Yet another Mars architecture

Russel · 2013-06-18 11:39:31

Yes, I'm still alive. And here to run some thoughts past you guys.

I've been thinking about the problem of getting to Mars and back. (I'm less concerned with what we do there, or what stuff we need to land there permanently)

Key to that is the landing/ascent problem and I think critical is the ascent problem. Hence my interest some months ago in reusable landing/ascent vehicles.

Then I've considered how the overall architecture works. How to achieve simplicity, synergy and robustness. Which means integrating hardware in different ways.

First, I'm going to suggest something radical. And I'll come at it this way. Earlier I went into some detail as to how a fully (many years) reusable landing/ascent vehicle might look. It turned out that such a beast is technically possible, but with CO/O2 fuel is marginal - so marginal and also heavy as to be excluded. The next phase was again an attempt at a fully reusable landing/ascent vehicle using methane as the ascent fuel. This looked more promising. I found it also made sense that if you were to import (land) hydrogen then the descent phase was more economical run on hydrogen. But even so, given the low density of liquid hydrogen the tankage had to be scaled around the task of landing that volume of hydrogen.

As a side note here. For all the talk about mass leverage in importing hydrogen and then converting it to methane, for me the mass leverage is at best 4:1 (methane is 25% by mass hydrogen). And the reason I say this is that the process of making oxygen on Mars can happen without any imported precursor. Oxygen in a sense, is essentially free. (Yes, it does take energy). For me at least, it makes most sense to land methane on Mars, rather than hydrogen. At least until there is a good indigenous source of hydrogen.

What this points to is an ascent powered by methane/oxygen where the methane is imported (landed) and the oxygen is produced locally. I felt I had to spell that out in case it wasn't obvious what I"m going to say next.

The other problem that keeps arising with an ascent vehicle is sheer mass. Neglecting the crew, the mass at takeoff basically scales around the mass of the structure itself. Now, if that structure is also a lander then the mass increases for two reasons. One is heat shield and the other is the forces incurred by a conventional heat shielded landing.

As a consequence you inevitably end up with a lander/ascent vehicle massing well over 30 tonnes fully fueled. And that's optimistic.

And this raises the obvious question: what if you could reduce the mass of the ascent vehicle. Now stepping aside from any form of reusability the answer is a purpose built, single use vehicle. The problem is that in terms of overall mission design you've still got to land the ascent vehicle in the first place, and then you need a separate lander for the crew. Those mass multiplier effects have to be kept in mind before ruling out a reusable design. Now, within the context of reusable lander/ascent vehicle how do you minimise the mass at takeoff? Well the answer is to take away hypersonic reentry. That eliminates the need for a conventional heat shield. It also considerably reduces the stresses on the vehicle and allows it to be pared down to a few tonnes.

So to cut a long story short, this is a lander/ascent vehicle that masses 18 tonnes fully fueled. The essential vehicle is a few tonnes. With crew, and basic life support its closer to 5 tonnes. Those numbers give it a delta-V capability of 4.5Km/s with methane/oxygen fuel. Which is sufficient for an ascent to any orbit, with margin.

Which then takes me to my next controversial decision. That is to use mostly propulsive landing. The key assumption here is to start from a low Mars orbit and then deorbit leaving your velocity at just over 3.5Km/s that's a velocity reference to a inertial frame. Remember, the planet itself spins within this frame.

At a first approximation, were you to do an burn that zeroed your velocity you'd have used 3.5Km/s of delta-V from your engines. Do this at a sensible altitude (say 100Km) before you've the density is of any consequence. Now if there were no atmosphere (and the planet wasn't spinning) you'd free fall vertically and hit the surface 229 seconds later at 871m/s. In this theoretical case you might apply another 0.9Km/s of delta-V. Now you've consumed 4.6Km/s. That's your worst case. And as it happens, the vehicle, fully fueled is capable of almost that.

But physics is on your side. Even in the stall and drop scenario above, you've still the assistance of air drag. It turns out you hit the surface at typically 300-400m/s (depending on assumptions about Beta - mass to effective drag area). You never go past Mach 3.

And that's according to a simulator I wrote earlier this year. What I didn't factor into the simulator was that the planet (and its atmosphere) are rotating - typically around 400m/s in zones near the equator. What this means is that relative to the air and to the surface you're already going that much slower.

So in the worst case (non optimised) you need just over 4Km/s of effective delta-V from your lander.

Now, further optimisation means taking advantage of the modest thermal loads at supersonic or near hypersonic velocities, again relative to the air. Which means slowing down to a relative velocity or more like Mach 2 - about 500m/s. Taking that into account, and tweaking the simulator (bending the frame of reference) what happens is your vehicle again accelerates to over Mach 3 before again meeting the thickest part of the atmosphere (around 25Km) and then slowing to a terminal velocity closer to the above 300-400m/s mark.

Now, allowing for terminal maneuvering and other losses and my best guesstimate is you can land mostly propulsively with the equivalent of 3.8Km/s of delta-V. Now, this figure may be improved upon, but there is a diminishing return from doing so. Likewise there is a chance I'm out, but that doesn't make much difference to the masses of fuel we're dealing with below.

That's a mass ratio of just on 3, or an all up mass of 15 tonnes. 10 tonnes fuel.

What this amounts to is a light weight vehicle that relies upon propulsion to avoid hypersonic air travel. It does rely upon some modest and light weight thermal protection in critical areas. But the temperatures involved are in the hundreds of degrees C. It also enables other design refinements like not having to have deploy-able landing gear or engines hidden away behind a shield. There are basically very few surfaces not made of metal or protected by cryogenic fuel. There are key systems that do need protection but this involves only a small amount of thermal protection mass. Think high temp textiles and ceramic fibres. Even the landing legs are not going to get too hot for a decent high temp alloy.

Supersonic retro propulsion is unavoidable here. But SRS is unavoidable in any large payload delivered to Mars. Its a key enabler. We need to research it, and then we'll discover where the envelope is. You'll notice that I'm doing hypersonic to supersonic deceleration high enough that the density of the air poses no stability problems (by definition basically).

Now that I've argued for the feasibility of a most propulsive landing, at least for small payloads such as a crew, I'd like to show how that principle can then inspire and integrate into an overall architecture.

Aside from a reusable lander/ascent vehicle I need only two more kinds of vehicles. One is a big dumb booster (LH2/LOX). The other is what I've usually called a space hab, but lots of others refer to this as a transit vehicle. My transit vehicle is unsurprisingly a crew compartment, tankage, and propulsion. One element of design is to put a docking portal axially on both ends. An arrangement that allows chaining. One end of the vehicle the crew compartment narrows (think sleeping quarters, radiation shelter). Around that is the tankage and a minimal propulsion unit (I'll explain why later). So the pressurised crew compartment has the full diameter on one end but at the other is basically a tube that connects to the portal.

I'm going to propose that minus fuel but including life support, consumables and so on, this vehicle is going to mass 20 tonnes. Now I know what you're thinking. Just hang on a sec.

In transit to Mars, and return, there will always be (barring emergencies) two of these vehicles, either flying in formation or docked together. Hence an all up mass of 40 tonnes (without fuel) and a target crew of four. In practice during flight you will probably want to dock these two units and then optimise their internal space, allowing more room in one, and more mass and thus radiation protection in the other.

Its a concept I originally hit upon because it allows mass to be transferred out of one vehicle and into the other, so that one vehicle can do something particularly costly in terms of energy, like moving from high Mars orbit to low Mars orbit and then back. As it turns out I don't think that's needed, but there's a dozen good reasons for having two essentially identical vehicles, including recovery from damage, systems failures and so on. I should add that these have the capability to transfer fuel between them.

Now, as far as the overall architecture goes we're going to use L2 as a staging point. (Look I'm not that fond of integrating moon based activities but what appeals to me about L2 is basically avoiding a high orbit that dips in and out of the radiation belts).

So conceptually this mission begins and ends at L2, but we'll discuss how to get there an back too.

Each transit vehicle has fuel capability for around 2.2Km/s or about 18 tonnes of fuel each. That's rated as sufficient to return from high Mars orbit to L2 under nearly all circumstances and that also includes propulsive capture. Now I did toy with the idea of aerocapture but in the end, I'd rather have the reserve of fuel even if its not fully used. The on board fuel is there for the return journey from high Mars orbit to L2. A typical mission would use more like 15 tonnes of fuel or less (per hab unit). Minimal energy trajectories might use as little as 12 tonnes. Understandably the systems on the vehicle include the ability to keep the fuel at near zero boil off.

The forward journey from L2 to high Mars orbit is simply a boost with a conventional H2/LOX booster. There are of course other ways to do this but I'm going to keep the exposition simple.

One thing I would like to have (not essential) is a top side docking port for the lander/ascent vehicle. Doing so means it can dock with the transit vehicles (at either end).

Now, the stack as it travels to Mars is a booster docked with a transit vehicle, which in turn is docked with a transit vehicle, which in turn is docked with an (upside down) lander. What we have here is bump into place assembly. No need to get out with a spacesuit and a spanner.

The lander at the top of the stack is pre-fueled with enough fuel to reliably transfer to low Mars orbit. Which is about 2.5 tonnes of fuel.

Now, lets step back in time. Prior to any of this happening, a number of missions have taken place. Leaving infrastructure in place, including the oxygen production plant on the surface and a store of methane. You could go so far as to have a fully fueled ascent vehicle waiting as a backup. I'm also going to gloss over all the other stuff you need including the surface hab, rovers, etc. All of which is out of scope here.

What is important is that for every manned flight to Mars, there is another unmanned flight. Now this happens on a slower, lower energy trajectory, again using a big dumb booster. This unmanned flight places key hardware in low Mars orbit. This is time to arrive in low Mars orbit prior to the crew arriving in high Mars orbit. Arguably this could or should happen one cycle earlier.

The unmanned flight delivers to low Mars orbit two items. One is another identical transit vehicle / space hab. Identical in form and design, but limited in consumables. However its fuel tanks are completely filled. And one little quirk. The other item is a lander which is also fully fueled. One design modification places a liquid methane tank (3 tonnes) in place of the crew compartment. The basic mass of this vehicle is 3 tonnes. Again, the booster, space hab and lander stack end to end through their docking ports.

Upon arrival on Mars, the lander that forms part of the unmanned flight is free to land. Thereupon it delivers its cargo (3 tonnes of methane) to the surface. And an oxygen production plant in the form of a rover, provides the necessary cooling system. The lander is also used as an oxygen storage vessel. The space hab (it doesn't quite make sense to call it a transit vehicle) now remains permanently in low Mars orbit. Its function is to provide a safe haven - because the lander has limited life support.

Now back to the manned mission. On arrival in a high Mars orbit, the crew transfer to the lander and use its fuel to transfer to the space hab in low Mars orbit. They wait whilst the lander takes on fuel. The crew transfer again to the lander and it departs and lands on the surface. A full surface mission takes place. Sometime during this period the lander is refueled. The crew again use the lander to ascent to orbit, docking with the space hab. Again they wait for the lander to take on sufficient fuel (about 2.5 tonnes) to return and dock with the transit vehicle which has remained for this time in high Mars orbit.

The stack in high Mars orbit comprises the two separate space habs docked together (as they were when the crew arrived) and the lander which is docked to one end. Returning from high Mars orbit to L2 involves using the engines of the lander. Note that the two space habs have their own minimal propulsion which could achieve the same goal, albeit with reduce efficiency (a much longer burn). Note also that the lander draws down fuel from the space habs.

In a nominal mission this will leave roughly a third of the fuel on board. This fuel will be consumed to propulsively capture back to L2. During transit the fuel is shifted towards one end to provide a lower radiation environment on that end - up until its used.

From L2 it is possible to either use the lander to approach low Earth orbit, or to supply a lander capable of directly landing on Earth.

Now taking a longer view, what happens is that the space habs are cycled through several phases over multiple missions. A typical life cycle involves going to L2, becoming part of a manned mission, returning to L2, then returning to low Earth orbit to be used as a vessel for resupply. Then returning to L2 and delivering consumables. Then the same unit is sent as part of an unmanned mission being delivered to low Mars orbit. So a typical unit would see service over 2 cycles or close to 5 years.

You know personally, I'd prefer to see things being recycled more than this but in the end something has to deliver fuel to low Mars orbit. The practical consequence of all of this is that over several missions there may be 2 or 3 (or even 4) functional space habs in low Mars orbit. Which is probably to the good. In the end though, something that is designed to have multiple copies made is probably going to benefit from economies of scale and will improve over time.

Notice also that I've basically designed this so that one singular methane/lox engine design accounts for both the multiple engines (probably 8) on the lander and the 2 or so on each space hab. And as time goes by there are going to be spares everywhere.

And its exactly the same life support system and ancillary equipment everywhere.

The really nice thing about this architecture is that its very resilient to failures. You can for instance limp home in one space hab. You can transfer fuel where needed. Which means in emergencies you can borrow fuel from one vehicle at the expense of crashing the other. In a similar fashion its also resilient to irreparable damage to fuel tanks, life support systems etc.

Another feature is that for just a little more fuel you can use the lander as a taxi to the Martian moons. I think everyone would jump at the opportunity to tie together both a surface mission and sample returns from the moons.

Now, seen from L2 the mass of the manned mission is about 80 tonnes. And the unmanned part about 60 tonnes. Given that the assumed delta-V from L2 to high Mars orbit is 2Km/s then the mass ratio for a H2/LOX booster is about 1.56. So a booster of about 45 tonnes for the manned mission. Which places it strategically below 50 tonnes. Likewise the two fueled space habs and the lander are flown separately. Similarly for the booster for the unmanned mission.

Whilst this takes multiple separate parts, none are over 50 tonnes. The only point where mass does add up is considering the LEO to L2 part of the system.

Oh, before I forget. The lander that flew back with the crew. It switches roles back at L2 and goes on to the next unmanned flight and ends up as the vehicle that transfers fuel to the surface.

Now between LEO and L2 things get more complex because not everything has to be brought up to L2 every time. So I'm not going to bore with detail, but in the worst case, using conventional technology, you're looking at very roughly 250 tonnes all up in LEO terms. Plus some basic infrastructure at L2.

This is where I think ultimately we're going to end up with a solar tug and where it makes the biggest difference. And that's from LEO to L2 and back. Sooner or later may wish to refurbish things in LEO and also consumables and fuel make up the majority of the mass. If so the combined mass per mission is around, or a bit under 200 tonnes.

At this point I have to say that even the fully conventional boost to L2 situation is a considerable improvement in terms of overall mass than anything I've seen from NASA.

Now step back and consider this. Initial mass in low Earth orbit is not everything. I can think of ways to halve it. Cutting corners, using aeorcapture etc. But as the cost of launching fuel into low Earth orbit comes down, it will become irrelevant. Indeed, a solar tug only becomes really viable over multiple missions.

The thing I keep coming back to is simplicity, redundancy, interoperability of parts. I've basically proposed two key vehicles. And anything else is just a minor design change on those two vehicles. This is why I've proposed to land fuel (for the next ascent) propulsively with another lander. Its not like I couldn't save a few tonnes there, but it means design effort can be focused on just one thing.

Now, there will be other things we'd need to land on Mars. I'd argue there's no real need to land anything heavier than about 15 tonnes. And the reason is a lot of mass is just consumables and fit out. Its not like the crew aren't being paid well. There are design challenges for such large loads but I'd submit two things.

One is it isn't as bad as NASA thinks. You don't need 100 tonne class entry mass vehicles to land 30 something tonne payloads. You'd never need more than half that. Second, whilst its a challenge, its not beyond our current knowledge plus (and again the key enabler) a better understanding of SRP.

My key point in all of this (and its why I came here in the first place) is to demonstrate why its a good idea to stop thinking in terms of landing people inside very heavy payloads and instead land them separately in purpose built landers with higher margins.

And the flip side of that is, it also makes it easier to design the landing system for the non manned large payloads, because we're not worrying about the crew there.

There are of course other features to this architecture I like. For instance you can extend it to provide (a small amount of) gravity en-route.

Anyhow, enough verbage. What do you guys think?

Last edited by Russel (2013-06-18 11:57:24)

JoshNH4H · 2013-06-18 12:20:13

So, in summary: You propose that, rather than using aerobraking for atmospheric entry, we should instead do a breaking burn followed by powered descent; and that further, rather than importing Hydrogen for use in ISRU that we should import methane. You also propose a reusable launch/descent vehicle for crew and (am I reading this correctly?) for cargo. Rather than starting in LEO you instead suggest that the mission should start in L2.

Your justifications for this (which I intend to argue, once I can be sure that I'm arguing with what you actually said ) are of course contained in your introductory post.

Russel · 2013-06-18 22:04:49

JoshNH4H wrote:

So, in summary: You propose that, rather than using aerobraking for atmospheric entry, we should instead do a breaking burn followed by powered descent; and that further, rather than importing Hydrogen for use in ISRU that we should import methane. You also propose a reusable launch/descent vehicle for crew and (am I reading this correctly?) for cargo. Rather than starting in LEO you instead suggest that the mission should start in L2.
Your justifications for this (which I intend to argue, once I can be sure that I'm arguing with what you actually said ) are of course contained in your introductory post.

Yes, I'm proposing a mostly propulsive landing. One aspect of which is a braking burn at high altitude. The key characteristic of this descent is making as much use as possible of aerobraking but within the capabilities of a modest TPS.

Yes, I'm proposing importing methane. Not going into all the issues with importing hydrogen, that decision was made out of sheer simplicity.

Yes, I'm proposing a reusable vehicle for both landing and ascent. With the caveat that in this instance its only used 3 times in an atmosphere. First is a crewed descent. Second is a crewed ascent. Third as an uncrewed descent fuel delivery.

Yes, I've proposed L2 as a staging point.

