Landing on Mars

Russel · 2012-12-25 00:59:27

JoshNH4H wrote:

To address what is, I think, the most glaring error in your post:
Whether [multiple launches of a Mars Ascent Vehicle are] a reliability issue or not has yet to be established.
While you later suggested that each mission would only involve one launch and one entry, subsequent missions would reuse the same craft. So what you would have is a six month flight to Mars, followed by a long period of stasis (either in LMO or on the Martian surface). There will be an atmospheric entry at some point, followed by a 4 km/s launch. After two years of stasis, repeat. Two more years of stasis, repeat again. With an LMO orbital period of two hours that's hundreds or thousands of thermal cycles for each rocket, followed by a high thrust launch (even though the Martian gravity is lower, it's still pretty high thrust and high powered). After this launch is an aeroentry followed by rocket braking and vertical soft landing.

Your quite correct in saying that it would be challenging to design something that unattended could last for years - particularly long periods of stasis.

But, your assumptions here are wrong.

Firstly these vehicles don't spend more than a fraction of their lives in orbit. Most of the time they're parked on the surface.
Secondly, they're not unattended. They will be inspected and if needs be serviced. Just like anything else we design here on Earth that is intended for years of service.
Even the fuel ferry will have the opportunity for inspection and service.

Its the same with the transit vehicle. Its not going to be re-usable without some level of human intervention. But re-usability is an inevitability.

Yes, you can make the transit vehicle a throw away item. But the price to be paid is having to loft it with every mission and the consequences of that is you get a bunch of people finding ways to mass reduce and the inevitable consequences.

JoshNH4H · 2012-12-25 01:10:33

Modifying my numbers for Carbon Monoxide for an Isp of 260 s (My source said 250 s but the number is not cited and I don't know where it comes from; I'm perfectly willing to accept a NASA report as more authoritative) that Atlas IIIA stage would have 18.6 tonnes and 26.2 tonnes of payload to LMO for Isps of 260 and 290 respectively, for a payload to surface mass ratios of 1.36 and 1.91 respectively. Clearly in this matter Isp is key, and I would like to retract my statement on Carbon monoxide vs. methane fuels because it is no longer borne out by the data and because the use of CO instead of CH4 as fuel is a simpler process: It requires only electrolysis of compressed, liquefied CO2. I look forward to a more extensive reply.

Russel · 2012-12-25 01:40:14

JoshNH4H wrote:

The ERV has more payload than both your Fuel Ascent Vehicle and your Personnel Ascent Vehicle. However, even ignoring the reusability issues of your architecture Mars Direct (and derivatives) have a major benefit that yours doesn't: Fewer rockets and fewer transfers. Each engine firing is a chance for something to go wrong. When your crew and all of your payload are separate in two rockets, the failure of either rocket results in the death of the crew. This approximately doubles the chance of a loss of crew (technically, it is [1-(chance of one rocket succeeding) x (chance of another rocket succeeding)] ). The use of two rockets where one can be used is done not once but twice in your architecture, both in sending things to Mars and in launching things from Mars to LMO. While you try to compensate by suggesting that the fuel rocket could carry crew to orbit (without suggesting a way to mount the capsule atop the rocket), separating the rockets does not enhance crew safety because without fuel for return the crew is also dead.

When considering safety issues you have to consider the time scale and sequence. If for instance you were to climb into a rocket designed to launch from Mars surface and go directly to Earth, you'd have to be pretty sure that one rocket will work. Of course you can build in redundancy and even testing - but something of that scale is a little hard to test fire.

If you stage it and instead rely upon two rockets to get you home - one to Mars orbit - then of course you've got two vehicles you have to rely upon. But its a bit more complex here.

Firstly the vehicle parked in orbit has a twin. A fully fueled and provisioned twin. Its a consequence of re-usability that you can afford to do this. In the worst case scenario you've got yourself a few days to fall back onto the spare. Likewise with the ascent vehicle. The reason I am still fond of a vehicle with commonality of design is that either your fuel ferry can be a spare or can be sacrificed or cannibalized in other ways.

