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5:21 am

March 5, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 1 0 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? Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

5:46 am

March 5, 2010

DenisG Saarbrücken, Germany (GMT+1) Member posts 69 2 0 I think it makes much sense to have many different concepts (and sub-concepts) beforehand; they should all be evaluated and then the most promising one should be chosen. Starting work on the first thing that comes to mind is a bad idea. From that POV I very much welcome your new concept. The work on the current concept is not useless though. Every concept must be researched to have a base on which to evaluate. Also, I think the round lander concept sounds good. I propose preparing a document with all of the info you find on round lander concepts that you have. Same thing for the current concept. That will allow everyone to understand the differences and pros/cons quickly for judgement.

8:00 am

March 5, 2010

Rocket-To-The-Moon Grand Forks, North Dakota, USA Member posts 641 3 0 This is definitely a major deviation from our current plans and we will need to have an in depth discussion of the merits of it before we commit to changing everything. Here are a couple of brief observations. 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 think that the spherical CM is a likely candidate, but the one piece lander will require a lot of analysis. Main Workgroups: Propulsion & Spacecraft Engineering

8:05 am

March 5, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 4 0 Rocket-To-The-Moon said: Such designs have been abandoned in the past because they are fuel inefficient Sorry, could you point out exactly what you were referencing in that article? I gave that section a quick skim and didn't see anything that immediately jumped out at me as being similar to my design. Are you thinking of the direct ascent plans from very early in Apollo, or did I miss something? Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

8:36 am

March 5, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 5 0 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? Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

8:46 am

March 5, 2010

Rocket-To-The-Moon Grand Forks, North Dakota, USA Member posts 641 6 0 Iwas referring to direct ascent being discarded in favor of LOR due to mass issues with a heavy lander. I would have to research a little bit further, but 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). If this is the case, then us doing LOR with the lander and the PM might solve the mass issues that direct ascent had. It would sort of be a hybrid between direct ascent and the final Apollo. Once again, sorry for the lousy resource. I was rushed to find something to support my point and came up short. Main Workgroups: Propulsion & Spacecraft Engineering

10:07 am

March 5, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 7 0 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. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

10:13 pm

March 5, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 8 1 Luke Maurits said: 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. A most unusual source of data, however… Free Spirit Spheres are large, 3.2m diameter fibreglass spheres that people of a certain lifestyle choice like to hang from trees and live in. Each to their own. According to the site's FAQ, the spheres have a mass of 500 kg. Now, 3.2 m is way too large for our purposes. I think at the absolute most we would use 2 m, and more probably something like 1.5 m, but let's stick to 2 m for now. A 3.2 m diameter (1.6 m radius) sphere has a surface area of 32.17 m^2, whereas a 2m diameter (1 m radius) sphere has a surface area of 12.57 m^2, or about 39% of the surface area. 39% of 500 kg is just 195 kg, which is incredibly light. Now, the Free Spirit Spheres are 3mm thick, and we may want more than that, but we could also use lighter carbon fibre instead of fibre glass to counteract that. Maybe around 200-250 kg for the structure (i.e. not including heat shielding) is reasonable? Compare this to 340 kg for Mercury and 638 kg for Gemini (average 489 kg) and we are looking pretty light (admittedly a huge part of this is the use of composite material instead of metals, which we could have done with the conical CM as well – this just means that our current estimates for the conical CM are probably quite pessimistic). Of course, if we go with a ballistic spherical reentry, the heat shielding will need to be quite heavy, compared to a lifting conical reentry. This is another argument in favour of the ballute (slightly eliminating the benefit of easy spherical aerodynamics). Playing around with a roughly modified version of the CLLARE fuel analysis code, if we can get our unmanned spherical CM down to 500 kg, eliminate the OSM entirely (by moving all life support supplies into the CM – I think this was how Vostok worked, and it apparently had an endurance of 10 days!), and get the leg/engine/tank attachments onto the sphere down to 200 kg, then we can do this (10314 kg fully fuelled with 450 s Isp). Those are quite bold figures, but perhaps not impossible. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