JoshNH4H · 2013-06-19 11:54:32

Assuming that these are the fundamentals of what you're proposing (nothing else major pops out to me) I would simply like to raise objections that you probably expected, but that I don't think were sufficiently addressed in your post.

The mass penalty for a propulsive landing is very significant. Launch costs may have been decreasing of late but I don't think you can justify that kind of mass increase in the mission. Especially not given that you don't intend to import Hydrogen. By the way, the sabatier reaction is not complicated. Zubrin demonstrated via experimentation that it's actually quite simple. Hydrogen storage itself is not an issue, as long as you're willing to accept some level of boiloff. Given proper insulation and a reflective coating to minimize sunlight, plus perhaps some level of active refrigeration I would expect it to be quite small. This gives you a fourfold reduction in imported fuel mass over importing methane. By the way, Oxygen may be present but the Reverse Water Gas Shift reaction is probably the best way to get at it. While it has been shown that electrolysis of Carbon Dioxide is possible, it's a very inefficient procedure relative to water electrolysis. It's also less well-developed, meaning that the equipment will have worse specific power and specific daily output.

Further, enabling technologies for landing on Mars are not far away. The big one is inflatable heatshields, and work is being done on these at NASA as we speak. Beyond that, we need to get a better understanding of the dynamics of the Martian atmosphere, which could be accomplished by sending balloon-type probes to different levels and observing their drifts, and allowing them to measure temperatures (Thermometers and GPS units are very light, easily less than 50 grams together. I'd imagine a barometer could be built to be extremely lightweight using piezoelectrics.) We might also want to fit out a couple with mass spectrometers to analyze local composition, including humidity.) or perhaps even through the use of a more typical weather satellite, of the kind we have orbiting Earth. Beyond that, improved non-propulsive deceleration (both in the form of aerocapture and aerobraking) is an important enabling technology for the colonization of the inner solar system, and it's better that we develop it sooner rather than later, so as to squeeze out maximal benefit.

With regards to L2 vs. LEO or another Earth Orbit, it doesn't really matter where you do the assembly. If you do it in LEO you have the benefit of shorter round-trip communication times with the hardware for more direct control/backups. While the outbound delta-V is higher, it takes fuel to launch to L2 and this will cost every bit as much as launching straight out of LEO.

I would raise significant concerns regarding the use of a reuseable vehicle on the very first mission, when there are no people out there who can fix the rocket or even check it out properly. The second use represents a huge safety issue, easily leading to loss of mission and loss of crew if anything goes wrong.

Russel · 2013-06-20 00:18:37

JoshNH4H wrote:

Assuming that these are the fundamentals of what you're proposing (nothing else major pops out to me) I would simply like to raise objections that you probably expected, but that I don't think were sufficiently addressed in your post.
The mass penalty for a propulsive landing is very significant. Launch costs may have been decreasing of late but I don't think you can justify that kind of mass increase in the mission. Especially not given that you don't intend to import Hydrogen. By the way, the sabatier reaction is not complicated. Zubrin demonstrated via experimentation that it's actually quite simple. Hydrogen storage itself is not an issue, as long as you're willing to accept some level of boiloff. Given proper insulation and a reflective coating to minimize sunlight, plus perhaps some level of active refrigeration I would expect it to be quite small. This gives you a fourfold reduction in imported fuel mass over importing methane. By the way, Oxygen may be present but the Reverse Water Gas Shift reaction is probably the best way to get at it. While it has been shown that electrolysis of Carbon Dioxide is possible, it's a very inefficient procedure relative to water electrolysis. It's also less well-developed, meaning that the equipment will have worse specific power and specific daily output.
Further, enabling technologies for landing on Mars are not far away. The big one is inflatable heatshields, and work is being done on these at NASA as we speak. Beyond that, we need to get a better understanding of the dynamics of the Martian atmosphere, which could be accomplished by sending balloon-type probes to different levels and observing their drifts, and allowing them to measure temperatures (Thermometers and GPS units are very light, easily less than 50 grams together. I'd imagine a barometer could be built to be extremely lightweight using piezoelectrics.) We might also want to fit out a couple with mass spectrometers to analyze local composition, including humidity.) or perhaps even through the use of a more typical weather satellite, of the kind we have orbiting Earth. Beyond that, improved non-propulsive deceleration (both in the form of aerocapture and aerobraking) is an important enabling technology for the colonization of the inner solar system, and it's better that we develop it sooner rather than later, so as to squeeze out maximal benefit.
With regards to L2 vs. LEO or another Earth Orbit, it doesn't really matter where you do the assembly. If you do it in LEO you have the benefit of shorter round-trip communication times with the hardware for more direct control/backups. While the outbound delta-V is higher, it takes fuel to launch to L2 and this will cost every bit as much as launching straight out of LEO.
I would raise significant concerns regarding the use of a reuseable vehicle on the very first mission, when there are no people out there who can fix the rocket or even check it out properly. The second use represents a huge safety issue, easily leading to loss of mission and loss of crew if anything goes wrong.

Yes, you're correct. These were expected objections

The mass penalty for a mostly propulsive landing is what it is. Two thirds of your mass on entry is going to be fuel. So question is the difference in mass between a mostly propulsive landing and a conventional landing. And the more important question is whether that can be justified in terms of:

a) overall dollar budget
b) development cost and complexity
c) simplicity of design
d) robustness, resilience and safety

Considering only the mass involved in transporting the crew from low Mars orbit to Mars surface and back again, my proposal looks like this.

The crewed lander is posited as being 5 tonnes of mass, without fuel, but including crew and life support consumables.
The lander which delivers methane to the Martian surface is posited as having a dry mass of 3 tonnes.
The landed methane. That's 3 tonnes.
The fuel needed to land the methane. 12 tonnes.
The fuel needed to land the crew. 12 tonnes.

So accounted for in this way, the actual process of landing a crew on the surface of Mars, and bringing them safely back into a low orbit, involves 35 tonnes delivered to low Mars orbit. I hasten to add that with the architecture I propose, all the mass (fuel, hab and lander) delivered to low Mars orbit has first been delivered to high Mars orbit and then aerobraked to low Mars orbit. (And strictly accounted for, the mass of the crew does not belong in the above accounting, but I'll leave it this way to be conservative.)

Now before going to look at alternative approaches to the problem, I'm going to make the observation that the 35 tonnes delivered to low Mars orbit in order to perform this function, compares to the 70 tonnes of mass delivered to high Mars orbit, which is the transit vehicles and their fuel supply. In other words, if a third of your mass budget is spent on safely landing the crew and returning them as far as low Mars orbit, then I'm comfortable with the result.

I should also add that there is additional mass in the habitat vehicle that is transferred to low Mars orbit and is used in part as a fuel depot. Its difficult to know where to account for that because in this context, much of the mass of that vehicle is actually life support and serves the purpose of providing a home in orbit for the crew. But I do acknowledge that these numbers are neither perfect nor final.

A conventional single use lander suffers its own mass multipliers. It needs to withstand the forces of a conventional landing. Meaning several tonnes of basic structure. Add to that the heat shield. Add to that unavoidable propulsion system which still needs to provide more or less the same thrust even if for less duration. And now its easy to see why Dragon heads north of 5 tonnes. Factoring in the uncertainties and the need for larger drag devices for Mars, I'd suggest that 7 tonnes is a conservative figure for a fully fueled lander of this type. Some have suggested much higher than this.

In addition you need an ascent vehicle landed separately. Now, with a minimalist vehicle (such as my own) then you have a take off mass of 5 tonnes (plus fuel). Some of that mass is crew. And some of it is life support consumables that can be produced on the surface. So lets be fair and call it 4 tonnes of ascent vehicle.

Then there is the imported methane fuel (I'll consider hydrogen shortly). A 5 tonne ascent vehicle needs 13 tonnes of fuel. That's 2.88 tonnes of methane. If its imported you probably need 3 tonnes of methane to start with. Now to import the methane you need a vehicle to land it in. The most economical way of course is to land the methane with the ascent vehicle, bringing the total ascent vehicle landed mass to 7 tonnes.

A conventional landing craft has a mass penalty of about 2 to 1, so 7 tonnes of payload means 14 tonnes at entry.

Which takes us down to a comparison between 35 tonnes delivered to high Mars orbit, then aerobraked. Versus 21 tonnes at Mars atmospheric entry (we assume direct entry). The former amounts to 111 tonnes in LEO. The latter amounts to 57 tonnes in LEO.

Total saving: 54 tonnes in LEO.

Now, that's nothing to be sneezed at. At $2000/tonne that's a saving of a bit over $100M in launch costs per mission.
But that should be seen against the backdrop of billions in development costs. And development costs are driven by complexity and time.
I do agree that in the long term, if we decide we want to return to Mars many times over, then an indigenous source of hydrogen is the key.

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Now, to do a fair comparison regarding the importation of hydrogen I'm going to use as a baseline the conventional mission described above. A conventional 7 tonne single use lander and a previous landing of an ascent vehicle with methane fuel on board which added up to 14 tonnes.

From this baseline we derive a version where hydrogen is imported instead of methane.

I'm first of all going to introduce hydrogen losses into the equation. Its not just boil off, its also the losses incurred by passing hydrogen through the plumbing and heat exchangers of a ISPP plant. In short, hydrogen is innately leaky. I'm going to assign a factor of 50% for overall loss. That's even with very good insulation and some measure of active refrigeration, but allowing for leakage and for a safety margin (we're assuming the hydrogen is our first and only source of ascent fuel). With that we need 1125Kg of hydrogen.

Landed in a conventional lander, 1125Kg of hydrogen versus 3000Kg of methane translates to a difference of 3.75 tonnes at Mars entry. With the direct entry delta-V of 4.6 assumed above, that translates to a saving of 12.8 tonnes initial mass in low Earth orbit.

Now that's the scale of the maximum possible theoretical saving. About 13 tonnes of mass, and about $26M of launch costs but in return all the issues and risks inherent in hydrogen.

Of course if you decided you wanted to use mostly propulsive landers for other reasons, and still wanted to deliver hydrogen rather than methane. Then the saving in IMLEO terms in delivering hydrogen would be roughly 6 tonnes. Of course that's the theoretical maximum saving.

Beyond that theoretical best case the real world detail doesn't look pretty.

For a start, the ISPP plant needed to produce methane from hydrogen is likely to be more massive than the alternative which simply produces oxygen. Possibly by a factor of 2. The reason is that it involves more parts and more plumbing and heat exchangers (or else you sacrifice efficiency). Now its fair to say that such a plant is likely to see service for years so it can't be accounted for on a per mission basis, but the issue is there. The same applies to the power source. And again the task of producing methane from hydrogen is much more energy intensive. For two reasons. One is the inefficiencies in the process, especially if done on a small scale. The other is the refrigeration costs. Often we're dealing with very hot gases but the ultimate product is a hard cryogen. I won't go further into this but its a real issue.

More importantly in terms of its effect on mass per mission, is that 3 tonnes of methane occupies a bit over 7m3 of tankage. 1.125 tonnes of hydrogen occupies closer to 16m3 of tankage. That presents problems both on the ground and more importantly in terms of how you land it. Any lander, mostly propulsive, or conventional suffers the issue that larger tankage translates to larger structural elements. Worse, the conventional lander is hit hardest because with the higher stresses, a large tank of hydrogen just multiplies the mass of structure and heat shield. This factor alone probably removes half of the real advantage of hydrogen.

There are also other system wide mass costs from importing hydrogen that go all the way back to supply of further power and refrigeration equipment right back to low Earth orbit. In the end the advantage of importing hydrogen over methane (even assuming conventional landers) may evaporate.

This is not to say that hydrogen isn't the future. It is. Its indigenous hydrogen we need. And for that we need machines that dig and process dirt. And that's quite probably an argument for having people there, to manage the prototypes. But I'm getting off track here.

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Back to the big picture. A singular lander and ascent vehicle design is an exercise in reducing complexity and saving development costs. It also affords you a platform that serves many needs. For one thing, a lander becomes the taxi that transports crew between high and low Mars orbits, thus enabling the practice of parking the transit vehicles in high orbit. That factor alone saves the 2.6Km/s of delta V otherwise involved in bringing the transit vehicles down to low orbit and back. That would double the overall mass in and of itself.

Next I'm using the lander as the primary propulsion source for the return journey. Now the transit vehicles have their own engines but these are for backup purposes and its because we're relying upon the lander as the main engine we can then mass reduce the transit vehicle.

Next, the lander is capable of running surface to surface hopping missions. Meaning that its capable of long distance hops from base to base.

Next, the lander is capable of simple missions to the Phobos and Deimos. Indeed for a reasonable amount of fuel one could stop to claim samples from those moons on the way back to high Mars orbit.

Next, a mostly propulsive landing isn't as given to the vagaries of the atmosphere and has more capability to land on higher landing sites.

These synergies actually reduce the total mass of doing all of these things in other ways.

....

Inflatable heat shields are a great idea - if you plan on an expendable vehicle. I've no doubt that they, or something similar will come in very handy when it comes to landing large objects on Mars. But the problem with an inflatable heat shield is that it is still added, mass complexity and risk. And the bigger it is the more stability issues it has. Again, a necessary evil for large payloads, but lets make these large payloads unmanned.

The more you rely upon stuff that you eject during landing, the more you end up with the cost (and risk) of garbage.

Now, I would be willing to consider that in the case of a mostly propulsive, reusable crewed lander, a supersonic decelerator (probably a towed one) could be useful. But only to provide further margin. In other words you can land without it but it means you have more time to consider your landing site. Here you have two advantages. Firstly you're not dealing with extremely high dynamic loads and temperatures and you're not contemplating operating above Mach 3. Which means a lighter decelerator. Think in terms of being man handled and replaced easily. The other advantage is that because of the lower velocities involved you can afford to deploy the decelerator virtually from interface, or as soon as there is enough density to make it worthwhile.

I agree that by the time we get to land on Mars we will probably have real time, detailed observations of the atmosphere. Even a mostly propulsive landing would (especially if the landing site is higher) require some consideration as to the atmospheric conditions.

Which brings me to another point. One bug bear I have with direct entry is that you don't get much choice about the atmospherics. But if you based your architecture around having a safe (life support etc) haven in low Mars orbit, you're then free to choose when you land. Which means you can time it for the best atmospherics. And again, having an outpost in low Mars orbit does come at a cost, but it does have its advantages in terms of safety issues like this, plus as I said, you can keep most of your transit mass in a high orbit.

----

In some ways I'm not particularly fussed about L2 versus LEO. Certainly if you're in the business of large scale fabrication (over sized heat shields, trusses etc) then I'd definitely go for LEO. Remember that what I propose does not involve anything more than bump together assembly. No EVA required.

However the bigger issue is that you get the choice between either assembling (or rather docking) everything in LEO and then using conventional propulsion to Mars, or if you're going to adopt a solar tug then L2 is possibly the better option. Yes, you could just use a solar tug to lift the whole thing into a high elliptical Earth orbit but the whole thing would be traversing the radiation belts on every orbit. You'd then have to time the crew transfer well and get them on board quickly. Possible, yes.

L2 does give you more time to do check out systems (again that's in the context of using solar tugs).

I guess if you pushed me, I'd go for a conventional launch from LEO. At least at first. But remember that if you're seriously concerned about mass delivered to low Earth orbit, you wouldn't do that. You'd develop the solar tug first. Again, its useful to have a sense of perspective.

....

Now regarding the final point about reliability. For me there really are two enabling technologies that we need to go to Mars under any architecture that we don't have.

One is supersonic retro propulsion. A fact of life even for more conventional landers. And I need to emphasise this. Large payloads are never going to go subsonic in the Mars atmosphere. We need to understand SRP. Simple as that.

The other is enabling technology is developing a methane/lox engine that's very reliable. And that's the case even without a propulsive lander. As soon as you rely upon methane/lox engines to get you home, then you need an engine that can survive long burns, multiple restarts and long idle periods in space conditions.

It also has to be added that there's no such thing as a non propulsive lander. Sooner or later you're going to have to use retro propulsion. And you'd hope your engines are reliable.

What I've done though in having an architecture that relies upon a mostly propulsive landing for the crew is that I can achieve the goal of having one universal engine design for all purposes. And only having two basic vehicles. Which means a lot of the supporting hardware is of common design. Which means the capability to swap components.

(The above sentence ignores the big dumb H2/LOX booster which I think everyone takes for granted).

Then I've built in multiple layers of redundancy. Typically the lander itself would be flying with 8 engines. Four directly down. Four pointed about 15 degrees to the side. There's a certain amount of redundancy there that I won't go into right here. Where redundancy also comes in is that instead of landing the ascent fuel in a purpose built, expendable vehicle, I am instead landing the fuel in another lander. Which means that there's a source of spare parts and spare engines waiting on the surface. Indeed it might be possible in an emergency to retrofit the fuel lander as a crewed ascent vehicle.

The redundancy continues in space. Instead of discarding the lander, I'm bringing it back to orbit and then on to high orbit. There it provides the primary source of propulsion for the return journey. But if it fails there are still 2 engines on each of the transit vehicles. Again, its designed so that you can swap over engines if needed.

Your final fall back is to abandon one of the space habs, transfer fuel across and limp home (Ok, its cramped but you're safe). I don't see any other architecture providing that final layer of safety.

And technically there is one final refuge, and that's the lander itself.

----

I hope that answers things?

Btw, I was hoping to be asked about turbulence

Edit: Ok, now I've managed to get this post into some readable shape.

Edit 2: Since I brought up turbulence I might was well answer it.

This is somewhat more speculative. I'm simply making an argument that a good solution may be possible, at least for a vehicle without a major heat shield.

