Completely new CLLARE design. | Mission Planning Workgroup | Forum

Post edited 6:28 am – March 5, 2010 by Luke Maurits

Yes, really!  I'm not trying to be a pain, but I have been thinking about alternative approaches to CLLARE overall later, and especially trying to be as ruthlessly simple and minimalist as we initially intended to be.  I offer the plan below not just to be judged on its own merits, but also to highlight rather a larger problem.  All of the fantastic work we have done so far on the Engineering Process etc. has been focused on fleshing out / refining the current overall plan, but that plan has not exactly been chosen via a rigorous, documented process, it's basically the best thing we have come up with so far.  I am not sure what we want to do for CLLARE right now, but looking forward to new projects, we will need to come up with a procedure for choosing the best possible overall structure before we use the Design Tasks Tree to flesh out details.

Anyway, onto the plan.

The first big change in this plan is that the CM is spherical, ala Vostok / Voskhod.  There are a number of reasons to prefer this:

• It offers the most internal volume ("living space") for any given CM surface area (hence mass), which will probably allow a lighter capsule compared to a cone.
• A sphere is also the optimal shape for a pressure vessel, which means the walls do not need to be quite as strong.
• The aerodynamics of a sphere are trivial: the cross-sectional surface area and drag coefficient are the same regardless of orientation, and no lift is generated.  I would have no idea where to start simulating the atmospheric reentry of an Apollo-shaped capsule, but I am confident I could do a sphere, no problems!
• I am willing to bet a lot of other analyses (to do with stress forces, heat distribution, etc.) are a lot easier for spheres as well.  This is more significant than it sounds: it might be the difference between an analysis requiring the use of multi-hundred (or thousand!) dollar proprietary software using all sorts of fancy finite element algorithms, etc. and being able to literally do the analysis on paper (or if not on paper using much simpler numerical methods we can comfortably code ourselves for free). 
• RCS is simpler with a symmetric shape.  In the current conical CLLARE CM design, the RCS thrusters in the nose are a little inconvenient in that it is impossible to translate the CM through space without also inducing rotation.  This was the motivation for adding RCS thrusters on the OSM – with this arrangement, the front and rear thrusters can both fire, eliminating any torque but still providing force.  With a sphere, an arrangement of 4 equally spaced thrusters around a "great circle" on the CM allows rotation-free translation without any additional thrusters on, e.g., an OSM.

About the only downside to a spherical capsule there is is that, due to the lack of lift, the g-forces experienced during reentry are quite high (around 9g).  On this front, I have two thoughts: first of all, 9 g is uncomfortable, but survivable: the Soviets did it with Vostok and Voshkod quite a few times.  This mission is not exactly going to be a picnic as it is, maybe we should be hard asses about it and tell the astronaut to deal with a short exposure to high gs as part of the deal: simplicity over comfort.  Secondly, if we decide we don't like this, then one alternative is a great big ballute.  Check out the "attached ballute" on page 2 of this presentation.  With one of these attached to our spherical CM, during reentry the spacecraft basically becomes a huge shuttlecock, like in badminton.  This is a very high drag shape (so lower gs, lower thermal stress, etc.) and is also incredibly aerodynamically stable – it's certainly not going to flip over during reentry!  This point could save on the use of RCS during reentry.

The second big difference is that this plan largely does away with the separate lander!  Instead, the spherical CM is used for the trip down to the moon, by attaching a set of legs and an engine to it (this entire attachment is separate so that for non-landing missions it can be removed).  Now, this may seem a little crazy in that it involves carrying significantly more weight down to the surface and back up than our bare-bones open cab lander, but I am starting to think (admittedly I should crunch some numbers to back this up) that this is actually worth it, because this approach buys us so very much in terms of simplicity:

• It drastically reduces hardware duplication.  With a separate CM and LL like we have now, the LL needs to have its own computers, communications hardware, inertial navigation hardware, RCS system, batteries, etc., etc.  Huge chunks of functionality are duplicated, whereas in this new approach there is only one of all these things.  This is much simpler and more reliable overall.  It also will reduce overall stack mass a little (probably not enough to make up for the extra mass of the spherical CM over an open lander, though).  Also worth noting: we are not just saving needing to have a second computer, comms system, etc., but a second vacuum proof set of those things for the lander (either that or adding a pressure vessel on the lander, and a system to heat it, etc, etc).  These really are drastic increases in simplicity.
• Comms with Earth will be direct from the surface of the moon to Earth, rather than being relayed through an orbiting CM.  This means no radio blackouts during lunar EVA.  It might also mean the ability to send more data back rather than recording it, since our LL would probably have quite low-bandwidth comms gear.
• It removes the need for EVA transfer from one vehicle to another while in orbit.  While this is possible (e.g. it was part of the Soviet lunar landing plan), it's always seemed like one of the riskiest parts of the existing mission plan in my mind, and I'm not sad to see it go.
• It means the astronaut only has to rely on their suit for life support during lunar EVA, not during ascent and descent, which means we can make our lunar EVA a little longer.

Perhaps the one drawback to this approach (aside from the increased mass of stuff we have to lower into and lift out of the moon's gravity) is that it requires a genuine, full-blown redocking of the spherical CM with the orbiting PM, whereas the old plan required the LL simply to get close enough to the orbiting CM for an EVA jump across to be feasible.   Note, however, that docking a sphere to a cylinder is a lot easier than docking an open cab lander to a cone.  I think this could be relatively straightforward, and would be a good capacity for the CM to have from the point of view of it not being quite so single-purpose.

These two changes represent the core of my new plan, but another thought I have which we might also want to consider if we do opt to go with this approach is using an inflatable airlock, like the Volga airlock which was used on Voskhod 2, for transfering from the CM to the lunar surface after landing.  This has two main advantages: one, our avionics do not need to be able to survive cabin depressurisation like in the current plan, which means they do not need to be in a separate pressure vessel, saving on mass (this was the Soviets' motivation.  From the Wikipedia page above: "The airlock was necessary because Vostok and Voskhod avionics were cooled with cabin air and would overheat if the capsule was depressurized for the EVA").  Secondly, it means if something goes wrong with the suit life support on the lunar surface, there's a pressurised capsule on the surface the astronaut can run back to.  The biggest challenge here would be integrating the inflatable airlock with some sort of ladder.

I know these are very drastic changes which are coming kind of late: if we switched to this approach a lot of the work we have done so far (although by no means all of it, possibly not even most of it) would become irrelevant.  It's not a decision to be taken likely, but I think it is worth considering because all of the benefits listed above are not insignificant.  I feel like this approach genuinely is a lot simpler.  Ultimately it will depend on how light we can get our spherical CM and whether or not we can keep the total mass under the Falcon 9's limit.  Before I actually do the work involved in figuring that out, though, I very much welcome feedback on the overall approach.  Can anybody see any shortcomings with this approach, potential problems I've overlooked, etc?

Rocket-To-The-Moon said:

• The lander will be much more massive with means that its fuel requirement will be higher
• The PM's fuel requirements may be reduced if the new lander's mass is less than the mass of the old CM and lander
• Such designs have been abandoned in the past because they are fuel inefficient
• We will need a way to not only physically redock the CM to the PM, but also a way to disconnect/reconnect the life support supplies (difficult).

I agree with all of these points, by the way.  Additionally, the PM's fuel requirements will also be reduced by the removal of the need for an RCS on the OSM.

There is definitely a lot of interplay between all these changes, with increasing and decreasing masses, and it will take a little bit of work to figure out if the overall mass can be kept the same or lowered.  I am pretty sure that the most important of these points will be the first one: the lander's increased mass, and its fuel ramifications.  Given the considerable appeal of the change, we may want to direct a lot of effort into coming up with innovative/crazy modifications to the landing process to try to minimise the increased mass.  I wonder if reducing the lunar orbit's altitude will help with this?  Some sort of crazy balloon-cushioned landing with a higher vertical velocity than would be allowable for a legged lander?