Now, if you really, truly, want belts and braces you have 3 vehicles. One ferry, two crew. Again, assuming re-usability over a number of missions that's still a huge saving of fuel in LEO terms. I don't know if the extra spare is really needed. But the point here is that by the time you've climbed into you ascent vehicle you have had the opportunity (2 years on the surface) to verify/inspect/service these vehicles.

Also, being smaller vehicles, and vehicles that are better suited to repeat soft landings its not as difficult to test fire them either. Yes, you can test fire a "direct to Earth" counterpart but that begs the same questions about reliable repeat firings, doesn't it.

So in short, time is on your side when on Mars. Use it.

Now, going back to the amount of fuel. What we need here is some simple, albeit rough basis of comparison.

What I'd suggest is that if we were to cast aside all thoughts about mass that has to be landed on Mars, and will stay there, that will simplify the comparison.

But I'll grant you that the landing/ascent vehicles that I'm talking about do need to get there, at least once. So we can go into more detail about their mass and and the required fuel.

One other wrinkle here is this. What's going on in the background is the thought I discussed many pages back, and that is if you can draw a line between the mass you need to send to Mars just once, and the mass that you have to send with your crew, you then have the freedom to think about propulsion options.

For instance, its not impossible that you could send all the heavy plant and equipment (the Mars, hab, the landing/ascent vehicles, rovers.. etc.. etc) in the most fuel efficient (but slow) manner you can think of. Think as an extreme example, ion propulsion.

But even better, think about a relatively small nuclear plant - a MW or so. Useless for a fast crew mission, but very efficient for just this purpose.

(And there's lots of other alternatives in the pipeline)

If you can solve the problem of minimizing fuel for the bulky/heavy non manned part of the mission, then you're onto a winner because you have more freedom to send more mass and thus you can afford to add things that in short add up to more comfort and more safety.

Having done that, you're then left with the crewed vehicle. And again my approach (as you might see if you read back a number of pages) was to avoid having un-necessary mass travelling with the crew - and that's where not having to travel with your lander comes in. So I'm basically down to the transit vehicle and its modest propulsion system - again you don't have to blast your way out of the galaxy, you just need a few KN of thrust. Again, mass saving.

As a rough guide, we're talking about 20 tonnes of transit vehicle, give or take, all up. We can argue the toss about he exact amount but lets have a fair basis of comparison.

But at that mass level, and with methane/LOX you basically need to double your mass with fuel to get back from Mars orbit to Earth. (Assuming aerocapture, which this thing has to be designed for).

Oh and btw.. other missions that assume aerocapture end up with a lot of extra baggage which then simply adds to the task of successful aerocapture, and your mass starts climbing - rather like a bad car design

So, and again we can haggle, 20 tonnes of fuel equates to roughly 4.5 tonnes of methane (the oxygen comes from Mars).

Going back a step you've arrived in Mars orbit with a vehicle weighing (its not quite empty) about 27 tonnes.

Going to Mars we'll assume a delta V of about say 4.5Km/s (I could have cheated and boosted this thing to a high Earth orbit first but lets keep it simple). Again feel free to haggle over the numbers.

So a mass ratio of 3.2 and total vehicle mass of 86 tonnes. About 60 tonnes of fuel, of which about 13 tonnes if methane, not including the 4.5 tonnes you need to keep in Mars orbit.

That's the rough starting point. And again, you can easily fiddle with that by assuming more exotic things like a higher Earth orbit, solar powered tugs etc.

Anyhow.. I guess if we go further we need some basic assumed masses etc. So feel free

Xmas dinner! cya!

JoshNH4H · 2012-12-26 20:55:34

Russel- Hope you had a good Christmas!

