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Artemis Discussion Thread

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On 11/8/2019 at 1:18 AM, tater said:

 ... you keep acting as if that is certain to be the site, and that the landers will have RTGs (which are fairly weak, anyway)

An RTG and lithium battery can do quite a lot these days, I am sure it would be adequate for a lander to power electronics, and a small microwave for cooking.  It could potentially provide passive heating as well.

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7 hours ago, sevenperforce said:

For reusing an Artemis component like Boeing's ascent module, there are four basic possibilities.

Capsule Reuse. In this approach, the only thing that is carried over between missions is the actual pressure vessel, backup power, LS, and associated hardware. All propulsion, propellant tankage, propellant, and primary power elements are delivered to Gateway and mated to the capsule at the start of the sortie; this carries the capsule down to the lunar surface (with or without any number of additional stages) and returns it to LOP-G before being jettisoned and discarded. The advantage here is that you have the safety assurance of a brand new propulsion unit for every launch and you do not need to develop any new docking systems or propellant transfer; the disadvantage is that it is the least mass-efficient approach.

Capsule and Propulsion Reuse. Here, the capsule includes reusable propulsion and RCS, but the tanks are discarded at the end of every mission. This requires the development of dockable/mateable tanks which can expend their propellant and then be jettisoned. It is advantageous because you do not need to develop propellant transfer and you do not throw away perfectly good engines after every mission so you save mass. It is also the most mass-efficient approach, in a way. The difficulty in developing such tanks is the primary disadvantage here, though end-of-life considerations for the engines are another concern.

Descent Module Refilling. In this version, the entire capsule is reused, along with propulsion system and tanks, and the ascent module's tanks are refilled from excess capacity on the descent module. This has the advantage of only requiring a single connection event during mission construction (structural and for prop transfer) but does require full-scale propellant transfer development. It also is less mass-efficient with respect to the descent module, which is then required to carry the weight of empty tanks. If the descent module is multistage (e.g., if there is a transfer stage as well), then using the transfer stage props is ideal because that is jettisoned before descent.

Logistics Module Refilling. This seems to be the version that NASA favors, though I don't know why.

Gotta say I I like capsule re-use only.  Far safer to not rely on multiple engine re-use.  How to dispose of the engine assembly at LOP-G?  I am sure Boeing has not designed for it, and would have opted for replacing the entire AV at the end of the engine life.

 

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Yeah, disposal is troublesome, and again, what's the ballpark of mass in LEO vs mass at lunar orbit? ~25%? So if you need 15 tons of props at Gateway, you first have to get ~60t of props to LEO? That's gotta get there regardless, mind you, but tank mass is improved with fewer tanks.

That or they build an ion tug, and let it take a zillion days to get there.

44 minutes ago, jinnantonix said:

An RTG and lithium battery can do quite a lot these days, I am sure it would be adequate for a lander to power electronics, and a small microwave for cooking.  It could potentially provide passive heating as well.

RTGs have uses, but for a serious polar moon base, they want kilopower or something similar. ISRU is going to suck power if they try it.

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14 hours ago, jinnantonix said:

Gotta say I I like capsule re-use only.  Far safer to not rely on multiple engine re-use.  How to dispose of the engine assembly at LOP-G?  I am sure Boeing has not designed for it, and would have opted for replacing the entire AV at the end of the engine life.

 

There are arbitrarily low-cost disposal trajectories from Gateway, if you don't mind it taking a long time. Once the capsule docks to LOP-G, you separate the propulsion unit and burn RCS to a disposal trajectory.

14 hours ago, jinnantonix said:

According to wiki the upper stage (tank) should be 4.0t dry mass, 111.5t wet mass. plus a fairing base , let's say 500kg.  I will tweak the values in SSTU and expect now that I should be able to model boosting a 37t payload from LEO to NRHO.

Yeah, SSTU is simply wrong. The tankage dry mass is accurate. I have seen 111.5 tonnes quoted as propellant mass and as wet mass, depending on source. 

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The Lockheed ascent module definitely seems to have potential for crew module reuse, since the tankage and crew section are visibly separate.

G-flHuQi.jpeg 

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On 11/9/2019 at 5:42 AM, sevenperforce said:

I confess I am very fond of Boeing's lander design. Putting the airlock on the descent stage and the engines on the sides allows it to comanifest flatpacked cargo to the lunar surface or act as an independent cargo lander that doubles as a surface asset (showers, anyone?). It also permits the ascent stage to be limited just to what is needed for return.

