Staging from Earth-Moon L-2 Orbits - Gateway or Tollbooth?
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Staging from Earth-Moon L-2 Orbits -
Gateway or Tollbooth?
David W. Dunham, KinetX Aerospace, Inc.
david.dunham@kinetx.com
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
Robert W.
Farquhar
1932 – 2015
Master Celestial
Mechanician,
Father of Halo Orbits
and Asteroid
Exploration (NEAR-
Published in: Farquhar, R. W., “Future Missions
for Libration-Point Satellites,” Astronautics &
Shoemaker to Eros)
Aeronautics, Vol. 7, No. 5, pp. 52-56, May 1969.
2
Human Exploration of the Moon, Near-Earth
Asteroids, and Mars using Staging from Earth-
Moon L-2 Orbits and Phasing Orbit Rendezvous
My presentation is
largely taken from
this one.
David W. Dunham, KinetX Aerospace, Inc.
david.dunham@kinetx.com
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 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
NRHO
• 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 5
Cargo Mission Possibility to EM-L2 Halo:
Outward Lunar Swingby & SE-L1 WSB Transfer
Rotating ecliptic-plane view with
Horizontal Sun-Earth line
• The vehicle only has to launch into a
trajectory just reaching the Moon’s orbit
rather than launching to Sun-Earth L1
Lunar HOI distances to reach the WSB
orbit
swingby
• The vehicle could be launched into
• SE-L1 phasing orbits before the lunar swingby,
Earth allowing use of less V to correct launch
To Sun 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
TTI (from LEO) ∆V 3152 m/s Genova, NASA Ames
Post-TTI V: 21 m/s at apogee WSB and 5.4 m/s
halo orbit insertion (26 m/s total); Perigee Jan 13, • Robert Farquhar conceived many of the orbit
Halo insert July 18, 2018, TOF = 173 days ideas shown here
(apogee = March 26, 2018). Lunar swingby
• Farquhar’s Memoirs, “Fifty Years on the Space
altitude 9700 km.
Frontier: Halo Orbits, Comets, Asteroids, and
More” are available on amazon.com
6
LunaH-Map Transfer to Lunar Orbit
A low-thrust cubesat (from EM-1) example of the
Thrusting shown in
previously-shown trajectory by Anthony Genova
RED F
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 C B
weak capture around Earth); {L+80 hrs} D A
E) Begin 4-day Thrust Arc (anti-velocity) {L+156
hrs}
F) Apogee at 1 million km altitude (no maneuver); E
G
{L+ 34 days}
G) Begin 5-day Thrust Arc (anti-velocity) to target H
Moon and decrease approach speed {L+
64 days} Trajectory shown in Earth Inertial
H) Weak Capture into Lunar Orbit {L+ 69 days} Frame 7
Fast Transfers to the Earth-Moon L2 Point
Trajectories 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
Some work presented here was supported by “megagrant” Powered lunar swingbys at h = 100 km,
11.G34.31.0060 from the Russian Ministry of Education S1 from Earth & S2 to Earth
and Science. Besides these 4 methods, others are
HOI = Halo orbit insertion
described in paper AAS14-470, “Trade-
• Mission to EM-L2 halo via powered lunar
off between Cost and Time in
HOD = Halo orb.
Lunar Transfers” by
swingby
Francesco Topputo.
Departure
– CTV post-injection V = 308 m/sec (& 295
S2 Mar. 29
Return 210 m/s
2021 April 2 with
HOD
atmospheric m/sec to return) 33 m/s
•
re-entry capsule
L2
Mar. 24
EarthLaunch Moon •
2021 March 3 S1 Mar. 7 HOI
V 3,129 m/s 255 m/s 19 m/s
from LEO Mar. 13
MCC V’s are at changes
Rotating lunar orbit plane plot With return, the total from red to blue near L2 on
with fixed horizontal post-TTI V is 603 m/s March 10, 34 m/s and
Earth-Moon line. One-way trajectories March 27, 52 m/s
Many halo revs possible
from the EM-L2 halo to any point on the lunar 10
surface take about 6 days and 2500 m/s V (by LST); our paper has details. 10EM-L2 Halo Orbit Selection
• A northern (or Class 1) halo orbit
with a relatively small Z-amplitude
Northern
of 7,000 km allows continuous
Halo Orbits visibility with Earth and with most
(Class 1)
sites of interest on the lunar far
side, but poorer at lunar S. Pole.