With regards to the topic at hand, specifically your post numbered 326 in this topic:

The location where the Mars Ascent and Fuel Ascent Vehicles spend their time is the least important part of the safety issues inherent to this mission plan. While time spent in space and thermal cycling is never good for mechanical reliability, let me be clear: The real concern here is repeated launches and landings. Take a look at rockets on Earth: Even for one launch after which they are thrown out, you need hundreds or thousands of people working on it for months. Why do you expect that a rocket on Mars will be better than one on Earth? It's all well and good to say that it will need servicing similar to a boat or an airplane, but it's quite a job to build a rocket like that. It hasn't been done on Earth. Compare hundreds of highly qualified people, focused only on the task of making sure a rocket is safe to six, whose primary task is necessarily not the rocket but exploration, science, and prospecting.

Also remember that even on Earth, a fully reusable rocket with this kind of delta-V has never been built. Why can we posit one on Mars? In the long run it's inevitable, but given the technological challenges I think "the long run" will happen long after the first mission. The transit stage is somewhat more of a tossup, but unless it's a nuke it's probably not worth it- at least for a while. My reasoning here is that a lot more of the mission mass and expense is going to be caught up in the surface gear and the launching thereof than the transfer vehicle for the crew, which is the only part of this operation that could sensibly be reused.

Again, in the long term reusable transports of all kinds are not only a good idea but an essential one. But for the initial exploration missions it is not cost effective, and more importantly, not safe.

With regards to your post numbered 328 in this topic:

You make several arguments in the post. They boil down to a statement that the logistics of reusability mean that you can plan for more and safer abort modes for that reusable mission, thus making it safer. This is the first part of the post.

Your plan is to send two personnel ascent vehicles and two Fuel Ascent Vehicles to Martian orbit and leave one set there as a safety factor for every mission. Fully fueled? Or would the crew have to wait to refuel it, miss the launch window, and then return two years later? With what supplies will they last? How do you plan to store the hydrogen (yes, Hydrogen... Why are you recommending bringing methane? Insanity)

You suggest that testing of the Earth Return Vehicle (ERV) constitutes two firings and thus it's not a good idea. Let me ask you, what do we do with single-use rockets on Earth? We test rockets of the same design then build one, test the parts, check it out a few times, then launch it. Usually it works. Up to 99% of the time for some rockets. It would be, again, insanity to try it twice without a full team (hundreds if not thousands of workers, traditionally, perhaps tens if you run your operation perfectly well) to fully check out the rocket again.

Sending an ERV is simple. Easy. Reliable. Not really that expensive. The Mars Direct plan provides for backup in the form of the ERV for the next mission, sent two years early (at the same time as the crew) and fully fueled by the time they are leaving. If for any reason the ERV is deemed untrustworthy, there is a second one ready to launch.

It seems like you're trying to save mass. Let me throw the question: Say you save 10 full tonnes of mass to LEO per mission. SpaceX can do $3,000/kg. That's 30 million dollars. What's 30 million dollars in launch costs on a multibillion dollar mission at the gain of more safety?

Russel · 2012-12-27 05:34:19

Hi Josh.. had a Merry and all too fattening Xmas

Ok, lets step back a little and look at some assumptions and in particular the technology we may or may not rely upon.

Clearly, if we started with the technology that is available today, we'd either conclude we just don't know how to do it (particularly EDL) or we'd have to resort to some pretty heavy handed (and costly) techniques. Those that do say we can fly today, with today's technology are either making assumptions about what technology we can develop (ISPP etc) or openly declaring "this is forward work".

There are two technologies or strictly speaking capabilities I am fairly confident will be developed. One is the capability to launch into low Earth orbit at relatively lost cost ($1000-$1500/Kg). The other is the capability to launch and relaunch from Earth multiple times. Indeed, the two are actually sides of the same coin. Elon Musk figured this out and has promised both. And I think he will succeed, albeit I think he's going to prang a few rockets in the process. If he doesn't even I'll be impressed.