I think there is a bit of Werner Von Braun and Das Mars Projekt in the Boeing design.  Aerodynamics and outboard engines aside, the Mars Excursion Module has similar features built into the descent vehicle.

K7q4PgA.png

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10 hours ago, Barzon said:

the tankage and crew section are visibly separate.

Though is connected with struts and pipes, the RCS are placed on the tank, and both can be covered with an insulation blanket.

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As @sevenperforce advises it seems feasible to use 2 x FHe launch vehicles, and deliver a 37 t payload to NRHO with a naked FHe upper stage. The SSTU and Kerbalized SpaceX mods need a bit of tweaking to match the published Falcon Heavy specs, but after some effort I managed to model the transit successfully.  The launch sequence is similar to the Gemini 8 (Agena docking test) mission, as follows:

  1. The first FHe launches the lander craft to LEO
  2. 90 minutes after first launch, the second FHe (no payload) launches the naked FHe upper stage to LEO (into lunar plane)
  3. The first FHe and lander use its FHe upper stage to rendezvous with the 2nd launch upper stage (2 burns)
  4. The two craft dock, and the spent first launch upper stage de-orbits into Earths atmosphere
  5. The combined lander and 2nd upper stage craft boost to TLI
  6.  3 days later, the craft burns twice to rendezvous with LOP-G
  7. The FHe upper stage separates, and de-orbits for crash on the lunar surface
  8. The lander craft docks with the LOP-G.

The lander needs two pairs of radial drop tanks and two inline drop tanks, (wet mass = 32t) as below:

  1. Radial tank pair 1:  NRHO - LLO
  2. Radial tank pair 2:  LLO - Surface
  3. Inline drop tank 1:  Surface -LLO
  4. Inline drop tank 2:  LLO - NRHO

 

0eqVzMV.png

 

The re-usable capsule remains at LOP-G after the first mission , the expendable assembly for subsequent missions is as below, and has a wet mass = 30t.

 

QLQf7pl.png

 

Comanifesting a Cygnus resupply module increases the payload to 37t.  The Cygnus provides logistics, and also the avionics and communications to complete the rendezvous at LOP-G.  It all fits snugly into a standard FHe fairing.

 

Pc0gKpW.png

 

At the end of the mission the spent Cygnus module docks with the last inline drop tank on the lander, undocks, and then de-orbits to crash on the lunar surface. 

 

Edited by jinnantonix

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On 11/8/2019 at 11:25 PM, tater said:

That or they build an ion tug, and let it take a zillion days to get there.

Ions are the way to go if you want to break the rocket equation (you can bring temperature stable [non-boiloff] fuels via engines with four-digit ISPs).

For lunar purposes, I'd assume that ions would be largely considered as practice for Mars and beyond, as it takes "a zillion days" to reach escape velocity (mostly spiraling out to the Moon and then using the Moon to switch to an elliptical "orbit" into escape velocity) then a similar amount of time to get where you are going as a chemical rocket.

To get funding, anyone committed to a "Lunar Gateway Outpost" should accept that such a thing makes a lot more sense if it contains fuel delivered via ions.  Just don't expect to launch an ion and get there before 2024, they are for more patient astronauts.

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5 minutes ago, wumpus said:

To get funding, anyone committed to a "Lunar Gateway Outpost" should accept that such a thing makes a lot more sense if it contains fuel delivered via ions.  Just don't expect to launch an ion and get there before 2024, they are for more patient astronauts.

Boil off becomes an issue, then.

 

I suppose IVF could power the ions.

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Just want to point out that the plan of two falcon heavy launches 90 minutes apart is currently infeasible. It would require 2 falcon heavy Launch pads. SLC 40 could be modified to support fh but it might be easier to launch one of the payloads on a new Glenn or something.

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21 minutes ago, Ultimate Steve said:

Just want to point out that the plan of two falcon heavy launches 90 minutes apart is currently infeasible. It would require 2 falcon heavy Launch pads. SLC 40 could be modified to support fh but it might be easier to launch one of the payloads on a new Glenn or something.

90 minutes would require a lot of margin since KSC is not equatorial. Realistically, the best they would do is about 24 hours (+- phased to the orbit, use ISS launch windows as an example, miss the window today, and you have another shot tomorrow several minutes different from today).

Any distributed architecture either needs really large boil off margins, so that elements can be placed in orbit weeks apart, or storable props. Note that this includes a distributed architecture using 2 SLS launches. The idea that they could send a lander and Orion in 2 SLS flights not separated by at least months seems absurd to me.

