L1 WSB Transfer
• The vehicle only has to launch into a
trajectory just reaching the Moon’s orbit
rather than launching to Sun-Earth L1
distances to reach the WSB
• The vehicle could be launched into
phasing orbits before the lunar swingby,
allowing use of less V to correct launch
errors and time for spacecraft checkout
before the lunar swingby, as accomplished
for past missions such as Geotail, WIND,
WMAP, and STEREO
• Calculated with STK/Astrogator by Anthony
Genova, NASA Ames
• Robert Farquhar conceived many of the orbit
ideas shown here
• Farquhar’s Memoirs, “Fifty Years on the Space
Frontier: Halo Orbits, Comets, Asteroids, and
More” are available on amazon.com
Rotating ecliptic-plane view with
Horizontal Sun-Earth line
HOI
Earth
Lunar
orbit
• SE-L1
To Sun
swingby
TTI (from LEO) ∆V 3152 m/s
Post-TTI V: 21 m/s at apogee WSB and 5.4 m/s
halo orbit insertion (26 m/s total); Perigee Jan 13,
Halo insert July 18, 2018, TOF = 173 days
(apogee = March 26, 2018). Lunar swingby
altitude 9700 km.
6
LunaH-Map Transfer to Lunar OrbitA low-thrust cubesat (from EM-1) example of the
previously-shown trajectory by Anthony Genova
TRAJECTORY SEQUENCE
A) Launch on Earth-escape trajectory with
EM-1 on Oct. 7, 2018
B) Deploy from EM-1 {L+8.5 hrs}
C) Begin 2.5-day Thrust Arc (in velocity), 24 hours
after deployment {L+32.5 hrs}
D) Lunar Flyby (changes energy from escape to
weak capture around Earth); {L+80 hrs}
E) Begin 4-day Thrust Arc (anti-velocity) {L+156
hrs}
F) Apogee at 1 million km altitude (no maneuver);
{L+ 34 days}
G) Begin 5-day Thrust Arc (anti-velocity) to target
Moon and decrease approach speed {L+
64 days}
H) Weak Capture into Lunar Orbit {L+ 69 days}
ABC
D
E
F
G
H
Thrusting shown in
RED
Trajectory shown in Earth Inertial
Frame 7
Fast Transfers to the Earth-Moon L2 PointTrajectories shown in rotating system with fixed horizontal Earth-
Moon line, lunar orbit plane projection – Farquhar, 1971
A similar technique can
go to EM-L1 but is not
as efficient since the
powered lunar swingby
V is 800 m/sec
8
Mission Profile for a Lunar Shuttle System with
(Earth-Moon L2) Halo Orbit Staging Adding a mirror image of the bottom of the previous slide, Farquhar 1971
With certain geometries, very low V’s might be possible near L2 for a trajectory that
might be used for a quick mission that might spend about a week above the lunar far
side. A variant could rendezvous with a station in an EM-L2 orbit, which Farquhar
called a Halo Orbit Space Station, or HOSS. At the time, NASA proposed a Lunar
Orbit Space Station (LOSS) in a 60 n. mi. lunar polar orbit that would impact the
Moon in about 4 months unless it had considerable stationkeeping capability (about
400 m/sec per year). Bob Farquhar sarcastically stated that the LOSS would become
“a real LOSS”. This comment prompted NASA HQ to change the name of the lunar
station from LOSS to OLS. (p. 48 of Farquhar’s Memoires).9
Human Missions:LEO to EM-L2 Halo Orbit
• Mission to EM-L2 halo via powered lunar
swingby
– CTV post-injection V = 308 m/sec (& 295
m/sec to return)
10
••Earth L2
Return
2021 April 2 with
atmospheric
re-entry capsule
Launch
2021 March 3
V 3,129 m/s
from LEO
Moon
Rotating lunar orbit plane plot With return, the total
with fixed horizontal post-TTI V is 603 m/s
Earth-Moon line. One-way trajectories
from the EM-L2 halo to any point on the lunar
surface take about 6 days and 2500 m/s V (by LST); our paper has details.
Powered lunar swingbys at h = 100 km,
S1 from Earth & S2 to Earth
HOI = Halo orbit insertion
HOD = Halo orb.
