Artemis-2 used an elliptic departure orbit to greatly reduce the trans-lunar injection departure burn required of its Orion service module. It used the SLS core stage and SRB’s to enter a surface-grazing ellipse with a very-low apogee, and then used its ICPS second stage to raise the perigee of that ellipse out of the atmosphere with a small ICPS burn at apogee. Upon reaching the new perigee, it made another substantial ICPS burn to raise the apogee into a very extended departure ellipse. It demonstrated what I propose here!
The ICPS had just enough propellant to be disposed-of
safely, after the Orion capsule and
service module separated, after getting
onto the extended ellipse. The Orion and
its service module made one circuit about the extended ellipse, preparing for the lunar trip. Upon reaching perigee, a modest service module burn put it onto the
lunar trajectory, with propellant
reserves for course corrections both ways.
This lunar trajectory was a free-return loop around the moon, with a direct re-entry back at Earth. See Figure 1.
Figure 1 – How Artemis-2 Used Elliptic Departure to Reduce the
Lunar Departure Burn
Changes for a Permanent Capability
The perigee Artemis used was still too low at about 185 km
altitude, too close to the entry
interface altitude of 140 km, to be a
permanent orbit. A more permanent low
circular orbit would about 300 km or more,
as shown in Figure 2. An
extended elliptical departure orbit was selected by this author that had more
than 100:1 stronger pull of Earth gravity at apogee than lunar gravity, plus an integer ratio of its period to the
period of the low circular orbit! Such
would be stable for multiple circuits,
and the integer period ratio ensures that anything left behind in low
circular will be there just as you arrive back at that perigee! That makes rendezvous and docking much easier
and faster.
Figure 2 – Proposed Elliptic Departure/Arrival Orbit With Basing
in Low Circular
It would be unattractive to base directly in the extended
orbit, when its perigee speed is very
nearly Earth escape speed. The most
demanding portion of the ascent is surface-to-orbit, for which low circular is demanding
enough, just reaching circular orbit
speed with any significant payload. It
takes a much bigger launch vehicle to reach near escape speed with that same
payload! That’s more expensive, and may require dedicated designs!
Instead, one bases in
the easier-to-reach low circular, and
uses a convenient stage or vehicle as a tug,
to take its payload craft from there to the departure ellipse. From there, the departure burn demanded of the payload craft
is quite small, especially for lunar
missions. The “tug” for Artemis-2
was the ICPS second stage of the SLS launch vehicle. Using a tug to get onto a departure ellipse
is still a substantial reduction of the departure burn, even for faster-than-Hohmann Mars
missions, as the figure shows.
As we already now know,
reusability dramatically lowers mission costs! To accomplish that as a tug-assisted elliptic
departure, we need a reusable stage or
vehicle to use as our tug, and we need
to base it in low circular orbit. Such
is more easily reached for re-supply from the surface. The tug can stay on the departure
ellipse, after releasing the payload
craft to make its modest departure burn.
This tug then returns around that ellipse, and burns unladen (for low propellant
expenditure!) to return to low circular, where it is based.
Basing at a Space Station
That base ought to be an appropriate space station
located in the low circular orbit at low inclination (to reduce plane change
requirements for lunar or interplanetary missions). What we need of that station is twofold: (1) the means to assemble mission craft from
docked-together components, and (2) the
means by which to fill (and refill) such craft,
and the reusable tug, with
appropriate propellants.
We already know that we need manipulator arms and a
framework to support them, from the
space shuttle and ISS experiences. We
will need the means by which to transfer both room temperature storable
propellants, and cryogenic
propellants, from tank to tank in
zero-gee, without spinning big
structures or using orbit-changing ullage thrust!
Given that we can accomplish those things, the advantages are enormous, as detailed in Figure 3. The notation Vnear is also known
as c3, the speed with respect
to Earth needed at end of burn, close to
Earth, to accomplish the lunar or
interplanetary mission.
Figure 3 – Reusable Tug From Low Circular Greatly Reduces
Final Departure Burns
That means that our space station is a frame to which
multiple manipulator arms are affixed, with
arm operator cabins, mounting and
holding fixtures, plus the support
equipment for such crewed activities. It
must also have a multiplicity of appropriate propellant storage tanks, the plumbing for propellant transport
point-to-point on the station, and
docking facilities for the “supply tanker” transports bringing propellant
supplies to the station.
The solution to the transfer of storable propellants in
zero-gee has long been known: bladder
expulsion using gas pressure, to squeeze
the bladder within the tank walls. There
are no well-known cryogenic propellant solutions, other than large structure spin or ullage
thrust, since there are no polymers with
the enormous strain capability required (over 100% elongation), for bladder service at cryogenic
temperatures! Neither structure-spin
nor ullage thrust would be useful at a space station, for any number of reasons.