What architectures are possible with F9e and FHe? Those could literally be on their pads at the same time (of FHe plus DIVH, Atlas V, Vulcan, NG, etc).

Edited by tater

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1 hour ago, tater said:

Note that this includes a distributed architecture using 2 SLS launches. The idea that they could send a lander and Orion in 2 SLS flights not separated by at least months seems absurd to me.

There are two mobile launchers.

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14 minutes ago, jadebenn said:

There are two mobile launchers.

And one pad. What's the time for pad refurb? The most launches ever within a single year at pad 39A was actually 12 from SpaceX. The best Shuttle did from that pad was 9 in a year (best was 6 from 39B). Tough to tease out pad repair vs turn around, etc, but it seems like beating a month would be record breaking. You'd think a couple weeks would be the bare minimum. (SLS will use 39B).

Also, is the SLS program really paying enough workers to stack and ready 2 vehicles at the same time, then move and prep them for launch a few days apart? As it is I thought that even to make 2 in a year would require a substantial increase in money (presumably an increase that then goes forward in perpetuity). What is the nominal timeframe for an SLS launch campaign? The last Shuttle mission had Atlantis moving to the pad on May 31st, for a July 8th liftoff. That's probably not the shortest it could be on the pad, but they need to check everything out, test the GSE, etc, that all takes time---and that was independent of any repairs from the previous Shuttle launch.

Edited by tater

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3 hours ago, Ultimate Steve said:

Just want to point out that the plan of two falcon heavy launches 90 minutes apart is currently infeasible. It would require 2 falcon heavy Launch pads. SLC 40 could be modified to support fh but it might be easier to launch one of the payloads on a new Glenn or something.

New Glenn doesn't exist now either.  But no reason why NG wouldn't work for the lunar lander launch and rendezvous with the FHe upper stage.  Alternatively a second FH capable pad by 2024 isn't out of the question. Both are better options that 2 x SLS launches.

2 hours ago, tater said:

90 minutes would require a lot of margin since KSC is not equatorial. Realistically, the best they would do is about 24 hours

I tried both, and they both worked OK.    Since SLS is hydrogen/LOX, the time to transit depends on when the Orion is ready for TLI.  We are basically talking three launches, very  close to each other.

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11 minutes ago, jinnantonix said:

I tried both, and they both worked OK.    Since SLS is hydrogen/LOX, the time to transit depends on when the Orion is ready for TLI.  We are basically talking three launches, very  close to each other.

I'm still of a mindset that says that EoR would be best spread over longer time frames, even if it results in more required launches.

We've all watched enough launches to know how likely scrubs are.

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Just now, tater said:

I'm still of a mindset that says that EoR would be best spread over longer time frames, even if it results in more required launches.

We've all watched enough launches to know how likely scrubs are.

Of course.  And both FH and NG launched vehicle could potentially be in orbit for weeks before SLS/Orion is ready to go.  But not months.  I think that is a feasible architecture.

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4 minutes ago, jinnantonix said:

Of course.  And both FH and NG launched vehicle could potentially be in orbit for weeks before SLS/Orion is ready to go.  But not months.  I think that is a feasible architecture.

I'm not disagreeing at all, I was thinking more of the SLS-specific example (not yours, Boeing's).

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10 hours ago, jinnantonix said:

As @sevenperforce advises it seems feasible to use 2 x FHe launch vehicles, and deliver a 37 t payload to NRHO with a naked FHe upper stage. The SSTU and Kerbalized SpaceX mods need a bit of tweaking to match the published Falcon Heavy specs, but after some effort I managed to model the transit successfully.  The launch sequence is similar to the Gemini 8 (Agena docking test) mission, as follows:

  1. The first FHe launches the lander craft to LEO
  2. 90 minutes after first launch, the second FHe (no payload) launches the naked FHe upper stage to LEO (into lunar plane)
  3. The first FHe and lander use its FHe upper stage to rendezvous with the 2nd launch upper stage (2 burns)
  4. The two craft dock, and the spent first launch upper stage de-orbits into Earths atmosphere
  5. The combined lander and 2nd upper stage craft boost to TLI
  6.  3 days later, the craft burns twice to rendezvous with LOP-G
  7. The FHe upper stage separates, and de-orbits for crash on the lunar surface
  8. The lander craft docks with the LOP-G.