Departure
S2 Mar. 29
210 m/s
HOI
19 m/s
Mar. 13
HOD
33 m/s
Mar. 24
MCC V’s are at changes
from red to blue near L2 on
March 10, 34 m/s and
March 27, 52 m/s
Many halo revs possible
S1 Mar. 7
255 m/s
Some work presented here was supported by “megagrant”
11.G34.31.0060 from the Russian Ministry of Education
and Science. Besides these 4 methods, others are
described in paper AAS14-470, “Trade-
off between Cost and Time in
Lunar Transfers” by
Francesco Topputo.
10
EM-L2 Halo Orbit Selection• A northern (or Class 1) halo orbit
with a relatively small Z-amplitude
of 7,000 km allows continuous
visibility with Earth and with most
sites of interest on the lunar far
side, but poorer at lunar S. Pole.
• Rather easy to transfer to other
halo orbits if necessary Selected
From Fig. 5 of
IAC-13.A5.1.4
Northern
Halo Orbits
(Class 1)
Southern
Halo Orbits
(Class 2)
Moon EM-L2
5° Horizon
mask line
from a far
southern
landing site
From Paper IAC-13.A5.1.4 presented at
the International Astronautical Congress
in Beijing in Sept. 2013, J. Hopkins, R.
Farquhar, et al. (Ref. 7)
View of the selected halo orbit
as seen from the Earth
Has more visibility →
of the lunar South Pole 11
From EM-L2 Halo Orbit
Direct to the Lunar Surface
Rotating Lunar Orbit View with fixed Trajectories near the Moon
horizontal Earth-Moon line. Red to Tsiolkovsky, Blue to S. Pole, Green to Rainer gamma
Moon
HOD
EM-L2
MCC
These take 6d from halo departure (HOD, 18 m/s
for all) to the Moon; longer might have slightly
lower Vs, given in m/s in the table to the left.
MCC is the mid-course correction described
before. The trajectory to the near-side crater Rainer- has a lunar orbit insertion (LOI) into a 10km-alt. circular arc to the target, then it uses a
“Drop” V for a nearly vertical descent to the target. For most near-side targets, the Drop V is the
main burn and the landing V is reduced (less fall time) for lower altitudes in the circular orbit arc. All
trajectories might be like the one to Rainer-, with a low lunar orbit before dropping, for nav. 12
LEO to NRHO & Small Halo V Comparison
The “Total Orion Cost” = the Total post-TTI V
from Whitley & Martinez, Options for Staging
Orbits in Cislunar Space, 2016 IEEE Aero-
space Conference, pp. 1428-1436 (Ref. 9)
V comparison in m/sec for Earth-return
trajectories to a small (7000 km Z amplitude)
thalo orbit, and to a Nearly Rectilinear Halo
Orbit (NRHO), using. powered lunar swingbys.
The Orion can easily fly either of these trajec-
tories, but other vehicles might be more limited
by the higher NRHO V. Has an abort strategy
been worked out for the NRHO like that for the
small halo in Ref. 7? The shorter period of the
NRHO may help for that. Addition of MCC’s
between the NRHO and the Moon may
decrease the total V. 13
But should we go to EM-L2 at all, or construct the
Lunar Orbiting Platform Gateway (LOP-G)?:
Moon Direct:A Coherent and Cost-Effective Plan to Enable Lunar Exploration
and Development
IAC-18,A3,2C,11
Robert Zubrin
Pioneer Astronautics
11111 W. 8th Ave. unit A
Lakewood, CO 80215
14
Alternative Options
We consider five alternative mission modes. These are:
A. Program of Record: First construct a Lunar Orbit Gateway (LOG), and then use it
as a node to send the Orion spacecraft to low lunar orbit (LLO), and then conduct the mission
to the surface via LOR, with a LEV type vehicle going from LLO to the lunar surface (LS)
and back. Orion then returns the crew to aeroentry at Earth
B. LOR-Orion: Same as option B, except no LOG is constructed.
C. LOR-Dragon: Same as option C, except a Dragon is used instead of Orion.
D. Direct Return: Dragon delivered to surface. Dragon flies directly back to TEI,
aeroentry
E. EOR (Moon Direct): Crew to orbit in Dragon. Goes to Moon in LEV.
Direct return to rendezvous with capsule in Earth orbit.