There is,
however, a not-well-known
solution involving spin, but only
spinning the propellant inside stationary tanks! One does this with vanes, driven by electric motors. If one does this by spinning half the
propellant one way, and half the other
way, then all the spin reactions and
gyroscopic forces sum to zero at the tank mountings! See Figure 4.
Figure 4 – Proposed Cryogenic Delivery Tank Using Spin of
Only the Propellant
Note also that the most practical tank design, for (only) the payload of cryogenic
propellants delivered to the station,
would be that same vane tank approach!
Storables could use bladder expulsion,
same as the main propulsion tanks.
Or they could use vane tanks, but
the bladder expulsion approach is both lighter weight and long-proven.
Now the form of the space station becomes clearer: a long truss space frame, along a portion of which are disposed a
number of storable bladder tanks and cryogenic vane tanks, and along another portion of which there are
assembly arms, holding fixtures, and arm operator cabins. Plumbing and power lines get routed within
the frame. You put the crew quarters and
re-boost propulsion at one end, and
leave the other open, for tanker vehicle
docking, and for future propellant capacity
growth. See Figure 5.
Figure 5 – Proposed Refill and Assembly Space Station
Concept
Benefits of This Approach Going to the Moon
For 1-way lunar landings, your craft only needs 0.1 km/s to depart from
the ellipse where the tug took it. It
needs around 0.9 to 1.1 km/s dV to enter low lunar orbit at some convenient
inclination. From there, about 1.7-1.9 km/s dV will land it. That’s a total of likely-under-3 km/s dV
required of your 1-way lunar landing mission craft!
At 330 s Isp for the lunar lander using storables, that is about a single-stage mass
ratio of 2.53, or a propellant mass
fraction of 60%. If the lander has rough
field capability, its inert mass
fraction might be near 15% at most.
Which leaves a payload fraction near at least 25%, with storable propellants! This is how you send cargo 1-way to build a
base there!
Astonishing!
And it’s only about 6 km/s to return to the extended
elliptic Earth orbit, all the way from
the surface of the moon, unrefilled! How to
return such a vehicle unladen of payload,
but as a single stage, is a topic
for another time. But it can be done, even with storables!
What About Mars?
For a 2-way Mars orbit-to-orbit transport using low Mars
orbit, you only need around 1.5 km/s
to depart the ellipse, and about 2 km/s
to enter low Mars orbit, 1-way. That’s only
3.5 km/s required 1-way, and so only 7
km/s required for the complete 2-way trip,
all unrefilled in a single stage!
Which would amount to a reusable chemically-powered orbit-to-orbit
transport, even with zero infrastructure
at Mars! That’s a single-stage
mass ratio of about 8.7 at only 330 s Isp for storables, or a propellant fraction of 89%. If the inert fraction for a vacuum-only ship is
5%, the payload fraction could be
6%! And with only storable
propellants in a single stage!
Astonishing!
To deliver payloads 1-way to Mars for direct entry and
landing, you are looking at about 1.5
km/s to depart, plus some modest course
corrections, and the final landing burn
of 1 to 1.5 km/s. That’s at most 3 km/s
dV demanded of the craft. At 330 s Isp
with storables, that’s a mass ratio of
only 2.53, for a propellant fraction of
about 60%. If the lander, which must survive entry as well as be
configured for rough field landing, has
an inert fraction of 20%, that is still near
20% payload fraction, even with only storable
propellants! That could well be how to
send cargo 1-way to build a base there!
Even more astonishing!
See also Table 1.
What About Practical Tug Designs?
Now, what do we need
of the tug? Assuming it might have to
travel the extended ellipse for 2, maybe
3, circuits, that’s about 10-12 days’ time in space. If it uses cryogenic propellants for their
high performance, it needs a useful
“stage life” without serious evaporative loss, of only some 10 or 12
days! That rules out common bulkheads
between LOX and LH2 tanks, and it rules
out bare single-wall tank shells exposed to sunlight in space! But it does NOT rule out using LOX-LH2
for its highest performance! You
just need separate LOX and LH2 tanks, which
must be well-insulated externally, plus
with a shiny foil outer covering to shade them from sunlight heating. That’s probably closer to a stage inert
fraction of 10% than the usual vacuum stage inert of 5%.
We do not need months or years of “stage life”, only a couple of weeks, or so!
Which means we really do not need the weight and power-required
penalties of cryocooler equipment! We
just need a minimal tank redesign from what is otherwise basically Centaur
stage technology. And later
on, we might need to scale it up for
larger-mass mission craft!
See Figure 6.
Figure 6 – Probable Tug Tank Construction for LOX-LH2
Propellants and 2-Week “Stage Life”
What About Arrival Versus Departure?
The dV requirements for arrival are almost exactly the
same as departure, but the timing
requirements are different!
It only takes several seconds for the tug to undock from its payload
craft and move several meters away. That
means for departures, the tug can feasibly
fire for reaching ellipse perigee speed,
undock, and let the craft fire
for its departure, all in the one pass! The tug then makes 1 circuit about the
ellipse before burning at next perigee unladen, to get back into low
circular, going back to the station.