Right now he's testing his "grasshopper" because I think the hardest problem to solve is simply getting the thing to fly downwards and land itself. There is the related problem that Musk has, and that's getting rocket engines to operate over and over again, reliably, and in a way that is more like regular aircraft, with relatively few people on the ground doing the maintenance. Of all the things he has to do in order to get re-usability, I think the machinery is the least worst of his problems. And by the way, he's trying to routinely re-fly rocket machinery involving turbo pumps and the works. If you wee designing something that can be re-fired at Mars I doubt you'd go for that level of complexity.

So to put it bluntly, verifying that your hardware is re-usable on Mars means a visual examination, a systems check and a test fire. On Earth you may want to do other tests and measurements but then you're also probably dealing with much more complicated machinery.

Now, all of this is presumptive Josh. It all presumes something as if it exists now. We don't know. But we probably will in a time frame short enough that when it comes to someone actually flying to Mars, these capabilities probably will exist.

Another thing. If we get to the point where it is relatively cheap to send fuel into low Earth orbit, that is a game changer. A thousand tonnes of fuel yes? A billion dollars. Still a fraction of the overall development cost. Lets go in style. Lets not even worry about ISPP and instead send as much fuel as needed into orbit and simply brute force the problem. I really don't mind. And I'd like to cover this more later. Of course, its one thing to have cheap fuel but in general its not ok to waste it. But throwing more fuel at safety is a good idea in my book. In short you'd probably end up with an architecture that simply parks a bunch of fully fueled (methane/LOX) propulsion modules in Mars orbit and just use them as needed. Anyhow, that's going down another rabbit hole.

Now back to Mars landing/ascent vehicles. I'm not asking a vehicle to be re-usable dozens of times, I'm asking it to be re-usable at most a dozen times, and probably less than that if your basis of comparison is 3 missions. In a future where Mr. Musk does succeed in re-using his rockets dozens of times (and they will subjected to all kinds of stresses) we will probably know if my guess is right or not.

If not, you can blow me a rasberry

"Your plan is to send two personnel ascent vehicles and two Fuel Ascent Vehicles to Martian orbit and leave one set there as a safety factor for every mission. Fully fueled? Or would the crew have to wait to refuel it, miss the launch window, and then return two years later? With what supplies will they last? How do you plan to store the hydrogen (yes, Hydrogen... Why are you recommending bringing methane? Insanity)"

Let me clarify this.

There are permutations but you can choose between the two basic forms.

One form is a landing/ascent vehicle that runs on CO/LOX sized to deliver oxygen to the transit vehicle and capable of docking with a manned capsule. In this form we start with two such vehicles. One remains as a fuel ferry and the other remains "permanently" docked to the capsule. In theory you could remove the capsule from one and transfer it to the other but yes that's logistically a last-ditch procedure. One advantage of this configuration is that the capsule would be capable of landing in its own right. So you have the capability of surviving failure of the ascent.

Correct me if I'm wrong, but does Mars direct have this capability or do they figure "it just has to fly" ?

Another form is more specialized vehicles optimized for scale. One is crew specialized. The other is pure fuel ferry. Benefit being commonality of design if not structure. You would have to consider how to make the crew version more robust in terms of failures on ascent but it could be made more survivable than an "all or nothing" alternative. In this case you could go belts and braces and simply deliver a second crew vehicle. That's probably overkill but the extra mass is small compared to the overall mission.

In any case its not 2 of each.

As to the next question. You always have a second fueled vehicle capable of crewed ascent into orbit ready before your launch window. And just to be specific I'll break it down for you into the possible cases.