7:13 pm

March 6, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 9 1 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. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

7:47 pm

March 6, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 10 1 By the way, in case it isn't clear, some of the ideas in this thread are not necessarily linked to the prospect of a whole new CLLARE approach, and could just as well be brought into the "old" CLLARE plan if we thought it better: things like replacing a conical CM with a sphere, using an inflatable airlock to prevent CM depressursation, and using a "Block D" to assist in lunar descent could all, with varying degrees of difficulty, be applied to the familiar plan. We should expect and welcome cross-fertilisation of ideas between "competing" plans for CLLARE. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

5:43 am

March 7, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 11 0 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. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

7:15 am

March 7, 2010

Rocket-To-The-Moon Grand Forks, North Dakota, USA Member posts 641 12 0 That sounds promising. 800kg is just the CM portion right? Hopefully we can get it less than that, but it is good to hear that the Falcon 9 can support something that large. Do you have any estimates on the lander's fueled mass? I have to admit that this design is starting to grow on me from a simplicity standpoint. Main Workgroups: Propulsion & Spacecraft Engineering

10:25 pm

March 7, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 13 0 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. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

7:10 am

March 8, 2010

Rocket-To-The-Moon Grand Forks, North Dakota, USA Member posts 641 14 0 So with this plan we have a total of 5 PMs and each PM has its own engine and fuel supply. I can see that this is definitely a good way to make it modular. For unmanned non-return missions the number of PMs could be drastically reduced. As the PMs are jettisoned it is going change the center of mass quite significantly which will make the RCS a little bit more difficult to control. Since we are using a heavier lander will we still be able to use the single design engine that the PM uses? Main Workgroups: Propulsion & Spacecraft Engineering

7:23 am

March 8, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 15 0 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: 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. A proposal by Convair for a 1.6 m diameter sphereical capsule with a mass of 450 kg, of unknown orbital endurance. 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. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

7:24 am

March 8, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 16 0 Rocket-To-The-Moon said: Since we are using a heavier lander will we still be able to use the single design engine that the PM uses? Do you mean will we still be able to use the same engine on the lander and on PM clusters? I do not see why not. That engine will need to be able to produce more thrust to hover a heavier lander, but that just means that the PM engines will not need to burn for as long, which is actually good – the faster the delta v is applied, the more efficient. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

7:13 pm

March 8, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 17 1 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. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