What little evidence I have about SRP is what can be found on the web. Some of it discusses the reduction in effective drag coefficient caused by an engine pushing the air away from the base of the vehicle. Some of it is the video showing clearly the turbulence that is caused by firing an engine into a supersonic air stream. Its fair to say that a lot of discussion about SRP is framed by the assumption that the engine or engines will be project through a conventional heat shield. Obviously I'm stepping outside that convention. And what I can't find is information that relates the nature of the turbulence to the density and relative velocity of the stream. Nor is there information about the nature of the turbulence. What the physical and time scales are and again, how those are governed by density and so on.

There are basically 3 regimes in which a mostly propulsive lander would have to work. The first is essentially outside the atmosphere. Here you have no constraint since there is no atmosphere to provide turbulence. In doing my rough math I've figured that this regime extends down at least to about 100Km. Now the reason for getting this point as low as possible is that the lower you do your initial large braking burn, the more efficient you are because of lower gravity losses.

The key question is, when does the atmosphere become dense enough for significant turbulence to apply. It might exist, but the forces might be too small to be of consequence. If I recall correctly, at around 100Km the air is still measured in micro grams per cubic metre. Its also quite possible that thermal considerations might kick in before turbulence from propulsion becomes an issue. In other words, it becomes necessary to apply (some of) the initial braking burn to keep heating below an acceptable level. The acceptable level I'm thinking of is around 0.3W/cm2. In other words something akin to being exposed to 3KW per m2. That's about 2 suns in Earth orbit.

The third regime is below the point at which full retro propulsion is required in order to effect a safe landing. Which is probably somewhere below 15Km. The key here is that there is a trade off involved. Wait longer and normal atmospheric drag will continue to slow your descent and thus lower the effect of turbulence when you apply retro propulsion. Wait longer and you need to apply more thrust.

The second regime is obviously between those two extremes.

Now, one thing that can help is to move your engines away from the body of the vehicle. This has two potential benefits. One is that the region of lower pressure created by the engine can extend away from the body of the vehicle (especially with engines directed away from the vertical) and this should have the effect of reducing the effect of propulsion on normal drag. A lot of numbers I have seen give effective drag coefficients based on engine thrust, but do so on the assumption that the engine is directly below the vehicle and thus will create lower pressure under it. The other reason for moving the engines away from the vehicle is to remove as much of the turbulent wake of the engine as possible away from the body of the vehicle. Now the theory here is that the effects of turbulence are only important to the extent that changing air pressure can react against something.

So in essence you end up with a design that has engines mounted on a four star truss that projects beyond the body of the vehicle and also keeps the engine as high as possible. Quite probably integrated with the design of the landing legs. On each point you have two engines. One directed downwards. The other projected about 15 degrees off vertical and outwards.

Now, in regime 1, the key is to brake as late and as hard as possible. So you use all 8 engines at first, and then switch to the 4 engines that are pointed outwards and continue to complete the initial brake. My hope is that you can do this well below 100Km.

In regime 2, its yet to be determined what the optimum is. One suspects the optimal strategy is either a continued low thrust, or none at all (for the sake of engine efficiency at lower powers). An advantage of continued low thrust perhaps is to avoid re-ignition problems further down.

In regime 3 I suspect that you'd use the outer 4 engines at low thrust initially and then gradually build thrust as velocity falls. As velocity and thus turbulence becomes less of an issue you then bring in the 4 vertically projected engines and quickly go to full power. And of course as you approach final landing you'd throttle back and use the outer 4 engines.

Now all of that is based on the educated guess that turbulence can and will develop but so long as it has minimal surfaces to react against, its manageable.

And the nice thing about a mostly propulsive landing is that its easier to get away with doing that. Because your engines don't have to be hidden behind a heat shield they can be positioned where you like them and away from the body of the vehicle

Last edited by Russel (2013-06-20 09:21:49)

GW Johnson · 2013-06-21 14:11:32

Hmmmm.

Entry heating protection at Mars is nowhere near as difficult as at Earth. Peak entry heating varies as entry interface velocity cubed, and it's a whole lot lower at Mars, by factors of 2 to 3.

There's a really good semi-reusable solution with Spacex's PICA-X ablative, which could be flown several times, maybe even a few dozen times, before replacement. And, it's fairly low density at 0.27 g/cc, you need around 2-3 inches of it for "long" life, and to cut off the conductive heat load.

Except for the demonstrated fragility and high maintenance costs, low density ceramics of the shuttle-tile type could be a longer-life solution. It takes about half an inch or so of that stuff to cut off conduction, and be processible for bonding the tiles. Except for damage repairs, the life is theoretically infinite.

I have an oddball low-density ceramic-composite material that I made and tested for a different use over 30 years ago. It has the density of Styrofoam (roughly 0.03 g/cc) and very low conductivity (around 0.035 W/sq.cm). It's a two-component composite laid up sort-of like ordinary fiberglass, and features redundant retention, even in large panels. It proved very tough, very damage-resistant, when I tested it so very long ago. I'll present a paper on it at the convention in Boulder CO this August. Much less than an inch of it would be needed.

That gets you through the hypersonics. That's where aero deceleration is most effective on Mars, completely unlike Earth.

If your entry angle was shallow (say near 1.6 degrees at interface), you come out of hypersonics at local Mach 3. On Mars, that's around 0.7 km/s at around 5 km altitudes, typically (in anything big enough to carry people or tons of cargo). That's too low to deploy a chute at all, much less have it decelerate you any noticeable amount, so why bother? Just go to direct rocket braking-to-touchdown from there.

You will use less fuel in a lower mass vehicle than in any other imaginable scenario, done that way. I already looked at a multitude of approaches. That was the best.

Firing through openings in heat shields is no problem, just seal the engine compartment behind the heat shield so that there is no through-flow through the hole. You might need a tad of coolant gas injection into the sealed compartment to make up for the volume-filling transient as you descend, but it's a very minor massflow.

Plume stability for supersonic retro thrust is an issue often raised, but easily addressed by using multiple nozzles canted off centerline just a few degrees. Spacex knows of this stabilizing effect already, why else would the super Draco thrusters be arranged to fire canted at 45 degrees around the edge of the Dragon heat shield? You don't have to fire at 45 degrees around the edge, you can fire through ports in the heat shield canted at only 5-10 degrees and still get stability.

Landing big stuff on Mars is not really such a difficult thing, but you do have to do something different than we have been doing all these decades. Fear-of-the-new more than actual technical issues is holding us back.

GW

Russel · 2013-06-21 21:34:17

GW Johnson wrote:

Hmmmm.
Entry heating protection at Mars is nowhere near as difficult as at Earth. Peak entry heating varies as entry interface velocity cubed, and it's a whole lot lower at Mars, by factors of 2 to 3.
There's a really good semi-reusable solution with Spacex's PICA-X ablative, which could be flown several times, maybe even a few dozen times, before replacement. And, it's fairly low density at 0.27 g/cc, you need around 2-3 inches of it for "long" life, and to cut off the conductive heat load.
Except for the demonstrated fragility and high maintenance costs, low density ceramics of the shuttle-tile type could be a longer-life solution. It takes about half an inch or so of that stuff to cut off conduction, and be processible for bonding the tiles. Except for damage repairs, the life is theoretically infinite.
I have an oddball low-density ceramic-composite material that I made and tested for a different use over 30 years ago. It has the density of Styrofoam (roughly 0.03 g/cc) and very low conductivity (around 0.035 W/sq.cm). It's a two-component composite laid up sort-of like ordinary fiberglass, and features redundant retention, even in large panels. It proved very tough, very damage-resistant, when I tested it so very long ago. I'll present a paper on it at the convention in Boulder CO this August. Much less than an inch of it would be needed.
That gets you through the hypersonics. That's where aero deceleration is most effective on Mars, completely unlike Earth.
If your entry angle was shallow (say near 1.6 degrees at interface), you come out of hypersonics at local Mach 3. On Mars, that's around 0.7 km/s at around 5 km altitudes, typically (in anything big enough to carry people or tons of cargo). That's too low to deploy a chute at all, much less have it decelerate you any noticeable amount, so why bother? Just go to direct rocket braking-to-touchdown from there.
You will use less fuel in a lower mass vehicle than in any other imaginable scenario, done that way. I already looked at a multitude of approaches. That was the best.
Firing through openings in heat shields is no problem, just seal the engine compartment behind the heat shield so that there is no through-flow through the hole. You might need a tad of coolant gas injection into the sealed compartment to make up for the volume-filling transient as you descend, but it's a very minor massflow.
Plume stability for supersonic retro thrust is an issue often raised, but easily addressed by using multiple nozzles canted off centerline just a few degrees. Spacex knows of this stabilizing effect already, why else would the super Draco thrusters be arranged to fire canted at 45 degrees around the edge of the Dragon heat shield? You don't have to fire at 45 degrees around the edge, you can fire through ports in the heat shield canted at only 5-10 degrees and still get stability.
Landing big stuff on Mars is not really such a difficult thing, but you do have to do something different than we have been doing all these decades. Fear-of-the-new more than actual technical issues is holding us back.
GW

GW,

Your approach - an entry from low Mars orbit with a shallow angle - is basically my starting point.

As you've seen in my earlier posts, I had focused on how to design a reusable ascent and descent vehicle based around a conventional hypersonic heat shield. Its indeed possible, but the critical problem is the takeoff mass. Which then scales the amount of fuel you need to get it into orbit. And as you recall, the ascent fuel was headed into the several tens of tonnes region.

Taken in isolation that's not a problem - just a bigger ISPP plant. But when I went away and thought about how a reusable lander and ascent vehicle integrates into an overall architecture (which is the main subject of this thread) I kept coming back to the fact that the takeoff mass of the ascent vehicle (reusable or not) is an issue that has to be addressed.

Another point here is that no matter how I come at it, the problem of ascent fuel (the fuel, not the oxygen) doesn't have an easy solution. Despite the enthusiasm for hydrogen, the perceived benefits of importing hydrogen are reduced or negated by the engineering realities. And if you're importing methane, again its the mass of the ascent vehicle that is the root cause of most of your problems. Either way you can't afford to have an overly heavy lander and ascent vehicle.

Now, in designing a lander and ascent vehicle that uses a conventional heat shield I kept coming up with a dry mass in the order of 10 tonnes. And even I would concede that figure was, if not optimistic, certainly going to involve some careful design and a lot of testing. Even with ceramics or a thinner ablative, there is still the support structure. And in brief, the kinds of stresses involved translate to more mass. More mass translates to more fuel on take off. Which again scales the tankage and so on.

In brief, the problem seems to be how to build something as minimalist as possible as an ascent vehicle and then secondary to that is how to land that same vehicle without adding substantial mass. And the mostly propulsive approach does that. It doesn't mean you don't need *some* thermal protection, but as you're probably aware the temperatures involved in Mach 3 or 4 are a whole different ballpark when it comes to materials.

Remember that at the temperatures involved there's lots of things that don't really need protection. Much of the bulk of the vehicle is tankage and the fuel itself keeps that cold. You only really need a thermal blanket to lower the rate of boil off. There's really only a few external systems that need special protection (electronics etc). And the whole thing is likely to be more amenable to hands on maintenance - which I think is an aspect that's often neglected. Designing stuff to be field serviceable. One thing that keeps coming up when I visualise the lander is that the conventional heat shield tends to make stuff less accessible to maintenance and repair.

There are still good uses for light weight ceramics and I'll be keen to see your paper. As it happens, I keep toying with the idea of allowing the transit vehicle (space hab) to have the capability to aerocapture, and besides having its own native heat sink (tonnes of water) the thing I'd need is a light weight ceramic. I've already allowed for aerobraking (after capture) but that's something that can be done gradually. Aeorcapture back to Earth orbit would be a real advantage, but for the moment I've gone for propulsive capture but with the understanding that aerocapture is an option and therefore planning for propulsive capture leaves you with more margin.

Now, as far as SRP goes, my understanding of the problem is you wish to avoid having the turbulent gas stream reacting against the vehicle. Now the Spacex solution appears to be to mount the engines above the plane of the heat shield and at a reasonable degree of cant. Meaning their solution (if it works) solves the problem but only where you aren't demanding a lot of delta-V from the propulsion system.

I can't rely exclusively upon the same approach because I'm demanding a lot more performance from the system. Instead what I would like to do takes a little description. Imagine a vehicle that's essentially 4 tanks arranged symmetrically around the vehicles vertical center line. These are your methane and LOX tanks. There is a center volume between these tanks and that's a good place to put the crew compartment. The actual frame of the vehicle is (if you're looking down at it) a four cornered square truss. That also supports the landing legs, which remain deployed at all times. Along each truss there are 2 engines. One of those as at the far end and its canted at a modest angle (say 15 degrees). The other engine, which is inside the landing leg, but still outside the projection of the body of the vehicle is pointed vertically. Both engines are mounted above the center of mass of the vehicle.

Now the trick here is to maintain efficiency, but use cant where necessary. For a start the main braking burn can use all 8 engines (or even just the inner 4 if that turns out to be optimal). It doesn't matter because the air is too thin to matter. During the mid part of the descent you'd either run with no engines, or use a low thrust on the outboard 4 engines. As you approach a key point (lets say 25Km) you throttle up on the outboard 4 engines. You'd rely upon that method to reduce the Mach number before then throttling up on the inner engines. All of this technique to minimise cosine losses. Its just possible though that so long as you mount the engines high enough and far enough from the center of the vehicle, that there's no problem and you can do what you please.

Now for the wider picture. I wanted to keep as much mass as possible in high Mars orbit. But doing so meant for various reasons having a staging point in low Mars orbit. The reason I originally started with the idea of 2 transit vehicles is that one such vehicle could then make the journey to low Mars orbit and back. So then you'd offload mass to the other vehicle and leave it in high orbit and then aerobrake gradually into low orbit. Still, the final ascent back to high orbit did consume some fuel and presented other difficulties. What would be good is a Mars orbital "taxi". A vehicle capable of transferring crew between orbits but otherwise fairly minimal. I also kept running into the issue of how to provide fuel in low Mars orbit for the lander. And remember that's still an issue even for a conventional lander. And of course in any scenario there's still the question of ascent fuel. So I needed a "tanker" in low Mars orbit.

The key to the solution was to make that "tanker" one and the same as a transit vehicle, and to send it on a prior unmanned, low energy trajectory. Which made the lander even more interesting. If I could have a lander that would act as the taxi then things start coming together. But again, that meant having a low mass lander. And this is where the bigger picture comes into play. A conventional lander might use less fuel going down, but its going to be heavier, and thus use more fuel as a inter-orbit taxi.

More than this, there is also the issue that the transit vehicles need a propulsion system, and as it happens the lander happens to be the right size. Now, one option is to never return the lander to Earth space. In that case it would go on to land the ascent fuel and that's its final purpose (apart from a backup and a source of spares). But that would mean more engines for the transit vehicles. More importantly I felt that having the lander there attached to the transit vehicles on the return journey adds more resilience. The lander would not in itself have a long term life support system, but it would be very useful if for instance you suffered a fuel leak on the return journey. Then you could transfer fuel back to the lander's tanks, abandon the transit vehicles and use the lander to do the propulsive capture back into Earth orbit.

And then of course, its a useful craft for visiting the Mars moons.

A very handy vehicle indeed, and only possible if its as light weight as possible. Now, I can't make that happen if it has a heat shield and is designed to cope with the stresses of a standard re-entry.

JoshNH4H · 2013-06-21 23:50:26

Russel-

The important point that I think you forget is that there is more to be landed than just the crew. There will be a lot of other equipment that would need to be landed, and the mass multipliers for choosing not to use aerobraking would be truly horrific. Most of the equipment sent to the surface will be important mission technology; it wouldn't be sent otherwise. Loss of a hab could prove equally fatal as hitting the ground at 5 km/s.

I reject your estimate of 50% loss of Hydrogen. Given refrigerators and well-designed equipment, I would expect it to be closer to 10-20%, tops. source. Note that additional insulation, as well as the use of a reflective sunshield, could reduce boiloff to levels below those estimated by Dunnspace. By cooling the Hydrogen below its freezing point one can improve the density further, and potentially eliminate boiloff entirely. Do note that one would need the same size tank as for liquid Hydrogen because I would expect most if not all of it to melt.

Regarding loss through machinery: It all depends if you design for Hydrogen being present or not. Once on Mars, boiloff doesn't matter because a properly insulated tank will result in boiloff at a rate slower than the Hydrogen will be used for fuel production. Leaks and such should account for less than 1% of product mass. Once synthesized, meth-lox can be refrigerated fairly easily. In fact, if you design it right you can use the same cooling system for the Methlox as for the Hydrogen.

Regarding the size of the ISPP plant, Zubrin actually built one that would be capable of producing 20 times its mass in methlox fuel. This was done at a small scale and on a low budget ($50,000). I would bet a huge mass reduction would be possible with some development, perhaps fivefold. That would be leveraging of 100:1. Power requirements are relatively modest, too. It's also worth noting that this machinery would be fairly easily repairable by the crew were anything to break.