15
Comparison of Options
Option A. LOG B. LOR-Orion C.LOR-Dragon D. Direct Return E. Moon Direct
Ph 1 IMLEO 240 120 120 120 120
Ph 2 IMLEO 126 126 56 120 68
Ph 3 IMLEO 110 110 40 53 14
Total IMLEO 2692 2572 1032 1300 536
Surface % Access 3 3 3 3 42
16
Zubrin’s Conclusions
It can be seen that the Moon Direct approach is decisively the best. Its advantages include:
1. Lowest total program launch mass. (~1/2 that of closest alternative)
2. By far the lowest recurring mission launch mass. (~1/3 that of closest alternative)
3. By far the greatest exploration capability (14 times surface access as 4 km/s LOR-class
LEV)
4. No need for lunar orbit rendezvous.
There is no point going to other worlds unless we can do something useful when we get there.
Turning local materials into resources is the key.
The resourceful will inherit the stars.
17
Zubrin was not the first to criticize a station near one of the Earth-Moon
colinear libration points:
From p. 48 of Robert Farquhar’s Memoires, “Fifty Years on the Space
Frontier: Halo Orbits, Comets, Asteroids, and More”:
A space station at the Earth-Moon L1 point supporting lunar surface
operations was discussed in a novel by Arthur C. Clarke in 1961 [6].
He commented that a Moon-bound spaceship stopping at the L1 station
to pick up a passenger and some cargo would waste time and a lot of ΔV.
[6] Clarke, A. C., A Fall of Moondust, Harcourt, Brace and World, Inc.,
New York, 1961.
18
Some History and My Conclusions about LOP-G - 1
Farquhar’s idea for an EM-L2 space station, HOSS, was given in NASA
TN D-6365, “The Utilization of Halo Orbits in Advanced Lunar
Operations”, July 1971.
Farquhar advocated this idea at IAA cosmic study meetings at the IACs
in 2004 and 2008; he called it an International Exploration Station (IES).
After 2008, NASA switched from an EM-L2 orbit to a DRO for ARM.
In 2017, ARM was cancelled and NASA, remembering the IAA cosmic
studies, again became interested in EM-L2 halos, especially NRHOs.
In February-March 2018, NASA held a meeting in Denver about science
goals for the Deep Space Gateway (or DSG, as LOP-G was called then).
My impression was, there was little science discussed there that couldn’t
be performed much less expensively with robotic missions.
19
Some History and My Conclusions about LOP-G - 2
During the next several years, NASA and our international partners want to
concentrate on lunar exploration. For that, Zubrin has shown that a “Lunar Direct”
approach, without LOP-G, is significantly more effective.
I believe that something like LOP-G should be built, but with the aim of explora-
tion beyond the Moon. LOP-G is already planned to have a robust propulsion
system; just increase that to become the Deep Space Transport (DST), and that
should be its primary goal. I believe that there is no need, and we can’t afford to,
build both LOP-G and DST. But DST is certainly needed for human missions to
NEO’s and Mars, and libration point orbits provide a high-energy perch to
minimize departure & arrival Vs – see following examples. During the first years
of construction of DST, it could be used for some of the currently-envisioned
purposes of LOP-G, and that’s also possible between missions, while DST can be
“stored” in some EM-L2 halo.
As noted before, lunar comm is best handled by 3 robotic comm sats in a large EM-
L2 orbit; comm shouldn’t be a reason for LOP-G.
20
1-year Return Flyby
of Asteroid 1994 XL1 in 2022
Ecliptic plane inertial view
1994 XL1 was the first asteroid
discovered with a period (201d)
less than that of Venus. It is
estimated to be 250m across.
From EM-L2 halo back to the halo with ΔV = 432 m/sec with help
from SE-L2 and unpowered lunar swingbys, slow departure
2021 Sept 21 ITV departs EM-L2
2022 late July/
early Aug
CTV uses PhOR to change crew &
supplies at ITV
2022 Aug 11 Earth departure perigee
2022 Dec 13 1994 XL1 flyby, 14.7 km/sec
2023 July 30 Astronauts return to Earth in re-entry
capsule, or via PhOR, ITV perigee V
at h = 622 km to capture
2023 Nov 29 Uncrewed ITV returns to EM-L2 halo
BOLD = crewed portion
Earth
To Sun→
1994 XL1
1994 XL1 flyby
2022 Dec. 13
ITV
Ecliptic plane view with
fixed horizontal Sun-Earth line
• Sun
Earth
& S/C
Venus
Mercury
Flyby
21
1994 XL1 Trajectory with Return
to EM-L2 Halo Orbit
Geocentric rotating ecliptic-plane view with fixed horizontal Sun-Earth line
SE-L2A1
A3
A2
→
Earth
The Moon’s orbit is light blue with
radius 380,000 km. 3 lunar swingbys
at alt. 10,000 to 30,000 km transfer
from/to the low HEO orbits. The motion
near the Earth for orbits with apogees (A#) to the left is counter-
clockwise (direct); most perigees (P#) are close to Earth
-SE-L1
To Sun →HI
Maneuvers: ‘22Jun01, 0.9 m/s, A3; ’22July, add crew
‘21Sep21, 0.1m/s, HD = depart halo, ‘22Aug11, 180 m/s, P4 (to 1994 XL1)
‘22Jan20, 53 m/s, A1 uncrewed ‘22Dec14, 9.4 m/s, 1d after 1994 XL1
‘22Mar23, 0.2 m/s, P1 ‘23Jul30, 110 m/s, P5 capture V*
‘22Mar31, 9.9 m/s, A2 ‘23Sep19, 17 m/s, A6
‘23Nov09, 25.5 m/s, P6
‘23Nov29, 25 m/s, HI =
Halo Insertion
A6
P6
HD
Total V from, &
back to, the EM-L2
Halo is 432 m/sec
At 2005 IAC,
Howell and
Kakoi showed
similar 0 V
transfers from
EM-L2 to
SE-L1 halos.