 

Looks excellent! Really great work. A few questions/comments:

The MVac can make very accurate terminal burns, but it cannot downthrottle below 39% and so very very small burns are not feasible. For example, while the "naked" FHe upper stage can launch into an orbit that very very closely matches the original stack, the original stack would then need to rendezvous and dock under RCS. Accordingly, you would jettison and deorbit the first upper stage immediately after SECO, rather than trying to use it to perform the rendezvous. That helps with timing, incidentally. The lander is 100% storables, so there's no real penalty to launching the lander stack on FHe one day, letting it loiter for a week or a month, and then sending the FHUS tug at leisure. Then, once FHUS delivers to LOP-G and deorbits itself, the lander stack can sit docked to LOP-G indefinitely while SLS is set up and launched.

But back to the question: what are the approximate burn times for each MVac ignition you use? I have seen very brief burns on webcasts -- under 2 or 3 seconds -- but those are usually MRS. If the burns are briefer, it could be a problem. What residuals does the FHUS have after the NRHO insertion burn? There will be boil-off, but if it has sufficient residuals at that point then we know boil-off isn't a problem.

What is the end mass of the reusable crew module?

What are your thoughts, if any, on airlocks and egress? I've been trying to come up with a way to allow a docking ring that opens to a lower chamber for an airlock and then also serves as the ingress/egress port from LOP-G, since LOP-G is orientation-agnostic. Reasoning being: we already have structural androgynous pass-through docking ports, but we'd need to design propellant-transferring structural non-pass docking connectors. It would be easier to design connectors which only transfer propellant and are not structural. Then again you'd end up with more complicated sequences at LOP-G.

What downmass capability does the landing stage have in this configuration?

If the ascent stage (with or without drop tanks) is used as a one-way cargo lander direct from TLI, what kind of capacity does it have?

2 hours ago, tater said:

I'm still of a mindset that says that EoR would be best spread over longer time frames, even if it results in more required launches.

We've all watched enough launches to know how likely scrubs are.

All the more reason to build something in EoR that is ultimately storable.

Back to the OP -- can a reusable Falcon 9 send the crew cabin and drop tank to TLI on its own?

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12 hours ago, sevenperforce said:

The MVac can make very accurate terminal burns, but it cannot downthrottle below 39% and so very very small burns are not feasible. For example, while the "naked" FHe upper stage can launch into an orbit that very very closely matches the original stack, the original stack would then need to rendezvous and dock under RCS. Accordingly, you would jettison and deorbit the first upper stage immediately after SECO, rather than trying to use it to perform the rendezvous. That helps with timing, incidentally. The lander is 100% storables, so there's no real penalty to launching the lander stack on FHe one day, letting it loiter for a week or a month, and then sending the FHUS tug at leisure. Then, once FHUS delivers to LOP-G and deorbits itself, the lander stack can sit docked to LOP-G indefinitely while SLS is set up and launched.

But back to the question: what are the approximate burn times for each MVac ignition you use? I have seen very brief burns on webcasts -- under 2 or 3 seconds -- but those are usually MRS. If the burns are briefer, it could be a problem.

The MVac burns are as follows: 

1.  FHe Launch 1:  The Lander Launch requires full throttle burn to get to LEO, it then need a three short burns at minimum (39%) throttle of about 30 m/s (6 secs) each to rendezvous in LEO.  The upper stage is undocked and burns to suborbital, and the Cygnus provides the engines and RCS thrusters to complete the rendezvous in LEO.

2.  FHe Launch 2:  The naked FH upper stage has a full burn to achieve LEO and uses minimal thrusters to align with the rendezvous in LEO.  After docking, the next burn is a long full throttle boost to TLI.  The rendezvous at LOP-G cannot be accurately modelled in KSP * however in theory it requires a burn of approx 190 m/s (full throttle, 19 secs) at lunar Pe, and 220m/s (full throttle,  22 secs) to rendezvous with LOP-G  Again the upper stage is undocked and assuming there is any remaining fuel, burns to suborbital , and the Cygnus provides the engines and RCS thrusters to complete the docking with LOP-G.

* I have LOP-G in an elliptical polar orbit with Pe= 2000km over the north pole, and Ap= 60,000km over the south pole.  I need over 600m/s dV to rendezvous, and I did not quite have enough fuel in this simulation to successfully rendezvous. 

Quote

What residuals does the FHUS have after the NRHO insertion burn? There will be boil-off, but if it has sufficient residuals at that point then we know boil-off isn't a problem.

At lunar Pe there is enough in the tank to do about dV = 450m/s, and that's it. So boil off may be a problem.