Case 1: You've got yourself two nearly identical landing/ascent vehicles. One is your fuel ferry and the other is notionally your crew vehicle. Before the first manned launch you've arrived at a situation where both of these vehicles are fully fueled and the orbiting transit vehicle is also fully fueled - time is on your side here. The crew arrive, do their thing on Mars, and also check out the crewed ascent vehicle. It is found to be flawed. What do you do? Ok, that depends on the nature of the failure. Could be anything from avionics - in which case you cannibalize your fuel ferry (or more likely you have some spares. It could be that it has a major structural weakness. The last-ditch resort in this case would be to use the resources at your disposal. And this probably means first of all arranging a platform for the capsule to be undocked onto. Then you command the fuel ferry to make a short flight, unmanned, so that it parks itself over the capsule. Then using simple technology you remount the capsule to the fuel ferry. Now, that's pretty out-there, but its a capability you don't have with other architectures.

Case 2: You've got yourself two specialized landing ascent vehicles. No more. In this case you've got one crewed vehicle only. In this case before the first manned launch that vehicle is fully fueled and tested (flown a short distance). When the crew arrives you again have the opportunity for further systems checks and inspection. Again, in some cases you have the opportunity for cannibalization. But in this case you focus more heavily on spare parts and even some limited repair capabilities. Yes of course the same things are available in most cases in other architectures, but the nice thing here is you have a relatively light vehicle and its not quite so costly to test fly it.

There are orbit and trajectory options that give you a few days if not a couple of weeks to recover. Most likely what you'd want to do is to progress to a highly eccentric orbit a week out of your optimal launch window and from there you have similar options that you have from the surface.

Case 3: As above but now you've brought a spare crewed vehicle. My preferred option here is to have the spare vehicle fully fueled before the first manned launch from Earth. Depending on your risk analysis it might be ok to defer fueling the second vehicle whilst the crew is on the surface.

In option 1 an abort from ascent is relatively straightforward and you could recover.
In option 3 an abort from ascent depends on the nature of the problem and the robustness of your design. But assuming you can abort recovery is very straight forward.
In option 2 an abort from ascent an abort from ascent means you've a limited time to fix the problem (and transfer fuel) or else you're on short rations. That's why I'd avoid it, but its here for completeness. In short my original idea (option 1) has its attractions - not the least of which is every vehicle is identical.

Its not clear what your reference to hydrogen entails.

To clarify the landing/ascent vehicles are all CO/LOX powered. That entails on their very first landing there has to be some fuel that is shipped from Earth. Not a big deal. After that its over to ISPP.

The transit vehicle is methane/LOX fueled. The methane always comes from Earth. But the bulk of the mass (oxygen) is supplied from Mars. Its precisely because of the hydrogen storage problem that I've arrived at these two fuels.

As for the last question. If I were attempting to save 10 tonnes I'd say it aint worth it. Obviously I'm attempting to save hundreds of tonnes of mass (LEO). That occurs over 3 missions. I'm using that baseline because firstly NASA is targeting that. Secondly *anyone* would be crazy not to think in terms of at least 3 missions, all costs considered, and thirdly I think 3 missions is about right in that time and lessons learned will eventually make whatever technology you started with redundant.

Obviously cheaper fuel into low Earth orbit is a game changer here, but for us to make any progress on this we need an agreed basis of comparison. What masses we're actually sending. How many missions etc. Then we can look into both re-usability and overall costs.

Anyhow I'll throw this back to you.

Several things bug me about Mars direct and other related missions.

Firstly, there is a limit to how well you can check out the ERV (or its equivalent). Its got to work and its not clear to me what happens if faults develop that you simply cannot test for - especially if you have no option to test-fire, either because of the sheer scale, or because what you've designed has limited cycles built into it be design. There is a reason why Earth ascent vehicles have abort capability. They designed it out of the space shuttle and that wasn't such a good idea, and I'd be loathed to do without it on Mars.