Rocket-To-The-Moon said:

 I think that the original direct ascent mission profile had the entire lander fly back to Earth (the lander provided the delta-v for TEI).

I believe this is correct.  In the direct ascent profile there was only one spacecraft involved at all.  Certainly part of the reason this was hugely fuel inefficient was having to carry a large and heavy spacecraft down to the surface and back again, but I think the larger part, or at least a very significant part,would have been having to carry the TLI fuel down then up again as well before using it, which the plan here doesn't involve, so we might be able to get away with it.

I think the deciding factor will be how light we can get our spherical CM.  Lighter than a cone seems given, but I'm not sure how much lighter.

An alternative idea on how to make this feasible:  The biggest problem there is in this new plan is carrying the mass of the spherical CM down to the surface.  We need as efficient a landing profile as possible.  While doing research on possible solutions to these, I discovered that the Soviet moon landing plan actually had quite a different landing profile to Apollo, even though the landers look superficially similar:

Basically, the lander is mated to a "Block D" rocket engine which serves to do most of the work to bring it and the lander out of the initial parking orbit, down to about 4 km above the lunar surface.  At this point the Block D burns out and is jetisonned, crashing into the moon some distance away.  The lander takes over with its own engines to do the last little bit of the landing, and to do ascent.  I suppose this is somewhat like having a separate ascent stage (ala Apollo) that you get rid of a little earlier.  Anyway, the Soviet lander using this approach had about 40% the mass of the Apollo lunar lander.  This must have been due in part to it holding 1 cosmonaut instead of 2 astronauts, but I don't think that alone can explain all of the reduction.  Perhaps we should investigate an approach like this for New CLLARE.  To assist with simplicity, since we will need our own "Block D", perhaps the PM could be replaced by a chain of "Block D"s, each jetisonned after it was empty?  This might help lower fuel costs overall, too, since we are not carrying as much dead weight in the form of empty tank?

I will try to do the numbers on this approach soon.

I have done some extremely preliminary number crunching for this plan utilising the Soviet landing profile and the results are promising, although I'm trying not to get too excited until I refine them somewhat.  At this stage I think the Falcon 9 will be able to support this proposal with an unmanned spherical CM mass of 800 kg, not 500 kg as quoted earlier, which is rather more realistic.  I also suspect that refinements of my calculations will lead to this figure increasing, rather than decreasing.

Rocket-To-The-Moon said:

That sounds promising. 800kg is just the CM portion right?

Yep, just the spherical CM and it's non-human contents.  No legs, no landing engine, or anything else.

I have now done some much more refined work and offer the following arrangement, with a total launch mass of 10448 kg, which can just squeeze onto a Falcon 9.  As it turned out, the refinements ended up making no difference to the CM mass.

CM/lander arrangements:

• CM mass: 800 kg
• Landing leg attachment for CM: 100 kg
• Landing engine attachment for CM: 150 kg
• Landing engine propellant tanks: 40 g per 1 kg of propellant (taken from Space Shuttle external tanks)
• Total lander propellant supply: 970 kg
• Mass of fully fuelled lander: 2237.15 kg

PM arrangements (this is not the old PM, this is a new "Block D" PM):

• PM engine and structure: 150 kg
• PM propellant tanks: 40 g per 1 kg of propellant (as above)
• Total PM propellant supply: 1430 kg
• Mass of fully fuelled PM: 1642.14 kg