I would like to reiterate: Safe Martian aerocapture and aeroentry is not optional in the long term. It's something that needs to be developed, and as GW has said it's not that difficult to do. A one-use descent vehicle and a one-use ascent vehicle (Land it the orbital window before so that you have a chance to send another with the crew if necessary, in case of any failures). One-use because rockets are complex and prone to failure. I'd like to point out that there has never been a truly reusable rocket on Earth; The reason is that it's really difficult, requires a lot of development, and is expensive to develop. If you've ever been to Cape Kennedy, they have an RS-24 (Space Shuttle Main Engine) on display. This engine is "refurbishable" in that one doesn't dispose it after use. It looks like it has a million pieces. They spent a whole lot of money developing and operating it, and they still couldn't be sure that it would work for more than one launch. What about a Mars mission do you think will change that difficulty?

wrt L2: It is a benefit, if you have electric engines with high enough thrust. But I don't know just how much, and it adds another operation for unknown benefit. It does sure reduce the delta-V a lot. Why not do L1, out of curiosity? According to wikipedia it has the same delta-V as L2 to Mars. I'd like to point out that the delta V from LEO to L1 with low thrust is 7 km/s, instead of 3.77 km/s with high thrust. This will reduce the usefulness of electric propulsion methods, though not eliminate them entirely. For reference, the mass multiplier for H2/LOX from LEO to Mars transfer is going to be about 2.6-2.7 including rocket structure. From LEO to L2 using electric propulsion (Isp=3000 s), assuming a power source with a specific power of 500 W/kg, all inclusive. Ad Astra is very closed-mouth about the specific power or specific thrust of their device. I will estimate that it will consume 250 W/kg. Ad Astra says that their engine will produce 5.7 N/200 kW supplied to the RF generator. Because they are generally evasive, duplicitous, and nontrustworthy in their published figures, I will assume that means 250 kW input to the engines as a whole, giving a thrust-to-power ratio of .000023 N/w.

Assuming that you want your trip from LEO to L2 to take 6 months of continuous engine firing, you need .00045 N/kg of ship mass. This means that for every kg of wet mass, you need 20 W of power. This is 120 g of panel and engine per kilogram of wet mass. The mass ratio will be 1.3, so that's another 230 g/kg of fuel. Allow 50 g/kg of other structure (e.g., 5% of wet mass) and you end up with a mass multiplier of 2. It's a savings, but it's not huge. Engine mass is inversely proportional to travel time; The minimum travel time, when you have nothing but engine, fuel, and structure, is a bit over a month. Keep in mind that one must then launch to Mars, requiring an additional .74 km/s, for a second stage mass-multiplier of 2.5 if you use H2/LOX to boost. However, given that you are skeptical about the ability of Hydrogen to keep for six months (I'll be honest, I'm enjoying this part), using Methlox instead results in a mass multiplier of 2.6. The conclusion here is that, based on my calculations, Pure chemical is equivalent to stopping at L2. And yes, I did inflict a structural mass penalty for chemical as well--Not 5% of the wet mass but 12%.

As an aside, please note that this calculation suggests that the 39 days to Mars from LEO proposal by Ad Astra is simply ridiculous, having a total delta-V of 8 km/s and assuming Aerocapture at Mars. The mass multiplier for a 39 day trip would be 7, assuming additional structural masses to be zero. Assuming 5% of the total mass is other structure, the mass multiplier is 10. Most of this will be heavy engines and panels, which are a significant cost inandof themselves.

In this post I have assumed that firing time is the same as transfer time. While this is a good enough assumption in Earth-Moon space, for interplanetary distances it's not. Covering the 80 million km between Earth and Mars in 39 days would require a constant acceleration of .028 m/s^2, which would require 7.32 kg of engines and panels per kilo of wet mass. The total delta-V would be 94.3 km/s, for a mass ratio of 23.2. That is to say, for every kilogram of rocket, fuel, and structure, you can carry about -8 kg (Yes, that is negative eight kilograms) of payload. Given that we can't build our crew cabin out of exotic matter, I think this really demonstrates the duplicitiveness of Ad Astra's advertising. Even if the engines have no mass at all this mission is impossible.

Last edited by JoshNH4H (2013-06-22 00:12:18)

Russel · 2013-06-22 01:41:14

JoshNH4H wrote:

Russel-
The important point that I think you forget is that there is more to be landed than just the crew. There will be a lot of other equipment that would need to be landed, and the mass multipliers for choosing not to use aerobraking would be truly horrific. Most of the equipment sent to the surface will be important mission technology; it wouldn't be sent otherwise. Loss of a hab could prove equally fatal as hitting the ground at 5 km/s.
I reject your estimate of 50% loss of Hydrogen. Given refrigerators and well-designed equipment, I would expect it to be closer to 10-20%, tops. source. Note that additional insulation, as well as the use of a reflective sunshield, could reduce boiloff to levels below those estimated by Dunnspace. By cooling the Hydrogen below its freezing point one can improve the density further, and potentially eliminate boiloff entirely. Do note that one would need the same size tank as for liquid Hydrogen because I would expect most if not all of it to melt.
Regarding loss through machinery: It all depends if you design for Hydrogen being present or not. Once on Mars, boiloff doesn't matter because a properly insulated tank will result in boiloff at a rate slower than the Hydrogen will be used for fuel production. Leaks and such should account for less than 1% of product mass. Once synthesized, meth-lox can be refrigerated fairly easily. In fact, if you design it right you can use the same cooling system for the Methlox as for the Hydrogen.
Regarding the size of the ISPP plant, Zubrin actually built one that would be capable of producing 20 times its mass in methlox fuel. This was done at a small scale and on a low budget ($50,000). I would bet a huge mass reduction would be possible with some development, perhaps fivefold. That would be leveraging of 100:1. Power requirements are relatively modest, too. It's also worth noting that this machinery would be fairly easily repairable by the crew were anything to break.
I would like to reiterate: Safe Martian aerocapture and aeroentry is not optional in the long term. It's something that needs to be developed, and as GW has said it's not that difficult to do. A one-use descent vehicle and a one-use ascent vehicle (Land it the orbital window before so that you have a chance to send another with the crew if necessary, in case of any failures). One-use because rockets are complex and prone to failure. I'd like to point out that there has never been a truly reusable rocket on Earth; The reason is that it's really difficult, requires a lot of development, and is expensive to develop. If you've ever been to Cape Kennedy, they have an RS-24 (Space Shuttle Main Engine) on display. This engine is "refurbishable" in that one doesn't dispose it after use. It looks like it has a million pieces. They spent a whole lot of money developing and operating it, and they still couldn't be sure that it would work for more than one launch. What about a Mars mission do you think will change that difficulty?
wrt L2: It is a benefit, if you have electric engines with high enough thrust. But I don't know just how much, and it adds another operation for unknown benefit. It does sure reduce the delta-V a lot. Why not do L1, out of curiosity? According to wikipedia it has the same delta-V as L2 to Mars. I'd like to point out that the delta V from LEO to L1 with low thrust is 7 km/s, instead of 3.77 km/s with high thrust. This will reduce the usefulness of electric propulsion methods, though not eliminate them entirely. For reference, the mass multiplier for H2/LOX from LEO to Mars transfer is going to be about 2.6-2.7 including rocket structure. From LEO to L2 using electric propulsion (Isp=3000 s), assuming a power source with a specific power of 500 W/kg, all inclusive. Ad Astra is very closed-mouth about the specific power or specific thrust of their device. I will estimate that it will consume 250 W/kg. Ad Astra says that their engine will produce 5.7 N/200 kW supplied to the RF generator. Because they are generally evasive, duplicitous, and nontrustworthy in their published figures, I will assume that means 250 kW input to the engines as a whole, giving a thrust-to-power ratio of .000023 N/w.
Assuming that you want your trip from LEO to L2 to take 6 months of continuous engine firing, you need .00045 N/kg of ship mass. This means that for every kg of wet mass, you need 20 W of power. This is 120 g of panel and engine per kilogram of wet mass. The mass ratio will be 1.3, so that's another 230 g/kg of fuel. Allow 50 g/kg of other structure (e.g., 5% of wet mass) and you end up with a mass multiplier of 2. It's a savings, but it's not huge. Engine mass is inversely proportional to travel time; The minimum travel time, when you have nothing but engine, fuel, and structure, is a bit over a month. Keep in mind that one must then launch to Mars, requiring an additional .74 km/s, for a second stage mass-multiplier of 2.5 if you use H2/LOX to boost. However, given that you are skeptical about the ability of Hydrogen to keep for six months (I'll be honest, I'm enjoying this part), using Methlox instead results in a mass multiplier of 2.6. The conclusion here is that, based on my calculations, Pure chemical is equivalent to stopping at L2. And yes, I did inflict a structural mass penalty for chemical as well--Not 5% of the wet mass but 12%.
As an aside, please note that this calculation suggests that the 39 days to Mars from LEO proposal by Ad Astra is simply ridiculous, having a total delta-V of 8 km/s and assuming Aerocapture at Mars. The mass multiplier for a 39 day trip would be 7, assuming additional structural masses to be zero. Assuming 5% of the total mass is other structure, the mass multiplier is 10. Most of this will be heavy engines and panels, which are a significant cost inandof themselves.
In this post I have assumed that firing time is the same as transfer time. While this is a good enough assumption in Earth-Moon space, for interplanetary distances it's not. Covering the 80 million km between Earth and Mars in 39 days would require a constant acceleration of .028 m/s^2, which would require 7.32 kg of engines and panels per kilo of wet mass. The total delta-V would be 94.3 km/s, for a mass ratio of 23.2. That is to say, for every kilogram of rocket, fuel, and structure, you can carry about -8 kg (Yes, that is negative eight kilograms) of payload. Given that we can't build our crew cabin out of exotic matter, I think this really demonstrates the duplicitiveness of Ad Astra's advertising. Even if the engines have no mass at all this mission is impossible.

I've not forgotten that there is more to be landed than just crew. Indeed I've acknowledged that. But I've also treated it as out of scope. Why have I done that?

Because everything else that must be landed on Mars is on a one way trip. And conventional, heat shielded entry vehicles are almost certainly the way to land those masses. And I've deliberately left that part of the enterprise out of scope because essentially it costs what it costs. In other words, landing the crew in a mostly propulsive lander does not change the cost of landing other materials. It simply changes the cost of landing the crew. Hence I can logically separate the two.

It would be unfair to describe the mass multipliers for a mostly propulsive landing as horrific. In simple terms, the mass multiplier is 3, versus 2 for a conventional landing. The question is not the mass multiplier, its the mass being multiplied. As I've made clear, in rough terms, to get a crew down to the surface of Mars and back again, it costs around 35 tonnes per crew in Mars orbit. To do that conventionally the cost is around two thirds that. That 10 to 15 tonnes extra, delivered to low Mars orbit is a real cost, and I acknowledge that. But mass isn't the only thing that matters. Other metrics such as flexibility and safety matter more. Most missions that rely upon conventional landers have other points of failure. Its those points of failure I'm trying to address. So this isn't just about the landing method per se, its about having the right vehicles. And its about those vehicles performing multiple roles, and thus allowing for simpler, but more focused development.

As far as loss of hab goes. Yes, it doesn't do any good to your mission timeline when things get cratered. But its not loss of crew if all that mass is landed, and verified before a crew leave Earth. Most people agree on this point anyway so I didn't bother to mention it. I'll repeat the point though that in landing crew along with a large object, such as a hab, you actually increase the risk to the crew or the cost, because in the one case you're dealing with higher velocities and lower margins (and quite probably lower landing sites as a consequence) and in the other case you're giving a huge entry vehicle more margin because there is crew on board, but obviously that costs.

Like it or not, the sanest method of getting people to Mars is on purpose built vehicles. Ones with low mass and where you can then afford to suffer other mass multipliers because ultimately the last few Km of descent does matter, a lot.

In factoring a 50% loss of hydrogen, I'm not saying you're losing half your hydrogen, I'm saying you're losing a third for what you started with. Just to make that clear. Secondly its not just boil off losses. Hydrogen has a habit of permeating most materials. Even in a liquid state. And because you've got complex shapes inside a ISPP including manifolds, heat exchangers and so on its nearly impossible to avoid some loss. The only way to really reduce loss with hydrogen is to quickly form it into water, and that requires previously generating a commensurate volume of oxygen. Now, that's feasible. But, then there is always a cost.

Now I won't quibble with you about the precise percentage of hydrogen loss. In the scenario above you need theoretically 722Kg of hydrogen. I accounted for 50% over and above that for both boil off and other losses. That made it 1083Kg. If the loss accounted for is only 20% then you need to import 866.4Kg. The net saving is 216.6Kg. Its called diminishing returns. As I said, given the risks you'd normally adopt a more conservative margin for hydrogen than you would for methane. Nevertheless I can adopt a margin of 20% and it changes nothing to the argument. The cost of converting hydrogen to methane ultimately boils down to more mass having to be landed. At the risk of repeating myself, the ISPP plant needed to be built will be more massive (over one that purely produces oxygen) and the power source mass will also scale. Now, I'd like to think that such an ISPP plant would service 2 or 3 missions. But its a risk to assume it would last longer. And the main reason I think hydrogen isn't as good a mass-save as you'd assume is that the vehicle that lands the hydrogen has to be larger and heavier. And that's a mass multiplier that gets worse for a conventional lander because a conventional lander has higher forces and thus that extra tankage and support structure for the hydrogen scales.

And again, the wider picture is that even with some optimistic assumptions, landing 866Kg of Hydrogen versus 3000Kg of methane translates to a mass saving in the order of 10 tonnes back in low Earth orbit. Now, it must be said that any mass saving is, on the face of it, a good one. But compared to the 200+ tonnes of mass launched to low Earth orbit that's a difficult judgement call as to whether you want to spend the extra $20M or so on launch costs, or you want to face the development costs, and risks of going directly to importing hydrogen on your first mission.

Again, I'll state that hydrogen does make good sense if its obtained from the planet itself. That's a goal I'm fairly confident will happen. But as I said also, I think that the process of scraping Martian dirt and processing it, at scale, is probably going to require some human supervision. Making 100% propellant production a goal for the 3rd mission. Now, as and when that happens, my architecture evolves accordingly. We then don't need to land any fuel, and as a consequence we only need one lander landing. And as confidence in the technology builds it might be possible to reuse the lander for more missions. Again, that's another reason why I feel its a good idea to take the lander back into Earth space with you because it gives you an opportunity to check out and refurbish.

Let me make it clear. I don't think the benefit of importing hydrogen (over the 4 times heavier methane) vanishes entirely because of the problems I've noted (particularly as a consequence of scaling the lander that lands the hydrogen). But practicalities do reduce the benefit obtained. And in the end, you have to maintain a sense of perspective and make judgement calls where a few tens of millions in launch costs are not worth certain added complexities and risks.

By the way, I'm actually in favor of testing hydrogen to methane synthesis even at an earlier robotic stage. I hope that clears things up.

As far as safe Martian aerocapture goes. That would be wonderful. The hardest decision for me is choosing not to use aerocapture on either end. Were I to be convinced of the long term stability of light weight ceramic materials as GW is talking about, I'd actually be tempted to do aerocapture into a high Earth orbit first. And the reason is that on return to Earth you can put the crew into the lander and it has the capability to both capture and put the crew into low Earth orbit, and presumably dock with the space station. If the aerocapture for the main vehicle doesn't work, well, you're going to have to build a new one. But if aerocapture does work then you can extend the same concept to Mars. This time you give the lander more fuel and its capable of doing both propulsive capture and docking in low Mars orbit (see the beauty of having some form of habitat there?). Those two procedures knock about 40% off the original mass. So its not to be sneezed at. But when you look at the numbers, the whole mission is perfectly doable even without aerocapture. Its just going to be cheaper with.

However you ran aerocapture and aeroentry together and that's a bit unfortunate. Aerocapture has the potential to save nearly a hundred tonnes IMLEO. Aeroentry - at least for the crew - is a second order saving.

I've done a lot of thinking about every single technique that can be used to trim mass. Before coming back to realise that mass simply isn't everything.

Now the reason why there has never been a truly reusable rocket on Earth is that Earth has a much larger gravity well. Its simply a lot more difficult to do. 9Km/s versus 4Km/s. That's the long and short of it. The reality is that a mostly propulsive lander for Mars requires a bit less fuel to land than it does to ascend. You can't do that on Earth of course. The mass ratios would kill you. But on Mars the mass ratio is about 3.6 going up and about 3 going down. Well within the realms of well understood engineering.

I'm not sure why we're talking about Ad-Astra. I think their engine is way cool. I also think it would require a legendary power source to make practical as trans-Mars and trans-Earth vehicle. More to the point though when it comes to dealing with Earth's gravity well you don't need a fancy new engine. A NEXT ion drive (or a bank of them) would do the job satisfactorily. I'm not advocating doing that. Indeed I'm happy with the 200+ tonnes IMLEO that would be required to do this with conventional boosters and a fair chunk of that obviously would be ditched before the whole project left the Earth domain. I'd be even happier with that of course, if aerocapture could be relied upon. I've just not specified aerocapture as essential because when its all said and done, if we spend a few hundred million launching rocket fuel (per mission) then I can live with that until its clear how to safely improve on that.

There is a case for L2 and its with solar tugs between there and LEO. But the development cost for a MW class solar tug I believe is such that it might be ruled out as a possibility for the first mission. I could be wrong. These guys who want to go to the moon again, and establish L2 might win their argument. And in which case you'd piggy back off that infrastructure.

Let me also bring in something that is more speculative but reinforces my feeling that the design has some truth to it. When you build a space hab or transit vehicle if you wish, there's a minimum mass, partly because of consumables and water. And also because mass is a desirable thing when you're avoiding radiation. Now, in my architecture you're travelling home with a reserve of fuel that's intended for Earth orbital capture. So up until relatively late in the trip, the fuel you've kept in reserve creates extra shielding, at least for a portion of your combined vehicle. If you do without that you've got to deal with the trade off somewhere. That's one thing. The other is that because you're more than likely to carry several tonnes of water you've got yourself a good heat sink. It turns out that in aerocapture you can simply point the blunt end of your vehicle into the air stream and youu've got yourself a big kettle. Now you probably need a modest amount of thermal insulation (light weigh ceramic) and the rest is handled by the water. You might end up with several tonnes of water heated to 60C (anyone for a hot tub?) but essentially its a fairly non stressful way to do things. Again, provided you get the thermal protection right. Can such a thermal protection layer also compensate for and act as part of the radiation shield. I really don't know.