Quote

What is the end mass of the reusable crew module?  What are your thoughts, if any, on airlocks and egress? I've been trying to come up with a way to allow a docking ring that opens to a lower chamber for an airlock and then also serves as the ingress/egress port from LOP-G, since LOP-G is orientation-agnostic. Reasoning being: we already have structural androgynous pass-through docking ports, but we'd need to design propellant-transferring structural non-pass docking connectors. It would be easier to design connectors which only transfer propellant and are not structural. Then again you'd end up with more complicated sequences at LOP-G.

The capsule is 2.7 tonnes, including crew, food and water, and other human requirements.

I think the best place for an airlock is a solid structure inside the capsule.  It is simpler, so less risky than an inflatable external structure.  The internal airlock would be tiny and cramped, with just enough room for the two EVA suits and room to manouevre in and out, plus some means to clean off moon dust (a shower?) and a filter system in the floor.  This would leave very little living space in the craft although it could be stretched a bit at the top, potentially creating a separate living space, and still fit in the FH standard fairing.

mIzo6Ln.png

I have designed the whole system so there is no requirement for docking ports with a fuel transfer capability.  Fuel transfer is simply not required.  The docking port at the top of the lander capsule is structural and allows for direct pass-through to the LOP-G habitation module.

Quote

What downmass capability does the landing stage have in this configuration?

At this stage, I have not considered downmass capability.  I envisage that instruments would be needed to be deployed, and expect that these could be stowed by attaching to the outside of the thermal protection fairing at the base of the lander, so the crew access it while doing their EVAs, and not need to carry anything large through the airlock and down the ladder.  There isn't really a capability for many tonnes of equipment to be delivered, (e.g. a rover) although as I mentioned before, there is about 10% fuel remaining at each stage, so there is room for fine tuning and ability to add a bit of payload.

Quote

If the ascent stage (with or without drop tanks) is used as a one-way cargo lander direct from TLI, what kind of capacity does it have?

The craft landed on the surface as shown above has a wet mass of 9.2 tonnes.  If the craft were to burn directly from TLI to the surface, the payload would be higher.  I'll do a quick test tonight. 

Edited by jinnantonix

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4 hours ago, jinnantonix said:

The MVac burns are as follows: 

1.  FHe Launch 1:  The Lander Launch requires full throttle burn to get to LEO, it then need a three short burns at minimum (39%) throttle of about 30 m/s (6 secs) each to rendezvous in LEO.  The upper stage is undocked and burns to suborbital, and the Cygnus provides the engines and RCS thrusters to complete the rendezvous in LEO.

My primary concern with using the first launch's MVac to perform any sort of rendezvous would be loiter time. The stack placed in LEO needs to be able to hang out for a while in case of scrubs/scheduling conflicts. The really nice thing about the mission profile we've outlined so far is that if executed properly, we make maximal use of high-performance cryos on every possible burn, but never have any extended/unplanned loiter with cryos just sitting there boiling. It doesn't matter (within reason) how long the stack sits there in LEO waiting on TLI; it doesn't matter how long the stack is parked at LOP-G waiting on Orion. Accordingly, I think it is better to have the first launch relight as needed in the hours after launch to place the stack in the most ideal specified reference orbit, then decouple, and depend more on a very accurate ascent for the second stage.

4 hours ago, jinnantonix said:

2.  FHe Launch 2:  The naked FH upper stage has a full burn to achieve LEO and uses minimal thrusters to align with the rendezvous in LEO.  After docking, the next burn is a long full throttle boost to TLI.  The rendezvous at LOP-G cannot be accurately modelled in KSP * however in theory it requires a burn of approx 190 m/s (full throttle, 19 secs) at lunar Pe, and 220m/s (full throttle,  22 secs) to rendezvous with LOP-G  Again the upper stage is undocked and assuming there is any remaining fuel, burns to suborbital , and the Cygnus provides the engines and RCS thrusters to complete the docking with LOP-G.

Those burn times are plenty long and permit full engine transients and targeted downthrottle to ensure accuracy. My concern was that we'd end up with burn times under a few seconds and it wouldn't be long enough to make it accurate. I would note that Cynus, while nice, is not needed for final approach and docking to LOP-G; the lander stack itself should have ample RCS and propellant margin to do this on its own.

4 hours ago, jinnantonix said:

At lunar Pe there is enough in the tank to do about dV = 450m/s, and that's it. So boil off may be a problem.