With these figures, one fully fuelled PM is able to impart a 2200 m/s delta-v to the lander configuration (CM, lander, lander engine, fuel tanks and fuel), as per the Soviet landing profile.  The LL fuel is able to impart a 300 m/s delta-v to the lander config for descent (again as per Soviet profile) and 2220 m/s delta-v for ascent (as per Apollo) if legs are jetisonned after take off.  A single fully fuelled PM is also able to apply the LOI delta-v of 1000 m/s to the lander configuration and the PM it uses to descend, with fuel to spare, and then if the lander redocks with that half-full PM after ascent the remaining fuel is sufficient to apply the 700 m/s TEI delta-v to the CM (if lander engine and tanks are jetisonned).  A stack of 3 fully fuelled PMs, each jetisonned after burn out, is enough to do the 3200 m/s TLI delta-v on the lander configuration plus the two PMs needed for LOI, descent and TEI.  So our total stack is 5 PMs, the CM and the various landing attachments to the CM.

All of the above is based on a 75 kg astronaut and a 100 kg EVA suit (noteworthy is that if we use a mechanical counterpressure suit like that Webb / OLF are working on, that EVA suit mass should drop considerably, increasing allowable CM size), and allowing an additional 10% extra delta-v on all maneuvers other than TLI (if we dropped our delta-v safety margins somewhat, allowable CM size would again increase).

I have tried to overestimate rather than underestimate some things – e.g. 100 kg for the landing legs is probably a little much, perhaps also 150 kg from engines and structures.  If we drop these figures to 50 kg for legs and 100 kg for engines/structures we can get our CM up to 875 kg.  I do feel like the tank masses are absurdly small, even though they are derived from shuttle tanks.  Perhaps this is because the shuttle tanks are so damn huge that they get really good volume-mass efficiency.  Ours will probably be considerably heavier.

As a consequence of these super light tanks, the stacking of 5 PMs to handle propulsion is probably actually a bad thing: the mass saved by jetisonning empty super light tanks is more than countered by the extra 150 kg per PM of engine and structure.  It might be the case that with more realistic masses for these things the PM stack approach is best, but it would also be worth while considering a single large PM, as per the current CLLARE plan.

Anyway, I am starting to feel fairly sure that we can get this plan to work from a mass standpoint.  From a simplicity standpoint, I think there is not even room for argument, this plan is simpler than the other plan, in many, many ways.  The one other thing I want to convince myself of before I start really arguing with conviction that this plan is superior, is that we can design the spherical CM to endure both the deep space environment and the lunar surface environment, which are somewhat different, mainly from a thermodynamic point of view.

Another thing that needs some consideration is how all of this hardware sticks together.  The spherical CM is quite flexible in that you can mount things to it at just about any position at just about any angle, but finding a way to stick everything together so that the pilot is seated in an appropriate orientation to both the landing legs and the liftoff thrust, and so that the lunar descent PM is angled through the detached lander's centre of mass, and so that the TLI/LOI/TEI PMs are angled through the whole stack's centre of mass, is not necessarily trivial.  I don't think we can meet all of these criteria at once so it will be a matter of choosing a trade off.  Anyway, more on this later.

Some more sources for spherical capsule mass estimates:  In the 1950s, before NASA was created and so before project Mercury, there was a USAF project called "Man In Space Soonest" (MISS) with the goal of putting a man in space before the Soviets.  The USAF contracted a number of civillian companies to produce proposals for spacecraft to do this.  You can read about all of them here, but three of them are of particular interest:

1. A proposal by AVCO for a 2.1 m diameter spherical capsule with a mass of 690 kg, which would support a crew of one in orbit for 7 days.  Interestingly this plan had a very large high-drag device on top kind of like the ballute I have proposed for spherical CM reentry in this plan.
2. A proposal by Convair for a 1.6 m diameter sphereical capsule with a mass of 450 kg, of unknown orbital endurance.
3. A proposal by Goodyear for a 2.1 m diameter spherical capsule with a mass of 900 kg, which would support a crew of one in orbit for 5 days.