Secondly, I really don't like options where you land people inside something that weighs several tens of tonnes. The reason is simple. Given that what fixes the ballistics is the size you can launch from Earth, something that weighs under 10 tonnes is probably going to get you to around Mach 3 at a much higher altitude than something weighing tens of tonnes - even if you resort to landing belly first (multi-conic) NASA has at least given that some thought. Mars direct are still squibbing it a bit.

In the end given the choice, I'd rather land in a simple capsule with more heat shield area relative to mass, than in a much bigger structure that ultimately has to push the limits harder.

I can handle aerocapture of a crew inside a "space-hab" or "transit-vehicle" more easily than I can do that

Thirdly. The criteria of minimal launches from Earth is something that gets people into all kinds of strife. We either end up with brand new super-heavy launchers or we end up cutting corners on various systems. the first is ugly and expensive. The latter is taking all kinds of risks. I'd rather both NASA and the private groups put a whole lot more thought into this.

Fourthly. I've avoided producing methane on Mars. For the simple reason that if you want something complex to fail in unexpected ways, its a chemistry lab operating on another planet. For this reason, I'm not all *that* keen on simply mining oxygen either - and I'll get to that later -. But if you *have* to rely on Martian resources, its oxygen that is the heavy bit. Both Mars Direct, and NASA, and others have underestimated in my view the amount of time it will take to demonstrate the technology and I do believe that it should be demonstrated, with people present but where its not essential to the mission.

Like I said before.. cheap fuel... reusability.. and ISPP. All of these have yet to be developed and tested. So what we're doing now is playing what-ifs. Allow me some latitude since I also extend that to others who also plan things based on stuff that has yet to be proven.

Cheers

May not be here till after the New Year hangover.. have fun

Russel · 2012-12-27 08:46:09

Here's a footnote.

Just reading through DRM 5 http://www.nasa.gov/pdf/373665main_NASA-SP-2009-566.pdf

NASA want to land two payloads onto the Mars surface for each mission. One is the habitat. The other is the crew descent/ascent vehicle. Each has a landed mass of about 40 tonnes.

Here's their figures.
Entry mass 109.7 tonnes
Aeroshell structure 22.5 tonnes
TPS 18.2 tonnes

Even with the assumption of aerocapture they're apparently doing this from a 1 sol orbit, so a fair bit more energy to burn than slipping out of a low orbit.

Even so, look at their other table..

Ballistic coefficient 471Kg/m2
Altitude engine initiation 1,350m
Mach @ engine initiation 2.29

Ouch, ouch, ouch.. and then more ouch.

Now, what are the corresponding numbers for Mars direct?

Here's another interesting tidbit. Part of the reason the crew descent is so horribly heavy and doesn't go into final deceleration until you can spot the the wheel tracks of the rovers is that they throw in a bunch of supplies and other gear with the lander that carries the crew. They had to do this of course because the lander that carried the habitat was itself at max weight.

These are the sorts of decisions that detract from crew safety. And this is why I recoil from such designs. And the non NASA missions don't do much better, or simply avoid any serious EDL calculations.

All things considered a specialized crew lander is a better option. Whether that's going to be one and the same as an ascent vehicle I'm not absolutely sure yet.

Oh and btw, NASA does note that they have to throw in extra descent fuel in order to pull a maneuver that discards that 40 tonnes of aeroshell and heat shield hopefully "so
that the heatshield debris does not impact the surface near any highly valued pre-deployed assets"

One of these days Mars is going to need a garbage mission...

Last edited by Russel (2012-12-27 08:51:14)

Russel · 2012-12-27 09:08:21

A brief muse about the implications of cheap launches of expendable items. Fuel, tanks, engines.

What do you get? Well no ISPP. You just land a big fuel tank. Aerocapture? Perhaps but money isn't everything and a fully propulsive orbit insertion is unarguably safest.

One off Mars lander? Yep.
One off Mars ascent vehicle Yep.

Space habitat? I'd argue there that the sheer amount of $/Kg that goes into one of those might make a propulsive capture into Earth orbit worthwhile.