Last edited by Russel (2013-06-22 02:32:55)

GW Johnson · 2013-06-24 14:44:31

Myself, I am not cognizant of all the issues and numbers being discussed here. But I do know something about entry.

The peak heating rate per unit area (W/sq.cm) at the stagnation point is proportional to velocity-at-entry-interface cubed. Not squared, cubed. Entry from LEO here at Earth is pretty close to 11 km/sec. From Mars LMO, it is only about 3.7 km/sec. Even for direct interplanetary transfers, the entry interface velocity at Mars is only in the neighborhood of 5.6 km/sec. Entry at Mars is less demanding than from LEO by an order, to orders, of magnitude, in terms of peak heating.

Back in the mid-1950's a fellow named Julian Allen at NACA did the reentry stuff for warheads here on Earth. He and those working with him looked at heat sinks, ablatives, and (non-ablative) "refractories" for that mission. All would work, but heat sinking was by far the heaviest of the three. Back then, ablatives (in spite of their densities back then, such as silica and carbon phenolic, which sink very quickly in water) were a bit lighter weight than the refractory solutions, which depended upon high-density things like graphite and tungsten.

That's why ablatives were selected for Mercury. The refractories were to be tested on the X-20 Dyna-Soar, which got cancelled. This early Allen stuff didn't get published in the open until the mid-1960's, when it was finally declassified. You can find it on the web now, I did.

Since then, there's been two changes: (1) ablatives got lighter, and (2) refractories got lighter. Heat sinks never did, since the thermal capacitance per unit mass is density*heat capacity, which maximizes for high density and high heat capacity, and both of those things always correlate positively for all known substances.

Refractories got lighter with shuttle tile ceramics, but these could not be used near stagnation regions due to temperature limitations of no more than 2350 F. That's why the shuttle nose cap and aerosurface leading edges were carbon-carbon composite ablatives. These are very heavy, somewhat fragile, and require replacement every few flights. The tiles themselves were very, very fragile, and very, very, very labor-intensive to maintain.

More recently, ablatives also got lighter with the lower-density PICA and PICA-X materials. You can fly these a few times from LEO (presumably several to many times from LMO), and they are far less labor-intensive to install and maintain than refractory shuttle tile. I did finally find a published density for PICA-X at 0.27 g/cc. The panels on Dragon look to be about 1.5 to 2 inches thick, for 2-3 flights LEO.

The stuff I came up with is extremely experimental, but it handled as if it were commercial Styrofoam (somewhere near 0.03 g/cc). I used 0.2 inches of it in a ramjet combustor running at almost 4000 F, and it withstood the extremely-violent effects of rich blow-out combustion instability repeatedly, while serving for hours of accumulated burn in dozens of tests. I cannot go that hot for entry, I must avoid shrinkage cracks by staying under 2350 F, just like shuttle tile. but, that's where refractories can take us. I think they are the ultimate winner for entry heat shielding.

As for retro thrust, Dragon's Super Draco's are arranged to fire around the heat shield perimeter at 45 degree cant, just so that they do not have to "solve" the heat shield port problem. They have plenty of thrust, that's the landing system for manned Dragons, even here on Earth. it's far easier on Mars.

What I was talking about was firing a retro engine through an open hole in the heat shield. You can do that, if you seal the engine compartment to stop all gas throughflow through the hole. Don't do this on centerline, the plume won't know which way to flip-flop as it reverses, leading to induced destabilizing forces of sufficient magnitude to tumble the craft. Do this off-centerline with cant angles in the 5-15 degree range, for "perfect" plume stability at little sensible off-angle thrust loss.

The basic no-throughflow idea for holes in the heatshield already worked in 1969 with the Gemini-B test flight before MOL was cancelled.

GW

Mark Friedenbach · 2013-06-24 17:20:14

GW Johnson wrote:

The stuff I came up with is extremely experimental, but it handled as if it were commercial Styrofoam (somewhere near 0.03 g/cc). I used 0.2 inches of it in a ramjet combustor running at almost 4000 F, and it withstood the extremely-violent effects of rich blow-out combustion instability repeatedly, while serving for hours of accumulated burn in dozens of tests. I cannot go that hot for entry, I must avoid shrinkage cracks by staying under 2350 F, just like shuttle tile. but, that's where refractories can take us. I think they are the ultimate winner for entry heat shielding.

Have you patented this? You should think about doing so.

Russel · 2013-06-25 02:57:48

GW,

When I was considering a lander capable of a mostly propulsive landing, what I considered was the overall architecture. Hence the thread title.

The thought process goes roughly like this. A lander capable of landing mostly propulsively has a dry mass of around 3 tonnes, before you add the crew and basic life support. That's my starting point. Now, that vehicle does have tankage but the actual tankage can be less than a tonne. The rest of the dry mass is structure, attitude control, helium tanks and so on.

Now take the same lander and instead convert it to a lander that relies upon aerobraking and four things happen.

Firstly the dynamic pressure of entry increases and along with that structural loads and mass. So g forces go from a peak of 1.5g under the mostly propulsive case to closer to 4g.

Secondly relying mostly upon aerobraking means adding more drag area and thus despite less tankage, the vehicle actually needs a larger base.

Thirdly is the mass of the heat shielding itself. Now I believe (and I tend to share your optimism) that this is a second order problem. I think that light weight ablative materials capable of withstanding a few entries are going to cost a few hundred Kg. If your ceramic material is robust then halve that.

Fourth is that with hypersonic atmospheric reentry you need to add a back shell which adds further mass.

All of this adds up to a lander that probably tops 7 tonnes with crew. That includes life support, RCS fuel, but not terminal descent fuel. Terminal descent fuel is probably in the order of a further 400m/s so that adds a further tonne. More if you're not using methane and instead opt for lesser but less complex (in a small engine) fuels.

Now in order to use the same lander for ascent you need to deal with the issue that 4.5Km/s of delta V gives you a mass ratio of 3.6 which takes you to over 25 tonnes on take off. But now you've added extra tankage and structure. Which illustrates why its easy to go beyond 30 tonnes. Most of that mass is fuel and this has to be accounted for in the design of the ISPP system. And it has a knock on effect in terms of the mass of fuel that needs to be imported. This is true with hydrogen or methane.

You can of course eject the heat shield and back shell during landing. But what you end up with is a vehicle which as an ascent vehicle is somewhat fragile. And your abort to land options are more limited.

In essence you have to trade the fuel you'd use to land a lander mostly propulsively against the extra mass of fuel (and corresponding systems) that need to be placed on the surface. All of which have their own mass multipliers.

Now, viewed in isolation from the rest of the architecture the lander problem is not a huge deal. You build an ascent vehicle and wrap it in a conventional lander and land it on the surface prior to your crewed mission. I might add even in this case, even without the landing wrapper, the ascent vehicle is necessarily heavier than one landed propulsively. Then you refuel the ascent vehicle on the surface. Then you send the crew in a conventional expendable lander.

Its such an obvious process that it tends to be take for granted in many Mars architectures.

The problem starts getting interesting though when you consider what happens to your crew when they get into low Mars orbit at the end of a mission. Obviously there has to be some form of habitat there waiting for them. And that's where you start running into the philosophical differences between direct and semi direct camps. What I came to realise about this problem is that there is a lot of mass to be saved from keeping as much mass as possible in a high (near escape) Mars orbit. So what I've proposed is an architecture where most of the mass you bring from Earth is parked in a high Mars orbit. There is also a functional habitat that arrives in low Mars orbit. Its the same vehicle but its simply not as heavily provisioned. Its mostly being used as a tanker. But it also serves the purpose of a safe haven in orbit.

Once you think in those terms it becomes apparent you need a taxi of sorts. And the lander I've proposed is that taxi. And minimalism again counts in terms of the amount of fuel needed to transfer between orbits. Or for that matter transfer to and land on Phobos or Deimos.

I would consider a lander that uses a hypersonic heat shield and then ejects that and the back shell during landing and then goes on to become an ascent vehicle (even if by doing so it is necessarily a heavier vehicle during ascent. As I said, this still raises the concern that as an ascent vehicle is has less margin in abort to land scenarios.

What ultimately steers my hand though is how safe the landing phase is. A conventional hypersonic aerobraking landing is by its nature one where you have to make trades between the size/mass of the vehicle and the drag forces you experience in the lower atmosphere (below 15Km) and thus the final altitude at which you go into propulsive terminal descent. You're also committing yourself to a much more complex landing sequence where in essence, more things can go wrong. With a mostly propulsive lander there is much more freedom to control where you want to be and when. Its the difference between controlled flight and being a brick with a parachute attached. And with a lander that maintains propulsion its a lot easier to land safely at high altitudes. Even Olympus if you wish.

Try doing that with a parachute

Now back to thermal protection systems. As you've correctly pointed out the heating is a cube law. Which means that the first 2Km/s of braking overcomes much of the problem. That's achievable with an initial braking burn - perhaps down to 80Km. What I've tried to do from there on is to limit the heating to about 0.5W/cm2. That's basically 3 suns of heating. I did that because setting that limit in the simulator and then controlling the amount of propulsion to stay within that target was easily doable. I can do less heating but it does fractionally add more fuel.

What I'd like to do is to rely more heavily upon aerobraking, but again without adding too much in the way of structural stresses. The key challenge for me is that this vehicle is all out in the open. Its not hiding behind a heat shield. So things that stick out leg landing feet are going to get hot. Again, given the rule of thumb that the temperature in Kelvins is roughly equivalent to velocity I feel fairly confident in limiting the peak relative air speed to about 800m/s.

However, I could go further. And that's where light weight thermal protection comes in. So far I've left it at that because what is interesting me more now is the overall architecture.

Oh and one other issue regarding thermal protection. In my days before being an engineer I spent a lot of time being an apprentice mechanic. Which meant I got to appreciate stuff that was simple and designed to be serviceable. And I got to hate Austin 1800s In terms of a lander, what I found is that wrapping the whole thing in a heat shield and and back shell tended to make a lot of stuff harder to service, or even discover if it was damaged. I like systems that can be quickly visually inspected. And I'm definitely aiming for a lander design where whole engines can be manually replaced - even in a space suit. So for the moment what appeals to me is thermal protection in blanket form that can be manually wrapped, unwrapped and repaired.

Now, when it comes to ceramic materials the real question is, can you do it in a modular clip on, clip off fashion. That would appeal.

GW Johnson · 2013-06-25 10:01:30

Hi Russell:

I think we're dancing around the extremes of a trade study here. It is quite possible to maximize aerodeceleration hypersonically at Mars, and then cope with a low altitude when you come out of hypersonics. In my studies, I just went to a retro-thrust powered landing, since for 0.7 km/s (local Mach 3) at 5 km altitude, even at low path angle you are a mere two minutes from impact.

In your study, retro thrust is used throughout the entry, which is going to reduce peak heating and total heat to absorb, however it is resisted. This will also increase the altitude at which one comes out of the hypersonics. That makes chutes and ballutes feasible, and therefore a thing you could consider. Although, I don't recommend a chute or ballute simultaneous with retro thrust, except for last-second touchdown of extremely heavy loads. It's too easy for flight control to be upset by wind gusts laterally.

The trade study would be when to start retro thrust during hypersonic entry. You say at the beginning of the hypersonics, I say at end of hypersonics, maybe there is an optimum in between! The idea here is to land the greatest tonnage of payload for the least tonnage of propellant. That would be true whether or not you plan to make return propellant while on the surface of Mars.

To my knowledge, the "when-during-hypersonics-do-we-start-retro-thrust?" trade study has never been run by anyone. People have shied away from concepts like that because of (what I consider to be relatively groundless) fears over firing through holes in heat shields and stability of retro plumes versus vehicle attitude control.

There is something very serious to consider about hypersonic flight conditions and vehicle configurations: extreme shock-impingement heating above about local Mach 5 or 6 speeds. This is a demonstrated risk from the old X-15 program. You must fly a very clean shape, not an assembly of nacelles, because each nacelle sheds a shock system that impacts the other nacelles, and the structures connecting them.

The X-15 flight with the scramjet test article replacing the ventral fin was the flight that reached Mach 6.67. The scram nacelle bow wave impinging upon the adjacent undersurface of the fuselage nearly cut the tail section off the bird in a matter of seconds, once about Mach 6 got exceeded. And don't forget that the skin was exotic refractory Inconel-X.

Josh:

There's rocket engines, and there's complicated rocket engines. Hydrogen-oxygen things tend to be quite complicated and therefore expendable one-shot items. Its only advantage is 450+ s Isp. Methane-oxygen might be complicated, or maybe not. It is 300-class Isp.

XCOR has the habit of making very reliable, very long-life reusable engines without much of an Isp penalty. They've certainly done it with kerosene-oxygen (300-class Isp), and I know they're working on methane-oxygen. Their engines don't look much like those NASA has always procured heretofore. They look like something you could safely put in an airplane. Some of them use piston pumps, not turbopumps.

The easiest, simplest, most reliable engines of moderate thrust that I know of are the hypergolic-ignition units that use hydrazine-nitrogen tetroxide (also 300-class Isp). Could be any of the hydrazines, but MMH seems favored. These can range from 1-pound thrusters all the way up to the huge OMS units on the shuttle. Pretty reliable technology, we've been using it since the late 50's. Something like it powered the old Titan-II’s that launched Gemini.

Unfortunately, MMH-NTO doesn't yet sound like something we might make in-situ on Mars. But it is quite dense and easily shipped, and extremely easily-storable for years at a time. I dunno, but MMH-NTO sounds like an ideal Mars lander/ascent vehicle fuel. Until we figure out how to make it there, we'll have to bite the bullet and ship it from Earth.

For the transit vehicle, cryo boiloff problems can be solved with only a modest loss for using LH2-LOX (450-class Isp) or LH2-nuclear (750-1100+ class Isp as solid core, depending upon details). But I doubt cryo tankage like that could also be successfully adapted to resist the heat and air loads of aerocapture at Mars. Too many conflicting requirements for the designs.

MMH-NTO (300-class Isp) is a serious mass ratio penalty for transit, but could easily be protected for aerocapture at Mars. There's a lot off tradeoffs there. Including LH2-LOX to Mars, aerocapture, and MMH-NTO for the return. Haven't seen anybody look at that yet.

I'm not yet sure at all how the ISRU propellant restrictions feed back into the lander design, much less aerocapture vs transit vehicle propellant restrictions. My point is that there are very serious restrictions on your mission design imposed by these choices. Your choices are not free.

If I was doing it without aerocapture or ISRU, I'd used LH2-nuclear for the transit vehicle, and MMH-NTO for the landers. I'd stage the landers out of LMO and refuel them there for reuse with propellant I brought along. I build this from 20-50 ton modules launched to LEO, with rockets we already have by next year, and docked there. I’d spend my money on the mission that way, instead of developing a giant rocket that can launch 100-ton modules, but that nobody else has any use for, yet.

There’s a brute-force baseline, with the best stuff we have. Now, try lowering the launched mass by aerocapture and ISRU, while staying within the propellant-choice restrictions they impose on your designs. That’s the way to reduce mission costs.

GW

GW Johnson · 2013-06-25 10:05:47

Oops, I forgot. My oddball ceramic heat shield is retained (two ways) upon backplates that are part of the vehicle shell assembly. The idea is to bolt-on directly to the interior structure, big extended panels of shell plating that have integral heat protection. They'd come off the same way for interior access.

GW

Russel · 2013-06-25 11:10:12

GW Johnson wrote:

Oops, I forgot. My oddball ceramic heat shield is retained (two ways) upon backplates that are part of the vehicle shell assembly. The idea is to bolt-on directly to the interior structure, big extended panels of shell plating that have integral heat protection. They'd come off the same way for interior access.
GW

Ok well in response to that one what I'm thinking here is that for my purposes most of the surfaces that need some form of heat protection are actually cryogenic tanks. The surface is actually kept cold by the contents so the purpose of the thermal protection is really to reduce the rate of boil off. Hence small shingles even with gaps is quite sufficient.