*crew to Earth
in capsule;
ITV uncrewed
from P5 to HI
22
The Trajectory near the Moon
Moon L2
Halo
orbit
from
1994 XL1
and the
SE-L1
region
To the
SE-L2
region
To Earth
This shows the trajectory in a rotating
lunar orbit plane view with fixed horizontal
Earth-Moon line, centered on the
Earth-Moon L2 point (thus, the Moon is
shown as a short line due to the eccentricity
of its orbit). The motion in the halo orbit is
clockwise in this view, which shows the
departure from the halo orbit, and return to it
2 years and 2 months later.
Also shown are 3 lunar swingbys that
drastically changed the orbit, with the two
inbound trajectories passing above the
Moon from upper right, and the main
outbound trajectory under “Moon” from
left to lower right; farther in the lower left,
there was also a distant intermediate pass
that had only a small effect on the orbit.
A .
2022 Lunar
Swingby Dist., km:
A – Mar. 19, 23,095
B – Apr. 13, 23,269
C – July 19, 10,289
C
B
23
Phasing Orbit Rendezvous (PHOR),
CTV & ITV (DST), Slow 1994 XL1 Flyby
The period of the ITV phasing
orbits is 12 days. The
opportunities for the CTV to
rendezvous with the ITV with just
one orbit occur on dates near the
ITV perigees on 2022 July 19, July
31, and Aug. 11. The light blue
trajectory is that of the ITV, but
dark blue from the S3 lunar
swingby to the first phasing orbit
perigee on July 19, and yellow or
orange during the times when the
CTV is staying with the ITV (for 2
days) for some CTV trajectories.
The CTV trajectories are in pink
outbound and dark green for its
Earth return. The ITV last phasing
orbit perigee on Aug. 11 has the
180 m/s Oberth V to 1994 XL1
• Earth
Rotating ecliptic-plane view with fixed
horizontal Earth-Sun line
To Sun →
S3 lunar swingby 2022 July 16, distance 10,289 km
Lunar
orbit
24
Phasing Orbit Rendezvous (PhOR),
CTV & ITV/DST, Slow 1994 XL1 Flyby
There are 2 weeks with almost daily consecutive opportunities for a CTV launched from the ETR
with (in this case) an incl. 39 orbit with C3 < -1.4 (apogee just beyond the Moon) to rendezvous
with the ITV/DST for post TTI V <400 m/s. All but the last, with V >400 m/s shown with red
font, have an alternate 2-orbit solution using V’s at the 1st orbit apogee and perigee. Lowest V
direct rendezvous occurs with launch on phasing orbit perigee date (green). 1st orbit rendezvous
dates are purple, 2nd orbit dates are blue, and last orbit dates are brown.