I don't know why Goodyear's spacecraft was so heavy, but the AVCO and Convair proposals are both very light, much lighter than the 800 kg maximum which a Falcon 9 can support, which is excellent news.  These spacecraft had sufficient heat shielding to survive reentry from orbit: reentry on a return-from-moon trajectory will require additional shielding, but considering (i) that these masses have at least 100 kg spare before hitting our limit, and in the Corvair case as much as 350 kg to spare(!), and (ii) these masses are from the 1950s (no composite materials, big heavy analogue electronics, heavy batteries, etc.), it seems entirely feasible to me that we could get a modern spherical capsule with sufficient heat shielding for a lunar return to be under 800 kg.

Post edited 7:34 pm – March 8, 2010 by Luke Maurits

Very detailed technical article on using ballutes for reentry on lunar-return trajectories.  I've only quickly skimmed it so far, but this will be a very useful thing to read later on to fully assess the feasibility of this idea.

EDIT: Another similar article, although less detailed.

Post edited 4:45 am – March 11, 2010 by Luke Maurits

More updates:

1. I found an oversight in the fuel analysis code: it wasn't counting the mass of the fuel in the lunar lander tanks for the purposes of the TLI and LOI burns.  Including this reduced the maximum CM mass back down to around 800 kg. :(
2. I thought about the 10% extra delta-v allowance and thought it didn't make much sense.  In some situations, like lunar descent and ascent, where nominal requirements are modest (compared to TLI), there is considerable uncertainty and running out of fuel would be a huge problem, it makes sense.  In others, though, like TLI, it doesn't.  An extra 10% ontop of TLI is 320 m/s, which is not a small delta v, and there is no real uncertainty in the requirements for TLI.  That safety margin is excessive.  So I have used 10% safety on lunar ascent and descent and 5% on other maneuvers and this carried the maximum CM mass back up to around 900 kg. :)
3. I actually crunched the numbers on the "many small PMs for TLI/LOI/TEI vs one large PM for same" issue and it turns out the many small PMs approach really is better, it saves an incredible 2000 kg overall!  A note for the "Old CLLARE" concept: instead of having a Light Propulsion Module and Heavy Propulsion Module, perhaps just having the light one and clustering it on lunar landing flights would save mass?
4. Worried that a tank mass fraction of 0.9 may be optimistic, I did some Googling.  I didn't find much, but did find this page which says "Thin gage corrosion resistant steel (CRES) construction combined with normal fusion welding as used on LM's Centaur has already been demonstrated to provide the highest cryogenic tank mass fraction (~.90) for large scale, cryogenic propellant storage".  So 0.9 is, actually, achievable.  The article goes on to say that new welding techniques may be able to increase this.
5. The point above is good news because I made the following table of tank mass fraction vs maximum allowable CM mass, and was startled by how strong the dependence is (sorry for bad formatting):

   Tank mass fraction: 0.990000    Maximum CM mass: 976
   Tank mass fraction: 0.980000    Maximum CM mass: 968
   Tank mass fraction: 0.970000    Maximum CM mass: 961
   Tank mass fraction: 0.960000    Maximum CM mass: 953
   Tank mass fraction: 0.950000    Maximum CM mass: 945
   Tank mass fraction: 0.940000    Maximum CM mass: 937
   Tank mass fraction: 0.930000    Maximum CM mass: 929
   Tank mass fraction: 0.920000    Maximum CM mass: 920
   Tank mass fraction: 0.910000    Maximum CM mass: 912
   Tank mass fraction: 0.900000    Maximum CM mass: 893
   Tank mass fraction: 0.890000    Maximum CM mass: 859
   Tank mass fraction: 0.880000    Maximum CM mass: 830
   Tank mass fraction: 0.870000    Maximum CM mass: 803
   Tank mass fraction: 0.860000    Maximum CM mass: 736
   Tank mass fraction: 0.850000    Maximum CM mass: 727
   Tank mass fraction: 0.840000    Maximum CM mass: 718
   Tank mass fraction: 0.830000    Maximum CM mass: 709
   Tank mass fraction: 0.820000    Maximum CM mass: 680
   Tank mass fraction: 0.810000    Maximum CM mass: 656
   Tank mass fraction: 0.800000    Maximum CM mass: 633
   Tank mass fraction: 0.790000    Maximum CM mass: 598
   Tank mass fraction: 0.780000    Maximum CM mass: 540
   Tank mass fraction: 0.770000    Maximum CM mass: 531
   Tank mass fraction: 0.760000    Maximum CM mass: 521
   Tank mass fraction: 0.750000    Maximum CM mass: 498
   Tank mass fraction: 0.740000    Maximum CM mass: 462
   Tank mass fraction: 0.730000    Maximum CM mass: 426
   Tank mass fraction: 0.720000    Maximum CM mass: 403
   Tank mass fraction: 0.710000    Maximum CM mass: 360
   Tank mass fraction: 0.700000    Maximum CM mass: 350

