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

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.

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

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