• Rather easy to transfer to other
Selected halo orbits if necessary
Moon EM-L2
5° Horizon
mask line
from a far
southern
landing site View of the selected halo orbit
Southern as seen from the Earth
Halo Orbits
(Class 2) From Paper IAC-13.A5.1.4 presented at
From Fig. 5 of the International Astronautical Congress
Has more visibility → IAC-13.A5.1.4 in Beijing in Sept. 2013, J. Hopkins, R.
of the lunar South Pole Farquhar, et al. (Ref. 7) 11From EM-L2 Halo Orbit
Direct to the Lunar Surface
HOD
Moon EM-L2
MCC
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
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. 12LEO to NRHO & Small Halo V Comparison
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
The “Total Orion Cost” = the Total post-TTI V been worked out for the NRHO like that for the
from Whitley & Martinez, Options for Staging small halo in Ref. 7? The shorter period of the
NRHO may help for that. Addition of MCC’s
Orbits in Cislunar Space, 2016 IEEE Aero- between the NRHO and the Moon may
space Conference, pp. 1428-1436 (Ref. 9) decrease the total V. 13But 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
14Alternative 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.
15Comparison 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
16Zubrin’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.
17Zubrin 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.
18Some 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.
19Some 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.
201-year Return Flyby
of Asteroid 1994 XL1 in 2022
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/ CTV uses PhOR to change crew &
Ecliptic plane view with
fixed horizontal Sun-Earth line early Aug supplies at ITV
2022 Aug 11 Earth departure perigee
1994 XL1 flyby 2022 Dec 13 1994 XL1 flyby, 14.7 km/sec
2022 Dec. 13 2023 July 30 Astronauts return to Earth in re-entry
Flyby capsule, or via PhOR, ITV perigee V
Earth
at h = 622 km to capture
2023 Nov 29 Uncrewed ITV returns to EM-L2 halo
& S/C
BOLD = crewed portion
ITV Venus
Mercury Ecliptic plane inertial view
To Sun→
Earth • Sun 1994 XL1 was the first asteroid
discovered with a period (201d)
less than that of Venus. It is
1994 XL1 estimated to be 250m across.
211994 XL1 Trajectory with Return
to EM-L2 Halo Orbit
Geocentric rotating ecliptic-plane view with fixed horizontal Sun-Earth line
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 P6 ‘23Sep19, 17 m/s, A6
‘23Nov09, 25.5 m/s, P6
‘23Nov29, 25 m/s, HI =
→ Halo Insertion
SE-L2 A2 SE-L1
-
A1 Earth HD
*crew to Earth To Sun →
A3
HI in capsule;
At 2005 IAC,
ITV uncrewed
Howell and
from P5 to HI
Kakoi showed
similar 0 V
A6 transfers from
The Moon’s orbit is light blue with
Total V from, &
EM-L2 to
radius 380,000 km. 3 lunar swingbys SE-L1 halos.
at alt. 10,000 to 30,000 km transfer
back to, the EM-L2 from/to the low HEO orbits. The motion
Halo is 432 m/sec near the Earth for orbits with apogees (A#) to the left is counter-
clockwise (direct); most perigees (P#) are close to Earth
22The Trajectory near the Moon
from This shows the trajectory in a rotating
1994 XL1 lunar orbit plane view with fixed horizontal
and the Earth-Moon line, centered on the
SE-L1 Earth-Moon L2 point (thus, the Moon is
region 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
To Earth A 2 years and 2 months later.
C
Moon L2
.
B Also shown are 3 lunar swingbys that
Halo
drastically changed the orbit, with the two
inbound trajectories passing above the
orbit
Moon from upper right, and the main
2022 Lunar outbound trajectory under “Moon” from
Swingby Dist., km: left to lower right; farther in the lower left,
A – Mar. 19, 23,095 To the there was also a distant intermediate pass
B – Apr. 13, 23,269 SE-L2 that had only a small effect on the orbit.
C – July 19, 10,289 region
23Phasing Orbit Rendezvous (PHOR),
CTV & ITV (DST), Slow 1994 XL1 Flyby
The period of the ITV phasing
Rotating ecliptic-plane view with fixed orbits is 12 days. The
horizontal Earth-Sun line 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
Lunar To Sun → 31, and Aug. 11. The light blue
trajectory is that of the ITV, but
orbit dark blue from the S3 lunar
swingby to the first phasing orbit
• Earth 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
S3 lunar swingby 2022 July 16, distance 10,289 km
24Phasing 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 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.