   As you can see, even dropping from 0.9 to 0.8 makes the mass constraints very uncomfortable.  By 0.7 the mission is basically impossible.  Literally every single percentage point of tank mass fraction matters significantly, so it would make sense for us as a team to be willing to invest a lot of time and money in finding innovations in this field.

Also, at last, the first concept images for this new plan.  These are very rough and preliminary, sorry:

The top left image shows the CM by itself.  Note that the human figure inside is not drawn to any kind of scale, it is merely to indicate the orientation of the seating.  The circles to the front of the CM are portholes.  There's actually no reason behind me using lots of small portholes instead of one big one or some other system in these images, I was just thinking of SpaceShipOne and thought it looked kind of cool.  We can make an actual informed choice for the real CM, of course.  The black thing with four triangles coming out of it in the middle is an RCS thruster, on the outside of the CM – there's another one on the top and bottom, and another one on the far side of the CM – they are spaced out equally along a vertical circumference line of the CM.  With this arrangement, independent translation and rotation of the CM is possible.  It's not shown, but when the CM is launched standalone on an orbital flight, a deorbit rocket pack can simply be strapped on ala Mercury.

The top right image just shows the CM with two PMs attached, in a configuration useful for circumlunar flights.  There are 2 PMs just for the sake of illustration, I've not done the numbers on how many would be needed for an actual free return TLI (although I think 2 would be right).  The appearance of the PMs is purely for illustrative purposes.  In the actual design I imagine they would have a higher radius and lower length, to make sure the lunar landing configuration can fit inside the Falcon 9 fairing.

The bottom image shows a lunar landing configuration: 6 PMs behind the CM (this is the actual number – I just noticed the diagram says 5 not six, that's a typo) and the lunar landing gear underneath the CM.  It's something of a shame that the landing legs make the whole arrangement asymmetric, since it will complicate propulsion somewhat, but I don't know what else to do.  If we point the legs forward to maintain symmetry about the stack of PMs, then the astronaut is facing downward at the time of landing.  When they remove their seatbelt they will fall straight into the control panel.  If we reorient the seat so that it would be in the correct position for landing with the legs pointing parallel to the PM stack, then the acceleration during Falcon 9 take off is parallel to the spine, and the body is less tolerant of g forces in this position.  We could have the entire insides of the CM mounted so that they can rotate inside the shell to change arrangement, solving the problem perfectly, but this will add mass and complexity.

One thing I've not yet figured out is where to put the CM door.  If it is on one of the sides of the PM, then a RCS thruster will be mounted to the outside of it.  This means the hoses/valves for the thruster will need to be integrated into the structure of the door.  Maybe this is not a big deal, but it sounds a little unpleasant.  I thought of moving the RCS thrusters so that, if one is facing the CM head-on, they are at NE, SE, SW and NW, instead of at N, E, S and W, but then translation is only problem along an axes which is diagonal from the perspective of the pilot, and that could be confusing.  The door could go on the front, but we would need to design the control panel such that the astronaut could exit over it easily.  The back is no good because it will not be accessible while the CM is on the launch pad.

Feedback welcome, as always.