25
1-year Return Flyby of Asteroid 1994 XL1 in 2022
with Fast Departure from EM-L2 Halo Orbit
2022 Jul 07 – HOD, ITV departs EM-L2 halo, 7.5 m/s
2022 Jul 09 – Mid-course correction, 30.0 m/s
2022 Jul 16 – powered lunar swingby to enter
phasing orbits, h = 50 km, V 198.9 m/s
2022 Jul 21 – Perigee h 2022 km, V 23.9 m/s
2022 late July/early Aug. – CTV crew PHOR with ITV
2022 Aug 8 – Apogee V 0.3 m/s targets Rper
2022 Aug 12 – Earth departure perigee, 201.8 m/s
2022 Dec 13 – 1994 XL1 flyby, 14.7 km/s
2022 Dec 15 – Earth targeting V 9.2 m/s
2023 Jul 31 – Astronauts return to Earth in re-entry
capsule, ITV capture per. V 111 m/s, h = 622 km
2023 Nov 29 – uncrewed ITV returns to EM-L2 halo,
V 50 m/s
Navigation easier with the less V method
HOD
MCC
Earth
Lunar
orbit
In halo
orbit
Rotating ecliptic-plane view
with fixed horizontal Sun-Earth line
Lunar swingby
2022 Jul 16
198 m/s
From EM-L2 halo back to the halo with V 633 m/sec
using powered lunar swingby for faster departure
26
Table of Selected Low-Cost 1-year Return Asteroid
Flyby Opportunities with Departure in 2026
The departure dates are the dates of the last perigee of the phasing orbits when the Oberth maneuver is performed to go to
the asteroid, so the actual departure from the halo orbit would generally be 4 to 6 weeks earlier, or 6 or more months if a
slow transfer, without a powered lunar swingby, is used with robotic operation. They show the rather frequent low-C3
opportunities; these are expected to increase significantly as new NEO surveys become operational. The objects are at
least 150m or more in diameter (since the albedos of these asteroids are poorly known, we give a range of diameters
based on a plausible range of albedos), and have arranged the table in order of increasing total V, which is just the sum
of the two Oberth maneuvers, the first being for departure to the asteroid and the second being for capturing the ITV back
into a HEO with perigee geocentric distance 7000 km and apogee 65 Earth radii, a little beyond the Moon’s orbit. About
500 m/s more V would be needed for the powered lunar swingbys, and the halo orbit departure and return, but if the
astronauts could rendezvous using a CTV during the phasing orbits before and after the Earth departure and return,
respectively, then the extra cost could be much less since the ITV, without crew, could be transferred from and to the EM-
L2 halo orbit using slow transfers, like those described previously. For PhOR, the departures must be near the lunar orbit
plane; 4 trajectories were removed to satisfy that constraint.
27
To 2000 SG344 Rendezvous
2029 Jun 17 – 8m/s, Leave halo orbit 2029 Sep 25 – 561 m/s, 2000 SG344 rendezvous
2029 Jun 18 – 55m/s, Mid-course V 2029 Sep 30 – 760 m/s, Leave 2000 SG344
2029 Jun 25 – 200m/s, lunar swingby 2029 Dec 25 – Pacific Ocean return, ITV perigee V 142 m/s
2029 Jul 11 – Depart 163m/s, at 2nd perigee 2030 Apr 12 – return to the EM-L2 halo orbit
Total V 1881 m/s; 2000 m/s back to EM-L2 haloThis traj., and most shown here, were calculated with high-fidelity models using the General Mission Analysis Tool (GMAT)
Rotating Ecliptic
Plane Views
with fixed horizontal
Sun-Earth line
In EM-L2 halo, just outside
lunar orbit (not shown)
Earth and
Lunar orbit
Zoomed out view; 2000 SG344
not shown after rendezvous
5d rendezvous
2000
SG344
S/C
Earth
2029 Jun 25th
lunar swingby
200 m/s
To Sun
28
2033 – To Mars From EM-L2 Halo
Heliocentric Ecliptic Plane Inertial View
2033 Feb 18 – 9m/s, Leave halo orbit 2033 Mar 23 – 4 m/s, at last phasing orbit apogee
2033 Feb 20 – 41m/s, Mid-course V 2033 Mar 27 – 358 m/s, Oberth ∆V, near last perigee
2033 Feb 27 – 202m/s, lunar swingby, h 50km 2033 Jul 19 – 605 m/s, Deep Space Maneuver
2033 Mar 04 – 13m/s, at 1st perigee 2033 Dec 01 – 1089m/s, Mars Arrival & Capture
Rather than the above, we prefer the WSB/unpowered lunar swingbys option with robotic operation until PhOR