Russel · 2013-06-25 12:06:35

GW Johnson wrote:

Hi Russell:
I think we're dancing around the extremes of a trade study here. It is quite possible to maximize aerodeceleration hypersonically at Mars, and then cope with a low altitude when you come out of hypersonics. In my studies, I just went to a retro-thrust powered landing, since for 0.7 km/s (local Mach 3) at 5 km altitude, even at low path angle you are a mere two minutes from impact.
In your study, retro thrust is used throughout the entry, which is going to reduce peak heating and total heat to absorb, however it is resisted. This will also increase the altitude at which one comes out of the hypersonics. That makes chutes and ballutes feasible, and therefore a thing you could consider. Although, I don't recommend a chute or ballute simultaneous with retro thrust, except for last-second touchdown of extremely heavy loads. It's too easy for flight control to be upset by wind gusts laterally.
The trade study would be when to start retro thrust during hypersonic entry. You say at the beginning of the hypersonics, I say at end of hypersonics, maybe there is an optimum in between! The idea here is to land the greatest tonnage of payload for the least tonnage of propellant. That would be true whether or not you plan to make return propellant while on the surface of Mars.
To my knowledge, the "when-during-hypersonics-do-we-start-retro-thrust?" trade study has never been run by anyone. People have shied away from concepts like that because of (what I consider to be relatively groundless) fears over firing through holes in heat shields and stability of retro plumes versus vehicle attitude control.
There is something very serious to consider about hypersonic flight conditions and vehicle configurations: extreme shock-impingement heating above about local Mach 5 or 6 speeds. This is a demonstrated risk from the old X-15 program. You must fly a very clean shape, not an assembly of nacelles, because each nacelle sheds a shock system that impacts the other nacelles, and the structures connecting them.
The X-15 flight with the scramjet test article replacing the ventral fin was the flight that reached Mach 6.67. The scram nacelle bow wave impinging upon the adjacent undersurface of the fuselage nearly cut the tail section off the bird in a matter of seconds, once about Mach 6 got exceeded. And don't forget that the skin was exotic refractory Inconel-X.
Josh:
There's rocket engines, and there's complicated rocket engines. Hydrogen-oxygen things tend to be quite complicated and therefore expendable one-shot items. Its only advantage is 450+ s Isp. Methane-oxygen might be complicated, or maybe not. It is 300-class Isp.
XCOR has the habit of making very reliable, very long-life reusable engines without much of an Isp penalty. They've certainly done it with kerosene-oxygen (300-class Isp), and I know they're working on methane-oxygen. Their engines don't look much like those NASA has always procured heretofore. They look like something you could safely put in an airplane. Some of them use piston pumps, not turbopumps.
The easiest, simplest, most reliable engines of moderate thrust that I know of are the hypergolic-ignition units that use hydrazine-nitrogen tetroxide (also 300-class Isp). Could be any of the hydrazines, but MMH seems favored. These can range from 1-pound thrusters all the way up to the huge OMS units on the shuttle. Pretty reliable technology, we've been using it since the late 50's. Something like it powered the old Titan-II’s that launched Gemini.
Unfortunately, MMH-NTO doesn't yet sound like something we might make in-situ on Mars. But it is quite dense and easily shipped, and extremely easily-storable for years at a time. I dunno, but MMH-NTO sounds like an ideal Mars lander/ascent vehicle fuel. Until we figure out how to make it there, we'll have to bite the bullet and ship it from Earth.
For the transit vehicle, cryo boiloff problems can be solved with only a modest loss for using LH2-LOX (450-class Isp) or LH2-nuclear (750-1100+ class Isp as solid core, depending upon details). But I doubt cryo tankage like that could also be successfully adapted to resist the heat and air loads of aerocapture at Mars. Too many conflicting requirements for the designs.
MMH-NTO (300-class Isp) is a serious mass ratio penalty for transit, but could easily be protected for aerocapture at Mars. There's a lot off tradeoffs there. Including LH2-LOX to Mars, aerocapture, and MMH-NTO for the return. Haven't seen anybody look at that yet.
I'm not yet sure at all how the ISRU propellant restrictions feed back into the lander design, much less aerocapture vs transit vehicle propellant restrictions. My point is that there are very serious restrictions on your mission design imposed by these choices. Your choices are not free.
If I was doing it without aerocapture or ISRU, I'd used LH2-nuclear for the transit vehicle, and MMH-NTO for the landers. I'd stage the landers out of LMO and refuel them there for reuse with propellant I brought along. I build this from 20-50 ton modules launched to LEO, with rockets we already have by next year, and docked there. I’d spend my money on the mission that way, instead of developing a giant rocket that can launch 100-ton modules, but that nobody else has any use for, yet.
There’s a brute-force baseline, with the best stuff we have. Now, try lowering the launched mass by aerocapture and ISRU, while staying within the propellant-choice restrictions they impose on your designs. That’s the way to reduce mission costs.
GW

Sorry for the long quote but there are points here and there I want to pick up on.

Getting to mach 3 at a reasonable altitude and relying only on aerodynamic braking means having enough effective drag area. Which makes for an awkward trade between the diameter of the heat shield and overall mass. Using propulsion to some extent side steps that issue. In any event a vehicle that is designed without a heat shield and with SRP in mind probably will place the engines well outside the main body of the vehicle - effectively on booms. And that arrangement tends to add to drag. Another reason for placing the engines above the center of mass is that then adds to stability.

As far as drag devices go, what I have in mind that might be useful is a trailing drag device deployed not long after the initial braking burn. So starting at around 80Km into a relative air speed of at most 600m/s. And then letting it go at around 20Km whilst throttling up. During the time the drag device is deployed there is either no thrust, or low thrust. Prudence may call for low thrust purely to lessen the risk of restarting into a denser air stream.

The trades as to the exact thrust profile are beyond my resources. I'd like to see it done but I think before that's done we need to gain a better understanding of how real thrusters behave. That's why I hope they use Dragon to deliberately test this. I suspect that what we will find is an envelope. A graph plotting on one axis velocity and on the other density. And within that there is a boundary inside which SRP can be used. Or even a set of boundaries defined by the level of thrust. And it would not surprise me at all if allowable velocity rises as density decreases, and vice versa. So for me the trickiest trade is probably about knowing how soon and how fast to throttle up. Even with aerodynamic braking the same problem will apply. Do you attempt SRP at local Mach 3 or do you attempt it at a higher Mach number but with less density air. Or more likely do you ramp up your thrust according to some criteria.

Fortunately I don't anticipate going anywhere near Mach 5 once the initial braking burn has been done.

Regarding engines, its my understanding that a well designed methane/LOX engine will develop an Isp of 350 in a vacuum. Which is another point. The density of air on Mars is near enough of a vacuum that this won't affect the Isp of a rocket engine designed for vacuum. But the same engine approaching at a relative velocity of Mach 3 will probably experience some reduction in Isp.

As far as reliability goes, I don't see an issue in designing reliable methane/LOX engines. Ones that can spend long periods idle. Ones that can be restarted multiple times. Ones that endure long burns. Its a matter of good design and a lot of testing. As I pointed out earlier, the reliability I need for methane/LOX engines in my lander is the reliability needed for the methane/LOX engines in most other architectures. And I suspect that the hardest requirement is actually long periods of idle. Again, I'd very much want to be in the position of being able to visually inspect the engine before I flew on it. And having a lander that also becomes the propulsion unit for the flight home is all part of that assurance.

I am not sure what you mean by transit vehicle propellant restrictions.

As for a nuclear LH2 engine, where I'd like to see one tested within the context of my architecture is this. I propose to transfer to low Mars orbit a transit vehicle which is also a fuel depot and a lander (the one used in the previous mission but is now the lander that lands the next cargo of methane) All of this mass - around 45 tonnes if I recall correctly - can be sent on a low energy trajectory well ahead of the crew. If the booster responsible were nuclear then the modest power requirement would mean a compact core. Eventually if you could prove the technology and demonstrate its safety then a scaled core version could become the booster for the manned transfer. I would still rely upon stored methane as the return fuel. In this case the stack would look like this. A lander docked on top. Then a transit vehicle docked to a further transit vehicle and finally at the bottom connected by the same universal docking port the nuclear booster. The crew would initially travel in the upper transit vehicle which would put a lot of mass between them and the core. I think I recall that in NASA's reference design they wanted 10 tonnes of shielding for the manned version. Eep.

RGClark · 2013-06-27 01:05:46

Mark Friedenbach wrote:

GW Johnson wrote:
The stuff I came up with is extremely experimental, but it handled as if it were commercial Styrofoam (somewhere near 0.03 g/cc). I used 0.2 inches of it in a ramjet combustor running at almost 4000 F, and it withstood the extremely-violent effects of rich blow-out combustion instability repeatedly, while serving for hours of accumulated burn in dozens of tests. I cannot go that hot for entry, I must avoid shrinkage cracks by staying under 2350 F, just like shuttle tile. but, that's where refractories can take us. I think they are the ultimate winner for entry heat shielding.
Have you patented this? You should think about doing so.

Second that. That 0.03 g/cc density is extraordinary, assuming it requires similar thickness to say PICA-X.
GW, there was another application of this that might be patentable I'll discuss with you in an email.

Bob Clark

Last edited by RGClark (2013-06-27 01:11:45)

JoshNH4H · 2013-06-27 21:28:39

A couple of introductory comments, then a more detailed response:

First, regarding Ad Astra: I'm sure you see the relevance of the first part of that section of my reply, if not the importance. I went rather off topic for the second part, but while I had my methods laid out I figured I would address a claim by Ad Astra that I find infinitely frustrating due to its blatant falseness. In any case, thought it may not live up to the extremely hyperbolic claims made by its developers, VASIMR is more or less flight-ready at a small scale-- assuming the claims made by Ad Astra are true.

As an aside, based on the numbers given in this report NEXT's performance is actually worse than I assumed in the above model. To improve upon the numbers I gave would require better engine performance, either in the Isp range or specific power, as well as better specific power for the power generation system. Again, please note that the numbers I used may be unrealistically forgiving on the power system. That report used 6 kg/kW, which is 167 W/kg. This is 1/3 of my assumed value, and changes the full-system mass multiplier for your proposal to about 3.3, which is actually worse than chemical thrusters out of LEO. This is pretty relevant to the discussion because it eliminates the reasoning behind assembly in L2. I know that you don't consider this to be central but where numbers can lead to consensus I see every reason to pursue them.

Regarding the mass multipliers for Mars Entry, I reject your figure of 2 for aerobraking. Zubrin estimates that the mass of a heatshield will be 15-20% of the mass that it needs to land, and inflatable ones are by their nature quite light. Fuel mass should be pretty low for the mild deceleration delta-V required, whether the fuel be methane or Hydrogen. I'd say Methane may be preferable in this case, though I don't have much evidence in support. I'd estimate a mass multiplier in the region of 1.5.

The logical interpretation of "50% Hydrogen loss" is that you lose half of your hydrogen. You're describing a situation with 33% Hydrogen loss.

Now, delving more into prime content:

I just don't see any benefit from making the mission components reusable. It's a safety hazard, it increases the seed mass and the per-mission mass (if you want to reuse the rocket it's gonna be more massive, and you want to send fuel for it), and it requires additional development seeing as you're already developing the reentry technology.

This is the fundamental part of my argument. Stated succinctly: You've proven beyond a reasonable doubt that your mission is different from the reference mission, and you've even argued that the increase and costs and the decrease in safety might not be too large (At least not the first mission; I do still intend to address powered landing). However, you have by no means demonstrated that it is better.

Before I launch (no pun intended) into a discussion of powered landing, I was wondering what your take on the L2 stuff was. I think I demonstrated fairly conclusively that using electric propulsion to travel to L2 does not represent any mass savings over chemical rockets from LEO. Do you agree with my assessment or are you of the opinion that it is invalid for whatever reason?

Regarding reusable rockets for powered descent:

You say that you're considering the landing of crew separately from the landing of cargo. The problem is that these things are not totally separate. There is no reason why one needs to be safer than another. Without a landed powersource, the crew is dead, period. Now, you claim that it is not necessary to develop "man-rated" landing equipment because you can always send vital equipment at the next launch window. Let's say we need each piece of equipment to have a 1% total chance of not making it. This means that, given two chances, the aerobraking would have to have a reliability of 90%. Honestly, 90% isn't that much different in engineering terms from 99%, or 99.9%. The difference is that you put in more time and money to really verify your method in all situations, vs. skimping on proper testing and simulating. If you have a 90% reliable method, you have a method that works, and it's just a matter of making it work every time.

This is a perfect time to develop this vital technology; Why skimp? We're going to need high-reliability aerobraking eventually. It's an investment that will pay for itself many times over, and we might as well start as soon as possible.

Now, you claim that the primary reason why EtO (Earth to Orbit) rockets have not been made reusable is that the delta-V to Earth orbit (~9.3 km/s) is much higher than the delta-V to Mars orbit. While this is true, it's not the only or even primary reason. Even a two-stage rocket has about 4.7 km/s per stage, which isn't that big of a difference from what you're proposing. The reason why it's hard to make a reusable rocket is that multiple firings can damage the engine; the thermal cycling of launch and re-entry, followed by spending time on the cool surface of Mars (not to mention the cryogenics in the tanks) weakens the tanks and any weld connections that may be there. Most rocket engines wear out after the equivalent of more than a couple minutes of continuous firing. None of these issues are unresolvable, of course-- but the primary way to resolve them is to trade payload for rocket structure. This makes your mass multipliers significantly worse.

And again--Rockets are complicated machines. You can't use them over and over and over again and just trust them not to break. Especially not without a large ground team on standby to fix anything that may have gone wrong during a flight. That's how you get someone killed. That's how you get colonization initiatives ended.

So, I pose this question to you: Given the obvious and hard-to-fix safety issues, why shouldn't we just aerobrake?

Russel · 2013-06-29 03:42:49

JoshNH4H wrote:

A couple of introductory comments, then a more detailed response:
First, regarding Ad Astra: I'm sure you see the relevance of the first part of that section of my reply, if not the importance. I went rather off topic for the second part, but while I had my methods laid out I figured I would address a claim by Ad Astra that I find infinitely frustrating due to its blatant falseness. In any case, thought it may not live up to the extremely hyperbolic claims made by its developers, VASIMR is more or less flight-ready at a small scale-- assuming the claims made by Ad Astra are true.
As an aside, based on the numbers given in this report NEXT's performance is actually worse than I assumed in the above model. To improve upon the numbers I gave would require better engine performance, either in the Isp range or specific power, as well as better specific power for the power generation system. Again, please note that the numbers I used may be unrealistically forgiving on the power system. That report used 6 kg/kW, which is 167 W/kg. This is 1/3 of my assumed value, and changes the full-system mass multiplier for your proposal to about 3.3, which is actually worse than chemical thrusters out of LEO. This is pretty relevant to the discussion because it eliminates the reasoning behind assembly in L2. I know that you don't consider this to be central but where numbers can lead to consensus I see every reason to pursue them.
Regarding the mass multipliers for Mars Entry, I reject your figure of 2 for aerobraking. Zubrin estimates that the mass of a heatshield will be 15-20% of the mass that it needs to land, and inflatable ones are by their nature quite light. Fuel mass should be pretty low for the mild deceleration delta-V required, whether the fuel be methane or Hydrogen. I'd say Methane may be preferable in this case, though I don't have much evidence in support. I'd estimate a mass multiplier in the region of 1.5.
The logical interpretation of "50% Hydrogen loss" is that you lose half of your hydrogen. You're describing a situation with 33% Hydrogen loss.
Now, delving more into prime content:
I just don't see any benefit from making the mission components reusable. It's a safety hazard, it increases the seed mass and the per-mission mass (if you want to reuse the rocket it's gonna be more massive, and you want to send fuel for it), and it requires additional development seeing as you're already developing the reentry technology.
This is the fundamental part of my argument. Stated succinctly: You've proven beyond a reasonable doubt that your mission is different from the reference mission, and you've even argued that the increase and costs and the decrease in safety might not be too large (At least not the first mission; I do still intend to address powered landing). However, you have by no means demonstrated that it is better.
Before I launch (no pun intended) into a discussion of powered landing, I was wondering what your take on the L2 stuff was. I think I demonstrated fairly conclusively that using electric propulsion to travel to L2 does not represent any mass savings over chemical rockets from LEO. Do you agree with my assessment or are you of the opinion that it is invalid for whatever reason?
Regarding reusable rockets for powered descent:
You say that you're considering the landing of crew separately from the landing of cargo. The problem is that these things are not totally separate. There is no reason why one needs to be safer than another. Without a landed powersource, the crew is dead, period. Now, you claim that it is not necessary to develop "man-rated" landing equipment because you can always send vital equipment at the next launch window. Let's say we need each piece of equipment to have a 1% total chance of not making it. This means that, given two chances, the aerobraking would have to have a reliability of 90%. Honestly, 90% isn't that much different in engineering terms from 99%, or 99.9%. The difference is that you put in more time and money to really verify your method in all situations, vs. skimping on proper testing and simulating. If you have a 90% reliable method, you have a method that works, and it's just a matter of making it work every time.
This is a perfect time to develop this vital technology; Why skimp? We're going to need high-reliability aerobraking eventually. It's an investment that will pay for itself many times over, and we might as well start as soon as possible.
Now, you claim that the primary reason why EtO (Earth to Orbit) rockets have not been made reusable is that the delta-V to Earth orbit (~9.3 km/s) is much higher than the delta-V to Mars orbit. While this is true, it's not the only or even primary reason. Even a two-stage rocket has about 4.7 km/s per stage, which isn't that big of a difference from what you're proposing. The reason why it's hard to make a reusable rocket is that multiple firings can damage the engine; the thermal cycling of launch and re-entry, followed by spending time on the cool surface of Mars (not to mention the cryogenics in the tanks) weakens the tanks and any weld connections that may be there. Most rocket engines wear out after the equivalent of more than a couple minutes of continuous firing. None of these issues are unresolvable, of course-- but the primary way to resolve them is to trade payload for rocket structure. This makes your mass multipliers significantly worse.
And again--Rockets are complicated machines. You can't use them over and over and over again and just trust them not to break. Especially not without a large ground team on standby to fix anything that may have gone wrong during a flight. That's how you get someone killed. That's how you get colonization initiatives ended.
So, I pose this question to you: Given the obvious and hard-to-fix safety issues, why shouldn't we just aerobrake?

I think the point you've not fully understood is that I've always operated on the basis that anything critical to safety will be landed on Mars and verified as operational before the crew leave Earth. So the safety issue you imagine isn't there. It a very simple principle. An unmanned mission lands what you need on Mars. Crew follows. There's no such thing as crew landing with life critical equipment.

I'd put it to you that the shoe is on the other foot here. You land crew with cargo and if the cargo doesn't function, you've got a safety issue. Even if you have landed successfully. Safety alone dictates you take the alternate route and land anything critical to safety before the crew leave Earth.

Once you've understood this, its possible to decouple the quite separate issues of landing large masses, and landing crew. In doing so you can provide more safety for the crew than you can possibly do were you to land the crew on a large vehicle.

I stand by a mass multiplier of 2 to 1. That's a 50% landed mass to entry mass. No landing on Mars so far has come close to this. NASA thinks that to get 30+ tonnes to Mars you need 100 tonnes at entry. Zubrin is a lot more optimistic. I'm adopting the middle ground. I think that a mass multiplier of 2 to 1 is possible with large payloads.