Or, if mission
preparation time is needed by the payload craft, both tug and craft can make one circuit about
the departure ellipse, with the craft
departing, and the tug burning to return
to circular, at the next perigee.
Arrival is different:
the tug cannot “be there” just as the craft arrives and burns
into the ellipse perigee! The
craft needs to make a circuit about the ellipse, before the tug can (1) rendezvous with it at its
next ellipse perigee, and (2) the tug
must then burn to get onto the ellipse with the craft. It then takes significant time to actually
get docked together. So the docked pair
must make a second payload craft circuit about the ellipse, before the tug can burn, to put them into low circular, and take them right to the station.
And the laden vs unladen weight statements are more beneficial
for departure, and not as beneficial for
arrivals. You’d like the bigger
propellant burn to be the first, but
that cannot happen when using a tug to assist arrival. Arrival retrieved craft sizes must then inherently
be smaller, for a given tug design. See Figure 7 for the estimated
departure and arrival data for a tug sized to put a 50 metric ton craft onto
the departure ellipse defined above, as
done with simple linked rocket equation calculations in a spreadsheet.
Figure 7 -- Typical
Tug Rough-Sizing and Performance by Spreadsheet
About Resupply Tankers
The resupply tanker vehicles sent to the station from the
surface could be the upper stages or payload items of almost any existing or
planned launch vehicle! If sending up a
storable propellant, the payload can be
a simple bladder-expulsion tank, quite
separate from the vehicle’s main propulsion propellant. If sending up a cryogenic propellant, the payload needs to be a vane tank, also quite separate from the main propulsion
propellant.
Done that way, all
the tanker vehicle need do with the space station is dock, and then hook up to the station
plumbing, to deliver its payload. It goes without saying that reusable tanker
vehicles would be much preferred for the long term.
The main caveat for tanker vehicles is that you do not want
to use the same payload propellant tank (of either type) for shipping different
propellants! Propellants do get into the
inherent slight porosity, of even metal
tank walls! You need to be able to swap
out the payload propellant tank, if the
same vehicle upper stage gets re-used to deliver other propellant species! Period! That nevertheless could be something quite
convenient, whether the payload tank is
a vane tank, or a bladder-expulsion tank! See Figure 8.
Figure 8 – Plausible Tanker Supply Vehicle Configurations
Prior Related Postings
As you can see from the list below, I have been thinking about the various
aspects of, and problems feeding
requirements into, this tug-assisted
elliptic departure space station scenario, for some time.
The dates are shown in MM-DD-YYYY format. There is an archive search tool on the left
side of this page. All that you need in
order to use it are the year,
month, and title. I suggest that you jot down the ones you
would like to see. Click on the
year, then the month, then the title if there were multiple postings
that month.
This site also has a keyword search option, in the sense that if you select a keyword, then you see only those postings that are
labeled with that keyword. The
current list of keywords is: aerothermo, airplanes,
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Only some of the more recent postings have received a search
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is something I have to give you. It is
the posting date in DDMMYYYY format.
Date............Keywords...............................Title
11-11-2025..space program.......................Where Should the New Space Stations Be Located?
10-16-2025..space program.......................Going Back to the Moon
7-26-2025....space program.......................Tank Design for Easy Cryogenic Transfers In Weightlessness
5-01-2025....space program.......................Vehicle Assembly and Refueling Facility in LEO
1-25-2025....space program.......................Initial Study for Tug Missions LEO to LLO
1-2-2025......launch, space program...........SpaceX’s ‘Starship’ As a Space Tug
12-1-2024....space program........................Tug-Assisted Arrivals and Departures
10-1-2024....space program........................Elliptic Capture
12-9-23........launch, space program............Overall Study Results: Propellant
From Moon
5-1-22..........Mars, space program..............Investigation: “Big Ship” Propellant From Moon vs From Earth
4-2-22..........Mars, space program..............Earth-Mars Orbit-to-Orbit Transport Propulsion Studies
2-1-22..........space program........................A Concept for an On-Orbit Propellant Depot
8-18-21........launch, space program...........Propellant Ullage Problem and
Solutions
3-23-21........space program........................Third Spacex Tanker
Study
3-21-21........space program........................Second Spacex Tanker
Study
3-17-21........space program........................Spacex Tanker
Investigation
7-3-20..........launch,
space program............Cis-Lunar Orbits and Requirements
11-21-19......Mars, space program...............Interplanetary Trajectories and Requirements
2-11-14........Mars, space prgm, spacesuit..On-Orbit Repair and Assembly Facility
10-2-13........space program.........................Budget Moon Missions
8-2-12..........Mars, space program...............Velocity Requirements for Mars
Orbit-Orbit Missions
8-2-11..........space program..........................End of an Era Need Not Be End of a Capability
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