So about 700 kg of props left? Let's see...that's about 56 m/s margin before the stack is dropped. Removing the capsule (or Cygnus, in the reflights) would free up a lot of margin. By my numbers, you're reserving 5.73 tonnes of propellant after the TLI burn. Assuming a six-day transit to LOP-G with powered lunar flyby, we'd expect less than 1.2% propellant boiloff (0.2% per day) which is 69 kg of propellant, so boiloff itself is not a problem here.

4 hours ago, jinnantonix said:

The capsule is 2.7 tonnes, including crew, food and water, and other human requirements.

If you cut the capsule from the nominal launch sequence, you can increase the mass of everything else by 7.9%, which means you can bump the capsule up to about 3 tonnes, which is nice

4 hours ago, jinnantonix said:

I think the best place for an airlock is a solid structure inside the capsule.  It is simpler, so less risky than an inflatable external structure.  The internal airlock would be tiny and cramped, with just enough room for the two EVA suits and room to manouevre in and out, plus some means to clean off moon dust (a shower?) and a filter system in the floor.  This would leave very little living space in the craft although it could be stretched a bit at the top, potentially creating a separate living space, and still fit in the FH standard fairing.

I have designed the whole system so there is no requirement for docking ports with a fuel transfer capability.  Fuel transfer is simply not required.  The docking port at the top of the lander capsule is structural and allows for direct pass-through to the LOP-G habitation module.

I would argue that we do need propellant transfer, in some capacity. Either we are transferring propellant into internal tanks on the capsule or we are using a new connector to feed the capsule's RCS from a drop tank. In the latter case (which is preferred for a variety of reasons) we are pressurizing the capsule's RCS using pressurized tanks and so we need a docking/mating connector which will permit that kind of high-flow-rate propellant feed.

We can have the structural docking port on top that allows crew to pass through in and out of LOP-G, and a structural propellant-feeding connector on the bottom which transfers force from the lower stages into the capsule while also feeding propellant into the capsule's RCS. That's what you have. However, this means we need a new double-duty connector element. It would be easier to develop a connector which only mates the tank itself to the capsule and provides propellant feed, without needing to bear structural loading from the capsule. Then, the "standard" docking port could be placed in the bottom of the capsule. This allows it to be used to mate to the ascent stage and the rest of the stack and also opens up the possibility of having a lower-stage airlock that is left on the surface (though dropping it while retaining the ascent engine(s) is a new challenge).

4 hours ago, jinnantonix said:

The craft landed on the surface as shown above has a wet mass of 9.2 tonnes.  If the craft were to burn directly from TLI to the surface, the payload would be higher.  I'll do a quick test tonight. 

I'm more thinking of one-way unmanned downmass for logistics and surface asset delivery. For example: remove the drop tanks and replace the crew capsule with a surface asset, then launch to TLI on Falcon 9R or Atlas V, using the ascent module's propellant to perform the LLO insertion and descent. What kind of downmass do you have then? What if you add another pair of drop tanks? A modular delivery system is critical.

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5 hours ago, sevenperforce said:

However, this means we need a new double-duty connector element. It would be easier to develop a connector which only mates the tank itself to the capsule and provides propellant feed, without needing to bear structural loading from the capsule. Then, the "standard" docking port could be placed in the bottom of the capsule. This allows it to be used to mate to the ascent stage and the rest of the stack

Ok, I see your point now, of course.  There needs to be both a structural and a feed capability.  KSP and SSTU Labs provides this is a single part, but IRL this is not so simple.  It could perhaps be modeled using two ports, one inline and the other (maybe a mini port) offset.  

Quote

and also opens up the possibility of having a lower-stage airlock that is left on the surface (though dropping it while retaining the ascent engine(s) is a new challenge).

I have tried numerous ways to design a lower stage airlock, and it beats me.  I just can't fit such a craft into the FH standard fairing, or it becomes stupidly tall and won't land on a slope.  I will double my efforts, because it is just such a neat idea.

Quote

I'm more thinking of one-way unmanned downmass for logistics and surface asset delivery. For example: remove the drop tanks and replace the crew capsule with a surface asset, then launch to TLI on Falcon 9R or Atlas V, using the ascent module's propellant to perform the LLO insertion and descent. What kind of downmass do you have then? What if you add another pair of drop tanks? A modular delivery system is critical.

OK, I get it.  I don't think I am ready to do that modelling yet.  I am working on the fine tuning now, adding a downmass capability, and this is changing the tanks sizing a bit.  Would love to add an airlock in the second stage.  Need to consider life support too.  This is going to take some time...

 

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