However with smaller vehicles you'd be quite lucky to get to that 2 to 1 (landed versus entry mass) figure. I was being fair. Zubrin btw is primarily interested in large vehicles.

I tend to agree that an evolved heat shield may amount to 15-20% of the overall entry mass. However one factor I keep repeating that doesn't seem to be fully understood is that apart from the heat shield the entire vehicle needs to be heavier because its being exposed to much higher forces during aerobraking than would be the case for a mostly propulsive landing.

So for instance, if you wanted to land an expendable ascent vehicle using aerobraking (inside an expendable lander), the payload itself - that is the ascent vehicle - must be heavier than if it were landed mostly propulsively. Hence one of the advantages of the lander and ascent vehicle being one and the same vehicle is that you can make the ascent vehicle lighter than could possibly be the case if the ascent vehicle were landed conventionally. Lets be more concrete about this. A mostly propulsive landing can be engineered with a peak deceleration of 18m/s/s. (About 2 Earth g). A landing involving aerobraking involves upwards of 40m/s/s. The reason for this in simple terms is that you've burned off most of the velocity in the initial burn and the final burn is of comparable scale to a terminal descent from aerobraking - so around the 500m/s mark, give or take. From there to zero vertical velocity in 40 seconds without needing to exceed the 18m/s/s mark. Now with aerobraking you're a prisoner to the density profile of the atmosphere and you're having to burn through all that velocity much closer to the ground.

Now, adding more drag. A larger heat shield. Parachutes. Ballutes. Anything you can imagine. They can only slow you down faster in the denser layers of atmosphere. Which means the more margin you want to build in, the higher the stress on the vehicle structure. And that affects everything. Framing, tankage, even down to the little bolts that secure the plumbing.

There are of course ways out of this. Generating more drag higher up will give you lower stresses. But this means something huge, and light, and thus fragile. And probably something you will have to jettison in favor of a smaller drag device (going from ballute to parachute) at a relatively high altitude and as well as creating more things that can go wrong, will be pushing the limits of the second tier drag device. There is also lift that buys you more time but also buys you into more complexity and more pushing of the limits of materials. Even a trailing ballute device engineered to provide lift still has the problem that as the air becomes more dense it will generate too much drag and either self destruct or put too much load onto the vehicle.

Worse, all of this gets harder and harder to do, the larger your vehicle is. Buying margin for your crew is a lot easier and less expensive on a smaller vehicle. Taking a vehicle that masses 60 tonnes at entry and trying to give it the same margin as is possible with a smaller vehicle is, well, frankly bordering on show stopping. And there will always be the temptation to cut corners. You either end up with outrageous situations where you've reached final maneuvering (most of your vertical velocity is gone) with a Km to go, or you end up discovering that all your margin is gone and the only way to make up for it is to land people on lower terrain.

The initial burn of a mostly propulsive landing takes away most of the velocity and most of the difficult and thus safety critical materials issues. I'll say that again. Aerobraking is a difficult and dangerous process because you're stressing materials to their limits. And no I'm not particularly worried about ablatives there. I'm talking about the materials you're going to use for ballutes and the like. I don't mind those things if they're part of landing a non manned cargo. (After all, your crew isn't going to leave Earth until critical cargo has been landed safely and is functional, right?). That way the risk of very large drag devices isn't an issue of crew safety. And lets not forget also that landing large masses on Mars absolutely begs for these kinds of devices.

Another point about safety. When you land propulsively you're given back the ability to set the parameters. You've restored authority over the situation. You can land anywhere on Mars. You can choose the margin. It does cost fuel sure. But you're in a position where any other drag device you deploy merely adds to the overall level of comfort but you can still land safely even if it fails. Crew survival doesn't then absolutely depend on the correct functioning of such devices. With a small vehicle and an aerobraking landing that depends on drag devices you're taking a risk on those devices. With a much larger craft that has to land cargo and at the same time maintain margins for crew you're further pushing the materials. You're further increasing risk.

Final point on this issue. Much of uncertainty in an aerobraking landing is due to uncertainties in the atmosphere itself. You can to a certain extent compensate for that. That's what MSL was programmed to do by using a small amount of thrust to essentially fly its way out of these uncertainties. Nevertheless this problem exists because you're relying on air to slow you down. A mostly propulsive landing is affected less by the uncertainty in the atmospheric profile and also has more ability to correct for it. If you care about crew safety its little details like that that matter. And landing crew with a large vehicle makes this problem also much more difficult. And more difficult still if the large vehicle is relying even more heavily on cutting edge drag devices, and has more inertia and is thus harder to "fly".

Lets go to the issue of engine reliability. Reliable engines are not black magic. It simply requires competent engineering. The problem with most engines are they are deliberately designed down in mass knowing they only have to work once. There are wel l known examples of engines that can survive multiple restarts and long burns. The price to be paid of course is extra mass. But that's extra mass on top of a component that overall is not a large fraction of your overall mass budget. Put it this way, given the choice between an engine with a 100:1 thrust to weight ratio (relative to Earth gravity) that is designed for single use and an alternative engine with a 70:1 thrust to weight ratio (its heavier) that is ultra reliable. Which one are they going to specify, even if the mission only (nominally) requires one firing?

I'll also repeat something I said before. There are many Mars architectures that rely upon methane/LOX engines that are capable of multiple firings and long burns. We cannot avoid either designing such an engine or testing it.

Putting an Earth return vehicle on low Mars orbit, or even resorting to landing it on Mars does not obviate the need for reliable engines. Instead a sensible mission involves the capability of multiple firings if not for course corrections then for unforeseen maneuvers or contingencies. Engines deliberately designed down (and I do mean deliberately.. its a choice) to the point of being single use, should never be sent to Mars.

Having a minimal mass ascent vehicle means less propellant manufactured on Mars. Less propellant (even if its just oxygen) means less mas both for the ISPP plant and for the energy source. Some architectures seem to take ISPP too much for granted. I'm merely making the point that if your sole concern is how much mass you need to get into low Earth orbit, sooner or later you need to deal with the mass of the ISPP system and thus the mass of the ascent vehicle itself. And the lightest possible ascent vehicle is the one you've landed via mostly propulsive means, since its a gentler landing.

Even in a single stage to orbit Mars ascent vehicle you still have to incorporate features that are specific to landing. Otherwise you've no abort to land. And designing a lander and ascent vehicle together guarantees that the ascent vehicle is guaranteed to have abort to land capability. How many Mars architectures have been thought through this far?

I'm not asking my engines to be reused over and over again for multiple missions. They're being used once for a crewed landing. Once for a crewed ascent. During those two critical times, there is redundancy. At all other times there is even more redundancy. Although I'm using the lander as the primary propulsion source for trans-Earth that's in the context where both the transit vehicles have their own secondary propulsion using 2 of the same engines. Beyond return to Earth the lander is recycled as a vehicle that will and fuel on Mars. That's a non manned mission. And a non safety critical process for exactly the same reason discussed above. And beyond this the lander becomes spare parts.

What I'd ask you to do is to specify exactly the failure mode and scenario and then I'll address a response to that.

And for the sake of completeness I'll point out again that one thing I like about a mostly propulsive lander and ascent vehicle is that its inherently simplicity. Not just less things to go wrong. Its also about servicability. Its all out in the open. Its easy to visually inspect. Its easy to repair. And the lander that ferried fuel to the surface is a ready made source of spares, including whole engines, which at this scale can be man handled. The crewed landing is effectively being done on pristine engines. You land, you inspect, you swap components if necessary. You've got an ascent vehicle that has a firing sequence that if there are problems in the first seconds you do a soft landing and go back and fix the problem. And having transit vehicles in both low and high Mars orbits gives you a very long effective launch window so these sorts of problems can be fixed.

Again, specific failure modes please?

As for your last question. Its aerobraking that has obvious and hard to fix safety issues. Its time to think laterally.

As for L2. Using L2 does not depend on solar tugs. But having a base at L2 might benefit from solar tugs - provided that we're doing this on a very large scale and over many missions. Otherwise the development cost of a solar tug doesn't save enough to be worth it.

My architecture does not depend on the means by which you get to high Earth orbit, or for that matter which high Earth orbit it is. It could be L2, or it could be simply a high, elliptical orbit. The latter being slightly less energy.

I have noted that the only real advantage (to me) in using L2 is that what you have at L2 isn't traversing the radiation belts on each orbit. Arguably a problem for systems. Definitely a problem for people. But again, the latter has workarounds. And the former probably means running systems in safe mode at times.

The big advantage for others in having a base at L2 is proximity to the moon. I'm not fussed. Its also having a steady platform at which you can do manned assembly. Again I'm not fussed because what I've proposed is bump together assembly.

As for large unmanned cargo deliveries I have no problem with simply very large heat shields. As much fun as inflatable heat shield extensions are, there is doubt in my mind that they cannot be matched by light weight but still rigid heat shields of the order of 30m diameter. Which of course requires manual assembly. And that to me means using low Earth orbit. Doing so (with a rigid heat shield) means taking the most advantage of emerging light weight ceramics (GW would be happy to hear me say this). Anyhow I'm not wedded to any particular solution and your mileage may vary.

Last edited by Russel (2013-06-29 09:17:01)

GW Johnson · 2013-06-30 11:56:08

When I ran my Mars lander studies, I was able to achieve 5 km altitude at local Mach 3, simply by shallow-angle entry from low Mars orbit. The required delta vee for ascent (complete with gravity and drag loss estimates) is about 3.7 km/s to 200 km. That's the absolute minimum for powered landing flown as ascent-in-reverse, but you'll actually need around another km/s more, to cover all the real-world maneuvering and hovering.

To aerobrake from LMO, with my rather large vehicles (at 400 kg/sq.m ballistic coefficient, some 60 ton size at entry), I needed about 50 m/s de-orbit burn, and hit the "air" at about 1.6 degree down angle. Local Mach 3 at 5 km altitudes is 0.7 km/s velocity. So, I braked by 3 km/s without using but 50 m/s worth of propellant. From that point, theoretically I need only 0.7 km/s to land, but the actual real-world maneuver delta vee was closer to 1.4 km/s.

Still, mass ratio for 1.4 or 1.5 km/s delta vee, or mass ratio for at least 3.7 and very likely 4.7 km/s delta vee, just to land? You decide which one is the bigger, more expensive thing to launch and bring from Earth. Bear in mind, the ascent will require about 3.7 km/s, no matter how you choose to do it. That propellant all has to come from somewhere.

The aerobrake entry velocity is 3.7 km/s, which is a lot of local Mach numbers, but also a lot less energetic than from LEO (11 km/s). Peak heating rates are proportional to entry velocity cubed (not squared). Mars is a lot less demanding. PICA-X ablatives are a nice choice at 0.27 sp. gr. and about 1.6 inches thick, for one-shot vehicles. My ceramic heat shield study showed that black-surfaced alumino-silicates would be fine, even for direct entry at 5.6 km/s, as long as the angle was shallow. Skin temperatures at the stagnation point fell well under the 1290 C phase-change limit for shrinkage-cracking in alumino-silicates.

My "stuff" would be reusable, damage-tolerant, and easily repairable for fully-reusable vehicles. Think under an inch thick at densities near 0.03. Like shuttle tile, but a whole lot tougher, and buildable in large bolt-on panels. Two-component ceramic composite, laid up a lot like fiberglass is.

There's no need to try inflatable heat shields at Mars (plenty of use here from LEO though). You just go from hypersonics to rocket landing. There's a lot of tradeoffs yet to be done to select such a flight path, but the feasibility is there right now, and with very large capsule-like vehicles massing many tons.

GW

Last edited by GW Johnson (2013-06-30 11:59:28)

GW Johnson · 2013-07-01 07:52:01

Josh:

Go take a look at XCOR Aerospace's web page. Select "our products" and go look at their rocket engines and their piston propellant pumps.

They're still down in the 5000 lbth (and smaller) class, but this is what long-life reusable liquid rocket engines look like. When I was there 3 years ago, I got to talk with their head and their chief engineer about these devices. The engines they flew in the rocker racer aircraft have an estimated time-between-major-overhaul exceeding piston engines.

These engines have failure containment built-in, too. Thrust/weight won't look anywhere near as high as the one-shot designs everybody is used to seeing, but Isp is every bit as good. The extra weight is what you have to pay to achieve long life and inherent safety.

This technology looks like it is scalable to larger sizes, to me. It is what they will fly in Lynx.

GW

Russel · 2013-07-01 07:57:13

GW Johnson wrote:

When I ran my Mars lander studies, I was able to achieve 5 km altitude at local Mach 3, simply by shallow-angle entry from low Mars orbit. The required delta vee for ascent (complete with gravity and drag loss estimates) is about 3.7 km/s to 200 km. That's the absolute minimum for powered landing flown as ascent-in-reverse, but you'll actually need around another km/s more, to cover all the real-world maneuvering and hovering.
To aerobrake from LMO, with my rather large vehicles (at 400 kg/sq.m ballistic coefficient, some 60 ton size at entry), I needed about 50 m/s de-orbit burn, and hit the "air" at about 1.6 degree down angle. Local Mach 3 at 5 km altitudes is 0.7 km/s velocity. So, I braked by 3 km/s without using but 50 m/s worth of propellant. From that point, theoretically I need only 0.7 km/s to land, but the actual real-world maneuver delta vee was closer to 1.4 km/s.
Still, mass ratio for 1.4 or 1.5 km/s delta vee, or mass ratio for at least 3.7 and very likely 4.7 km/s delta vee, just to land? You decide which one is the bigger, more expensive thing to launch and bring from Earth. Bear in mind, the ascent will require about 3.7 km/s, no matter how you choose to do it. That propellant all has to come from somewhere.
The aerobrake entry velocity is 3.7 km/s, which is a lot of local Mach numbers, but also a lot less energetic than from LEO (11 km/s). Peak heating rates are proportional to entry velocity cubed (not squared). Mars is a lot less demanding. PICA-X ablatives are a nice choice at 0.27 sp. gr. and about 1.6 inches thick, for one-shot vehicles. My ceramic heat shield study showed that black-surfaced alumino-silicates would be fine, even for direct entry at 5.6 km/s, as long as the angle was shallow. Skin temperatures at the stagnation point fell well under the 1290 C phase-change limit for shrinkage-cracking in alumino-silicates.
My "stuff" would be reusable, damage-tolerant, and easily repairable for fully-reusable vehicles. Think under an inch thick at densities near 0.03. Like shuttle tile, but a whole lot tougher, and buildable in large bolt-on panels. Two-component ceramic composite, laid up a lot like fiberglass is.
There's no need to try inflatable heat shields at Mars (plenty of use here from LEO though). You just go from hypersonics to rocket landing. There's a lot of tradeoffs yet to be done to select such a flight path, but the feasibility is there right now, and with very large capsule-like vehicles massing many tons.
GW

As you probably know from the landing thread, I coded a simulator for landing. I got generally similar results. Where we differed was because of the initial hypersonic trajectory curving under gravity. Nevertheless the results that matter to me were to do with a small manned lander (5 to 10 tonnes). In order to give that lander enough margin (better than 5Km at Mach 3) I invariably had a lander with a base diameter over 6 metres. That suited me because a 10 tonne lander and ascent vehicle needed tankage of about this scale anyhow. But it was still a beast.

Its not quite as simple as an ascent in reverse. On ascent, atmospheric drag works against you. On landing atmospheric drag works for you. If you refer to my above posts you'll see the justification for the figure of 3.8Km/s. That takes into account drag, plus the fact that the planet and atmosphere is rotating relative to the inertial frame of reference in which the initial 3.5Km/s velocity applies.

So I'll stand by the 3.8Km/s which does include allowance for final maneuvering. And with it a mass ratio of 3.

The propellant has to come from somewhere, but so does every other mass.

Let's take another look at the problem of landing people on Mars and getting them back to Mars orbit. This time I'll come up with a fair representation of a conventional landing and attempt to compare it. It uses an expendable ascent vehicle which has been landed previously inside a conventional lander. It also uses a conventional and expendable lander.

The expendable lander has a mass of 7 tonnes. Taking a requirement of 500m/s for maneuvering and an Isp of 250 for the lander and the lander needs 1.5 tonnes of fuel. So 8.5 tonnes in LMO.

The ascent vehicle has a mass of 7 tonnes. Taking a requirement of 4.5Km/s for ascent to orbit and an Isp of 350 and it needs 18 tonnes of fuel. So 23 tonnes before takeoff.

The ascent vehicle needs to be landed in the first case. Given a landed mass to entry mass of 2 to 1 that means the ascent vehicle started life as 14 tonnes in LMO

18 tonnes of fuel for the ascent vehicle comprises 4 tonnes of methane and 14 tonnes of oxygen.

Now, for the sake of the exercise I'm going to suggest that for every tonne of oxygen produced on Mars you need 50Kg of ISPP and power supply. And for every tonne of methane produced you need a further 100Kg of ISPP and power supply. This is rough but it reflects the relative complexity and energy intensiveness of simply splitting CO2 into O2 versus the steps taken in producing methane from hydrogen.

So now we need 700Kg of plant for oxygen, and 400Kg of plant for methane production. Now, in fairness, the plant is reusable, or will be when its well tested and well understood. So I'm going to allow it to provide the fuel for 2 ascents. Meaning the mass per mission is 550Kg.

Of course we had to land that so we now have 1100Kg in LMO being for the propellant production.

We of course need to land hydrogen. And this time I'm going to go beyond my better judgement and only allow for a 25% loss factor. So 1 tonne of hydrogen needed or 1.25 tonnes landed.