4:44 am

March 9, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 18 0 A number of updates on the mass analysis for this plan. In my original mass calculations for this plan, I used a 300 m/s descent delta v. The reason for this was simply because I had read this little snippet from Encyclopedia Astronautica: But as soon as detailed design of the LK began it was realized that the mass of the spacecraft in the draft project was completely unrealistic. The young engineers that had done the preliminary LK design had made numerous absurd assumptions. They had assumed a soft landing delta v of only 30 to 40 m/s (200 to 300 m/s was a more realistic estimate) I simply took the highest value, 300 m/s. Having read more carefully, I can see that the Soviets did 3 separate tests of the LK lander's propulsion system in LEO. On each test flight they did a test burn to simulate ascent and descent on the lunar surface. The different ascent delta-vs in the simualtions were: 263, 251 and 266 m/s. So 300 m/s was a little high. I have now set the "nominal" descent delta v to 275 m/s, and included a 10% safety margin (just like I did before with the 300 m/s nominal descent delta v). Now, I also used an ascent delta v of 2220 m/s, which I copied straight from the Apollo Lunar Module. However! Having read more, I see that the Soviets were very careful in choosing the lunar orbit the LK had to descent/ascend from. From the same page: The Chief Designers offered prizes of 50 to 60 rubles per kilogram of weight reduction identified by project engineers. 500 kg was saved just by optimizing the rendezvous orbit. It turns out the Soviet's LOI put their stack into a 175 km high circular lunar orbit (higher than Apollo's 100 km orbit) and then used the block D to put the stack into an elliptical orbit with 175 km aposelene and 40 km periselene. This low periselene means the LK did not have to descend or ascend so far. The cost for this was an extra "eliptification" burn, but (i) saving on ascent/descent delta v is well worth small increases in orbital delta v, and (ii) the 175 km high circular orbit requires a lower initial LOI in the first place, and a lower TEI later, somewhat (perhaps completely?) off-setting the cost of the eliptification. Because of this unusual orbit, the LK did not have to ascend very far. The three test flights mentioned before had these delta vs when simulating ascent: 1518, 1320 and 1333 m/s (I am not sure why the first delta-v is so much higher than the others). So I have changed the nominal ascent delta v in my calculations from 2220 to 1600 (just to be safe) and once again included the 10% saftey margin. This is a big and exciting change! Saving fuel on ascent is fantastic, because it has dramatic flow on effects. The less fuel you need for ascent, the lighter the load you carry for descent, and so the less fuel you need for that – and since fuel requirements increase exponentially as mass does, seemingly small savings in ascent make a big difference in the end – and 622 m/s is a large saving! So replacing these numbers brings our maximum allowable CM mass up a lot higher, up to around 1050 kg!!! However, as mentioned, the propellant tank mass estimates I used used (of 4 g of tank per 1 kg of fuel) are absurdly optimistic (I think perhaps the mass numbers on the Wikipedia page about the shuttle tanks I used to arrive at that figure are wrong). To correct for this, I have gone with a "tank mass fraction" approach, and assumed that the mass of a full tank is 90% propellant, 10% tank structure. This increases the tank mass considerably and subtracts somewhat from the big mass bonus we got by improving our delta v figures, but we still come out on top (remember that as tank mass increases, the benefit from staging our big burns using many small PMs increases) with 950 kg of allowable CM mass. As an aside, the same unrealistic tank mass figures were used to estimate the mass of the "Old CLLARE" PM, so those figures are unreliable too. Importantly, this means that the Old CLLARE design which seemed just to fit on a Falcon 9 now probably will not. Of course, that plan can be salvaged using many of the techniques outlined for this plan, e.g. using an elliptical lunar rendezvous orbit, using a separate engine to handle most of the descent. Back to this plan: the 950 kg calculated above does not take into account the increased PM fuel required for the "eliptification burn". This is because (i) I am very busy and also very tired tonight and I don't want to take the time to calculate the exact delta v for that maneuver (but I can do that, and will) (ii) my gut tells me this will make very little change (the decreased LOI and TEI delta vs due to higher circular altitude will, I think, largely cancel out the eliptification delta v) (iii) the stack of 2 PMs used for LOI and TEI in the above plan already had a little bit of extra fuel in them that we can apply to this task. Even if these assumptions are wrong, I don't think there is anyway it could drop the CM mass back down past the old 800 kg figure. It might take it down to 900 kg or perhaps 850 kg at worst. Even if it goes back to 800 kg where it was, at least we can be a lot more confident in that 800 kg figure. Looking at the mass figures from the spherical Man In Space Soonest capsules I posted this morning, even 800 kg should certainly be okay for our CM, and 900 or 950 kg would be luxurious! Note that everything above is using the conservative estimates for little extra bits: 100 kg just for the lander legs, 150 kg for the lander engine and associated equipment, and 150 kg for the PM engine and associated equipment. If we can reduce these figures then, of course, maximum allowed CM mass goes up. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.

4:19 am

March 11, 2010

Luke Maurits Adelaide, Australia Admin posts 1208 19 0 Post edited 4:45 am – March 11, 2010 by Luke Maurits

More updates: 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. :( 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. :) 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? 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. 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. Main CLLARE workgroups: Mission Planning, Navigation and Guidance. I do maths, physics, C, Python and Java.