In EM-L2 halo, just outside
lunar orbit (not shown)
Feb. 27th
Lunar
Swingby
∆V 202 m/s
To Sun
Rotating
Ecliptic
Plane View
with fixed
horizontal
Sun-Earth line
To DSM
and Mars
EarthPhasing
orbits
• Sun
Mars
Earth
Mar. 27th
Departure
358 m/s
Dec. 1st
Mars Arrival
1089 m/s
DSM 605 m/s
29
2033 – 2035, Phobos Rendezvous
Mars Arrival & Phobos Rendezvous Phobos, and then Mars, Departure
Inertial Mars Equatorial Plane Views; Mars at center; inner circle, Phobos’ orbit;
outer circle, Deimos’ orbit; capture/departure orbit apoapse distance 48 Mars radii
Ap. ∆V
76 m/s →
Subtract 1642 m/s if the ITV rendezvouses with a pre-positioned Mars Tug
that takes the astronauts to and from Phobos. 30
2035 – Return from Mars
• Sun
Nov. 22nd, 2035
Perigee
Astronauts return
In re-entry capsule
ITV ∆V 444 m/s at
radius 7000 km for
slow robotic capture
To Sun
Rotating ecliptic
plane view with fixed
horizontal
Sun-Earth line
Lunar
orbit
SE-L2 •
Earth
Heliocentric Ecliptic Plane Inertial View
Depart May 9th
V 893 m/s
Arrive Earth
Nov. 22
Apogee
HOI
2035 May 09 – 893 m/s, Periapse Departure V 2035 Nov 22 – Earth return, ITV V 444 m/s
2036 Feb 17 – Apogee, V ~45 m/s 2036 Mar 31 – Halo orbit insertion (HOI), V ~25 m/s
Mars
31
Ballistic/Hybrid Comparison Goals & Assumptions
• Computed with Mission Analysis Environment (MAnE)/Heliocentric Interplanetary
Low-Thrust Optimization Program(HILTOP)
• The goal is to minimize the mass in Earth orbit that delivers a final mass on return to
Earth orbit equal to the dry spacecraft mass plus the sample mass (58,500 kg).
• Array power at 1 AU = 150 kW with 10 kW reserved for non-propulsion purposes.
The power drops off as 1/r2 where r is the heliocentric distance in Astron. Units.
• SEP consists of 10 Hall effect thrusters, each with a max. PPU input power of
13.254 kW with Isp of 2290.18 sec, efficiency of 58.037%, and 90% duty cycle.
• Dry spacecraft mass = 58,000 kg (excludes high- and low-thrust propellant)
• Sample mass = 500 kg
• High-thrust Isp = 320 sec, velocity losses ignored
• Earth departure and return orbit = 7,000 x 414,579 km (HEO, apogee near Moon;
astronaut rendezvous with CTV). The Earth departure date for both missions were
chosen such that perigee of the Earth escape hyperbola lies within the plane of the
lunar orbit, needed for optimum linking with a trajectory from the EM-L2 halo orbit.
• Mars capture orbit is 3,696km (300 km alt.) by 163,017 km (48 Mars radii, period
8.4 days; for rendezvous with a MST)
• Ephemeris of Earth and Mars are from JPL DE430 and a JPL spice kernel (.bsp) file
for 2000 SG34432
Ballistic/Hybrid Comparison to 2000 SG344
The Hybrid mission departs the HEO with 99
metric tons, 6 less than for the Ballistic mission
33
Ballistic/Hybrid Comparison to Mars
The Hybrid mission departs the HEO with 158
metric tons, 10 less than for the Ballistic mission
With 300 kW, hybrid performance would be better. 34
Earth – Moon L2
Halo OrbitIES?? & ITV between missions
Mars Destinations
Phobos, Deimos, or
Mars surface
Phasing trajectories using
lunar gravity-assist
maneuver(s)
Crew exchange via CTV
Perigee ∆V for Earth escape
WSB transfer near SE-L1 or
SE-L2, possibly after 1 or 2
high-orbit loops
Crew Earth return via CTV Perigee ∆V to Earth phasing
orbit
Conclusions - Human ITV/DST Missions from an EM-L2
Halo Orbit to Mars and Return with Reusable Elements
Or NEA rendezvous & Departure, without
Lower 4 rectangles to the sides
via fast (can be crewed)
or slow (uncrewed, low V)
transfers
Can be fast or slow
transfer, uncrewed
(robotic operations)
Crew transfers to ITV that,
Uses periapse V
to escape Mars
Periapse V to Mars
Capture (10d orbit) & MST
rendezvous for crew exchange
ITV Earth to Mars
Small ITV periapse V raises
apoapse to Mars WSB to
move apsidal line for departure
ITV periapse V lowers
apoapse to 10d elliptical
Mars orbit
ITV Mars to Earth
MST to Mars destinationMST returns to 10d orbit
35