That would be 2.5 tonnes in LMO but for the fact that a hydrogen lander is going to involve large tanks than if you used methane. 4 tonnes of methane might require 1 tonne of hydrogen, but 1 tonne of hydrogen requires 1.5 times the volume of the methane. Given a tankage factor of about 50Kg per m3 (that's a very light weight, composite tank) it turns out that the hydrogen needed requires an 870Kg tank. So we're actually dealing with 4 tonnes of hydrogen, tank and requisite landing vehicle.

Adding this all up we get

8.5 tonnes in LMO for the lander
14 tonnes in LMO for the ascent vehicle
4 tonnes in LMO for the hydrogen
1.1 tonnes in LMO for the propellant plant

Now, lets do the same sums involving a integrated lander and ascent vehicle doing a mostly propulsive landing.

We start with a vehicle with a mass of 5 tonnes on entry. Its using methane/LOX at an ISP of 350 and requires a delta V of 3.8Kms so a mass ratio of 3.
Thus the fully fueled lander has a mass of 15 tonnes in LMO. Ok, so its heavier.

Now the ascent vehicle is one and the same. And since its 5 tonnes it needs less fuel for ascent. With the same mass ratio as above, it needs 13 tonnes of fuel.

13 tonnes of fuel breaks down to 2.888 tonnes of methane and 10.111 tonnes of oxygen.

By the above rules we need 500Kg of ISPP plant and power source. According to the same rules as above (used over 2 missions) that's 250Kg of plant per mission.

And I'm going to land the ISPP plant via a conventional lander so with a 2 to 1 mass multiplier thats 1 tonne in LMO.

Of course the expensive step is delivering the fuel. For that I need a 3 tonne lander (minus crew compartment and life support) and 3 tonnes of methane. That's 6 tonnes. Again with a mass ratio of 3 I need a further 12 tonnes of fuel in LMO. So all up in LMO I'm up for 18 tonnes. Ok, I could have landed the methane more economically, but I'll get to that.

Adding this all up we get

15 tonnes in LMO for the lander
0 tonnes for the ascent vehicle
18 tonnes in LMO in order to land the methane
0.5 tonnes in LMO for the propellant plant

Now, at this level we're comparing 27.6 tonnes for the conventional approach versus 33.5 tonnes for my approach.
A little commentary is in order.

Despite the complex sums involved in methane versus hydrogen, that choice is a second order factor in terms of overall mass.

The expensive factor in a conventional approach is landing an ascent vehicle, which because its being landed in a higher stress environment, has to be heavier, all else being equal.

The expensive factor in my approach is landing the methane fuel itself in a lander that uses mostly propulsive braking. I could have instead adopted the approach of landing the methane conventionally. That would have saved me the better part of 10 tonnes, and put me ahead of the conventional approach overall. But I didn't do so ultimately because this leaves a host of contingencies. Having an identical vehicle parked on the surface of Mars means at the very least you'll have spare parts, if not a fully functional vehicle. Its also serving another purpose and that's storage tanks.

Now for the bigger picture. The most commonly accepted approach in getting back from Mars is having a return vehicle, fully fueled and provisioned brought to LMO. That's a mass of tens of tonnes. To get to a near escape orbit around Mars that vehicle needs a delta-V of 1.3Km/s. Another mass multiplier of 1.45. So relative to the same vehicle parked in a near escape orbit you've lost tens of tonnes of fuel.

Now, what I've done is park the return vehicle in a high Mars (near escape) orbit so I've incurred a very large saving reflected back to low Earth orbit. That's the primary insight here. In order to make this happen though I need a ferry that will get me from LMO to HMO. That ferry comes in the form of the ascent vehicle. Now a 5 tonne vehicle requires only 2.25 tonnes to act as a ferry between orbits. That's an enormous saving.

Now of course you could take a conventional ascent vehicle and do the same thing. Only it would be (somewhat) heavier and would have to be required to do multiple burns (mind you, ANY such vehicle would, given the possible contingencies involved).

I'm also requiring a separate habitat be sent to low Mars orbit in advance. That's the staging point. But such a habitat whilst based on the same design as the main transit vehicle does not have to use the same degree of provisions or shielding. In other words about half the mass of a fully provisioned vehicle of its type. That extra mass I can justify within the overall budget because of the very large mass savings on the main transit vehicle being parked in high orbit. And because its sent to a low Mars orbit in advance, it and the fuel that gets used by the landers can be sent on a lower energy trajectory ahead of time.

One other thing. The primary propulsion for the transit vehicles is the lander. Now, since the lander must travel with the crew and must return to high orbit, there is minimal further cost in using it as the primary propulsion for the return journey. Instead I could provide the return vehicle with a scaled up propulsion unit and ditch the lander. That would actually save some mass. But at the expense of not being able to deal with some contingencies particularly on approach to Earth space. The lander remains at all times a last-ditch lifeboat. And I like it that way.

Ok, enough. I hope that clears things up some.

Now, regarding the thermal protection. My interest there is about the transit vehicle. One option is aerocapture. Using the thermal mass of the water on board and a modest thermal barrier I believe it would be possible to do aerocapture. I lean against aerocapture for two reasons. One is that the solar arrays would have to be stowed. The other is that the entire vehicle might have to be strengthened to cope with the stress of aerocapture. (Or its just possible it might be inherently strong enough thanks to all the layers of hull and radiation shielding anyhow).

Now assuming that aerocapture can be justified, what I need is a material that is light weight, modular and stands up well to long periods in space. That's why I'm curious about your ceramic composite.

Now the interest bit is you say it can be repaired. How? And can that be done in space? Or is it better to keep spare panels?

As I said in an earlier thread, the energy involved in Mars aerocapture is about the same as the energy involved in a full landing from low orbit. But it gets worse doing aerocapture into Earth orbit. That's where the material will be tested. Any guesses at the temperatures?

JoshNH4H · 2013-07-01 10:23:26

I think the point you've not fully understood is that I've always operated on the basis that anything critical to safety will be landed on Mars and verified as operational before the crew leave Earth. So the safety issue you imagine isn't there. It a very simple principle. An unmanned mission lands what you need on Mars. Crew follows. There's no such thing as crew landing with life critical equipment.

What do you do if your lander crashes? Give up? Tell the astronauts and the contractors that you appreciate all their hard work but this mission is cancelled? Do you push your mission timeline back 26 months (At a cost of potentially billions of dollars, plus the more intangible risks of bad press and loss of interest by investors or Congress and the American public*)? No, you don't. You pull out spares and build new equipment where necessary and send it in the same launch window as your crew. If possible send it so that it enters while the crew still has the option for a free return trajectory. I would say that this is okay even if it means using less delta-V and extending the crew's transit time by 2 weeks (It's a trade-off: Mildly more radiation for the crew for the increased safety of knowing that you're not screwed if the aerobraking fails again-- likely or not it's an eventuality that any good mission design will account for).

Again, I reject your mass multiplier of 2 for aeroentry. I disagree with the statement that the decelerations will be higher than powered landing. Inflatable heatshields are, if not entirely necessary to land large payloads, what I would consider an enabling technology. They make it possible to have a larger heat shield area given constant fairing diameter. This means that you will shed more velocity in the higher atmosphere and lowers the peak heating, as well as peak decelerations. I would expect rocket deceleration for the last half a km/s or so. This means that you only have one heat shield and a bit of fuel+engines.

Reliable, lightweight aeroentry is a vital technology for the colonization of the inner solar system. Why develop it later rather than sooner? This is the question that you need to answer (I am after all defending the Mars Direct plan). This is the question that I don't think you have an acceptable answer to.

*Yes, the US Congress and the American public. What other country could go to Mars any time soon?

Russel · 2013-07-02 07:04:20

JoshNH4H wrote:

I think the point you've not fully understood is that I've always operated on the basis that anything critical to safety will be landed on Mars and verified as operational before the crew leave Earth. So the safety issue you imagine isn't there. It a very simple principle. An unmanned mission lands what you need on Mars. Crew follows. There's no such thing as crew landing with life critical equipment.
What do you do if your lander crashes? Give up? Tell the astronauts and the contractors that you appreciate all their hard work but this mission is cancelled? Do you push your mission timeline back 26 months (At a cost of potentially billions of dollars, plus the more intangible risks of bad press and loss of interest by investors or Congress and the American public*)? No, you don't. You pull out spares and build new equipment where necessary and send it in the same launch window as your crew. If possible send it so that it enters while the crew still has the option for a free return trajectory. I would say that this is okay even if it means using less delta-V and extending the crew's transit time by 2 weeks (It's a trade-off: Mildly more radiation for the crew for the increased safety of knowing that you're not screwed if the aerobraking fails again-- likely or not it's an eventuality that any good mission design will account for).
Again, I reject your mass multiplier of 2 for aeroentry. I disagree with the statement that the decelerations will be higher than powered landing. Inflatable heatshields are, if not entirely necessary to land large payloads, what I would consider an enabling technology. They make it possible to have a larger heat shield area given constant fairing diameter. This means that you will shed more velocity in the higher atmosphere and lowers the peak heating, as well as peak decelerations. I would expect rocket deceleration for the last half a km/s or so. This means that you only have one heat shield and a bit of fuel+engines.
Reliable, lightweight aeroentry is a vital technology for the colonization of the inner solar system. Why develop it later rather than sooner? This is the question that you need to answer (I am after all defending the Mars Direct plan). This is the question that I don't think you have an acceptable answer to.
*Yes, the US Congress and the American public. What other country could go to Mars any time soon?

I'm struggling to follow you here. The first question you ask is "what do you do if your lander crashes?". Which lander in which scenario is my first question. So lets go through some possibilities here.

I think most Mars architectures call for some things to be landed robotically. Correct me if I'm wrong there. Now, for instance if you were to land an ISPP plant and then wait two years before landing humans, what do you do if the robotic lander crashes, or the equipment fails? Give up? Or do you fix the error, launch again, and tell the crew they're going to have to wait? As I see it, this particular outcome is a definite possibility in any architecture that seeks to establish mission critical infrastructure ahead of a manned landing. Can't be avoided.

Now, are you considering the type of mission where mission critical infrastructure is launched during the same launch window as the crew? If so you've got two basic options. One is you land all your eggs in one basket. That's the land everything on one very big lander approach. As we discussed above, this has some serious safety issues. Now an alternative which you've raised above is to send the mission critical stuff slightly ahead of the crew. Meaning separate landings.

Thing about this approach is now you've got two separate landings. And you're headed down the track where the manned lander (the one delayed by 2 weeks) is fairly minimal. And to the extent you're not weighing that lander down with stuff that could have been landed 2 weeks sooner, and thus you've got a lighter lander and one that you can afford to build more margin into, then I'm happy with it.

Problem though is that you've reached a point where if something goes wrong with the first lander then the crew are either committed to hanging around in Mars orbit for a long time, or a free return trajectory. And what you've done here is trade the possibility that the crew will have to wait another 26 months on Earth with the possibility that the crew will have to go through the return journey and then again wait until someone figures out what went wrong and there is an opportunity to fly again.

Now, I'm committed to safety wherever technically possible. Which means a principle of anything mission critical functioning, or people waiting. You've pulled me up on the consequences of that, but the reality is that mission scrubs and long delays are a possibility in absolutely every architecture.

I could have chosen to trade off safety in various ways. For instance, a free return trajectory is well within the capability of my system. So if you were really in a blind hurry you could indeed send everything on the one launch opportunity. Ok, so the non manned lander crashes. What then? I've sent two fully fueled transit vehicles towards Mars. They don't capture. Instead they return home. Nice thing here is that instead of having to be very fussy about the trajectory in order to guarantee a free return, I'm in a position to plan non free return trajectories that can be converted to a free return. Because I've got fuel reserves both in the lander - that was tasked to ferry the crew to low orbit, and in the two vehicles. At last resort I can limp home in on transit vehicle.

The other nice thing about not having to be too fussy about having a free return trajectory in the first case is to minimise the length of time the crew spend in space. An issue you've noted.

Now, honestly, faced with a choice between that scenario, and simply opting to send the non manned mission out 26 months earlier, I'd choose the later even before scratching my head about it.

You simply don't get a choice. There will always be situations where you have to scrub missions. And its better to do that with your crew waiting on Earth than all the other alternatives.

Quite apart from the ethical considerations of sending people into space in its own right, do you stop to think about the risks to reputation and public enthusiasm that are not just oh dear we'll have to wait but oh no we just cratered a 60 tonne lander and there were people inside? People will forgive mistakes if its just stuff. They won't forgive mistakes if needless risks were taken with people and you can stop and think about the risks taken with the shuttle program.

By far and away the safest Mars architecture is one where do everything to protect the crew, including avoiding hurling people through the thin Martian atmosphere strapped to a very large brick. And if your concern is that Mars should be an ongoing venture, the biggest risk to that is not delays - its the commission of inquiry that will follow a loss of crew and the revelation that you could have done it more safely.

If you don't get why an aerobraking landing is necessarily more physically stressful you don't get the fact that the Mars atmospheric density profile is essentially an exponential curve. The only way you avoid higher g forces is to slow down earlier. That's what a mostly propulsive landing does. To get the similar effect with drag you need bigger.. MUCH bigger. So that you're already getting significant drag higher up - around 80Km. Worse part is that anything you use to create drag either has to have a variable size, or you again suffer the fact that if you've created enough drag high up, you've created way too much drag further down.

Inflatable heat shields designed to create significant drag in the higher atmosphere will be huge - 50m for a small manned lander and well over 100m for the sorts of landers you're contemplating. And what do you do when the stress becomes too high? You have to shed it and rely upon a smaller drag device. This staging effect adds mass, complexity and risk.

How about a rigid heat shield? Well for a smaller lander (under 10 tonnes) you might consider 6m diameter providing you want to hit 4g lower down. But if you want to make the same heat shield big enough to create enough drag higher up to lower the g forces, then you're going for an in Earth orbit assembled 30m affair. Obviously the mass goes over the top. Try and do the same thing with an inflatable hybrid and you need to lose the inflated section before it shreds itself. More risk.

Essentially the laws of physics are against you. There is no known way to build something that provides enough drag higher up in the Mars atmosphere that would give you the benefit of lower g forces. And do so without defeating the purpose which is trying to avoid weight.

Now, there is lift. You'll notice that MSL used lift. But it did so at the expense of throw away weights.

You could really light weight wings. But you'd either need those to fold or more likely to jettison them. Again, more mass.

Even though a minimal lander might cost you 5 tonnes in low Mars orbit, there is the extra structure and the reason for that is that it's reacting against the shell of the heat shield. That shell has to be stiff in order to avoid buckling and cause aerodynamic instability. Then you have to have a back shell under aerobraking conditions. And every time you decide to make the heat shield bigger, you also scale the back shell. You're lucky to get out of it for 7 tonnes. Now, add your parachute or other drag devices if you deem them necessary (they may just make things worse) and then the terminal descent fuel which at 500m/s using lower Isp fuel gives you a further mass multiplier of 1.22. So a 7 tonne craft is now 8.5 tonnes. And that's not the end of the story.

See, I'm creeping up on 2 to 1 here? Now, we may do better than that. But not significantly better. Not when the scheme of comparison include the system wide consequences. A mostly propulsive landing will always require more mass in LMO. How much more doesn't particularly matter. What matters, and I've said this before, is that when you don't think about landing in isolation, and think about the whole architecture, it makes a whole lot of sense.

Now, some failure modes, please?

Russel · 2013-07-02 07:33:25

JoshNH4H wrote:

Reliable, lightweight aeroentry is a vital technology for the colonization of the inner solar system. Why develop it later rather than sooner? This is the question that you need to answer (I am after all defending the Mars Direct plan). This is the question that I don't think you have an acceptable answer to.
*Yes, the US Congress and the American public. What other country could go to Mars any time soon?

Reliable lightweight aeroentry is vital to landing large masses on Mars. And I'm not suggesting it be developed later rather than sooner. Seriously, I'm not. It just doesn't matter as far as the crew goes. It will always be safer to land the crew separately (And indeed that is what you're suggesting if you have the kind of mission where the crew arrive 2 weeks later).

I will differ on one issue and that's I think you don't need to land anything on Mars larger than about 15 tonnes. Practically anything you can think of can be built up from that scale. I see Mars drive as a response to the "big lander problem" but I believe they're going a bit too far in the other direction. In any case they don't really make the landing problem go away either.

You're defending Mars direct. Fortunately I'm not

What I do see though is a lot of frustration people have with NASA coming up with ridiculously large mission mass. And thus a lot of architectures that are attacking the problem from particular directions. Mars direct is largely a response to the NASA overkill. I kinda like it in some ways. Its got a certain minimalism. But, for me the solution ultimately lies in something a bit more considered, a bit more conservative, but still not wasteful either.

However we get to Mars, its going to take a decade or two. And that's time enough to for everyone to sit around and bang heads together and not get too wedded to their particular approach.

NASA isn't the only body that theoretically could do this. I think ESA has the resources to do this eventually. Japan, China could all play the part.

The thing about the US is that if you've got a great architecture, whose going to build it. If its NASA then you've got to get through its strange meld of hard core rationality, plain old fashioned bureaucracy, and US style corporate welfare. Not that the ESA doesn't have its own politics at play.

Point is, even the most spartan Mars architectures are not pocket money. Eventually the big space agencies will have to be won over. And if that's the case they're going to be won over with with architectural proposals that don't just save launch costs, but also simplify development effort and most of all don't make them look like they're taking unnecessary risk with their crew.

And that's about as far into politics as I dare venture. I'm interested in this as a problem worth solving that hasn't been solved well by anyone, yet.

We do have time to get the technology right, to scrutinize what we're doing in unprecedented detail and do it safely, and I might add, in style.

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#24 2013-07-02 07:04:20

Re: Yet another Mars architecture

#25 2013-07-02 07:33:25

Re: Yet another Mars architecture

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