251-year Return Flyby of Asteroid 1994 XL1 in 2022
with Fast Departure from EM-L2 Halo Orbit
From EM-L2 halo back to the halo with V 633 m/sec
using powered lunar swingby for faster departure
2022 Jul 07 – HOD, ITV departs EM-L2 halo, 7.5 m/s
HOD 2022 Jul 09 – Mid-course correction, 30.0 m/s
In halo
MCC orbit 2022 Jul 16 – powered lunar swingby to enter
Lunar
orbit 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
Earth 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
Lunar swingby 2022 Dec 15 – Earth targeting V 9.2 m/s
2022 Jul 16 2023 Jul 31 – Astronauts return to Earth in re-entry
198 m/s
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
Rotating ecliptic-plane view
with fixed horizontal Sun-Earth line Navigation easier with the less V method
26Table 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.
27To 2000 SG344 Rendezvous
Rotating Ecliptic
Plane Views
2000
with fixed horizontal
Sun-Earth line SG344
2029 Jun 25th
lunar swingby Earth and
To Sun
Earth
200 m/s Lunar orbit
S/C
In EM-L2 halo, just outside 5d rendezvous
lunar orbit (not shown ) Zoomed out view; 2000 SG344
not shown after 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 halo
This traj., and most shown here, were calculated with high-fidelity models using the General Mission Analysis Tool (GMAT)
282033 – To Mars From EM-L2 Halo
To DSM
and Mars Dec. 1st
Mars Arrival
1089 m/s
Earth
Feb. 27th
Lunar
Mar. 27th
• Sun
Swingby
Departure
∆V 202 m/s
358 m/s
Phasing Earth
To Sun orbits
Rotating
Ecliptic
Plane View In EM-L2 halo, just outside Mars DSM 605 m/s
with fixed lunar orbit (not shown)
horizontal
Heliocentric Ecliptic Plane Inertial View
Sun-Earth line
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
292033 – 2035, Phobos Rendezvous
Ap. ∆V
76 m/s →
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
Subtract 1642 m/s if the ITV rendezvouses with a pre-positioned Mars Tug
that takes the astronauts to and from Phobos. 302035 – Return from Mars
Nov. 22nd, 2035 Mars Arrive Earth
Perigee Nov. 22
Astronauts return
In re-entry capsule
ITV ∆V 444 m/s at
radius 7000 km for HOI
slow robotic capture
SE-L2 •
To Sun • Sun
Lunar Apogee
Rotating ecliptic orbit
plane view with fixed Earth
horizontal
Sun-Earth line Depart May 9th
V 893 m/s
Heliocentric Ecliptic Plane Inertial View
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
31Ballistic/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 SG344
32Ballistic/Hybrid Comparison to 2000 SG344
The Hybrid mission departs the HEO with 99
metric tons, 6 less than for the Ballistic mission
33Ballistic/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
Conclusions - Human ITV/DST Missions from an EM-L2
Halo Orbit to Mars and Return with Reusable Elements
Can be fast or slow Earth – Moon L2 via fast (can be crewed)
transfer, uncrewed Halo Orbit or slow (uncrewed, low V)
IES?? & ITV between missions
(robotic operations) transfers
WSB transfer near SE-L1 or Phasing trajectories using
SE-L2, possibly after 1 or 2 lunar gravity-assist
high-orbit loops maneuver(s)
Crew Earth return via CTV Crew exchange via CTV
Perigee ∆V to Earth phasing
orbit Perigee ∆V for Earth escape
ITV Mars to Earth ITV Earth to Mars
Crew transfers to ITV that, Periapse V to Mars
Uses periapse V Capture (10d orbit) & MST
to escape Mars rendezvous for crew exchange
MST returns to 10d orbit MST to Mars destination
Mars
ITV periapse V lowers Destinations Small ITV periapse V raises
apoapse to 10d elliptical Phobos, Deimos, or apoapse to Mars WSB to
Mars orbit Mars surface move apsidal line for departure
Or NEA rendezvous & Departure, without
Lower 4 rectangles to the sides
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