What are we left with? The usual stuff that has to get to Mars one way anyhow. And an Earth return capsule that is less of a hassle to design.

Rough estimate - close to 2,000 tonnes into LEO.

Launch cost $2B-$3B

Still considerably cheaper than all the other costs involved - especially the engineering that will go into the Mars surface stuff.

Puts things into perspective don't it?

Oh and btw.. with all the stuff that's just thrown away, some of which into orbits that criss cross Earth orbit, one of these days we're going to regret that

Have a happy new year!

JoshNH4H · 2012-12-28 14:56:04

Quite a post there. Again, to summarize the points to which I'm responding, to make sure there are no understandings:

You suggest that because we don't have the capacity to launch a Mars mission tomorrow (next launch window, whatever) any Mars mission will involve new technology and development. You further stipulate that by the time we launch a Mars mission, $1,000-$1,500/kg launch and multiply reusable rockets will simply "be there."

You then voice several complaints you have with Mars Direct. I would like to get this on the record now: I am not saying that Mars Direct is necessarily the best way to send a mission to the planet Mars. What I am saying is that it is a reasonably cheap, effective proposal to begin the exploration of the planet and it is a logical baseline against which other missions can be compared. Specifically, the technology of the Mars Direct mission involves minimal advances compared to that which exists today, and I would say that technological advances neither stipulated by Zubrin or existing in the current day should be proffered with necessary justification.

In your following post, you bring up DRM-5. Let me be clear: DRM-5 is useless and irrelevant, and will not remotely approximate whatever the architecture finally used to get to Mars is. The Design Reference Missions (DRM) were based on the idea of "Mars Semi-Direct" Where In-Situ Propellant Production is not a mission critical technology, due to NASA concerns. DRM-5 bears no relation to Mars Semi-Direct and deserves no consideration as a practical Mars Mission.

Why will $1,000 to $1,500 Earth to Orbit just be there? You have no reason to say you can count on it being developed. Because it's a publicly traded company, you can't even say that SpaceX won't declare bankruptcy tomorrow. I don't think they will; but if you look at the rockets that they have built thus far there isn't really any new technology there. All they did is took the technology that has been there since the 60s and did it cheaper. I'm not suggesting that it's impossible for them to develop the technology or bring the price down that far, but it would certainly be an achievement above and beyond what they've already done.

Again, baselining to Mars Direct: The new technologies that would be required are Entry, Descent, and Landing for large payloads, as well as a new HLLV (High Lift Launch Vehicle) with a payload above 100 tonnes, plus you have to build the equipment to send to space. While the equipment sent into space will generally be of new design, the principles are the same and it can be tested pretty thoroughly on Earth. In-Situ Propellant Production (ISPP) while as yet unused in space, was already demonstrated by Zubrin... About twenty years ago now. Finally, Mars Direct stipulates centrifugal gravity for the trip to Mars. While this will in my opinion be comically easy to demonstrate (launch a rocket with payload; bring three or four tethers; use attitude thrusters to spin system), plus very useful insofar as you could vary the spin rate or the radius to simulate different gravity levels and do useful biological experiments on animals or people.

So: One complicated new technology, one simple, one new rocket, and an ISPP system that , plus old technology applied to new uses.

I think the comparison to what you're saying makes itself. Technology development costs money and increases the unreliability of the mission; We have sixty years' experience with disposable rockets, how many with your reusable ones?

I have one word for your reusable rockets and your reusable Entry Descent Landing technology: Vaporware. It's not hardware because it doesn't exist now and just because you incorporate it into the mission with certain specifications doesn't mean that it will actually have those.

With regards to abort during launch for Mars Direct: It depends on the ERV design. There is no fundamental reason why it couldn't be incorporated and it would seem to be a good idea.

I would like to remind you: In your plan, the Fuel Ascent Vehicle is just as vital as the Personnel Ascent Vehicle. If you're going to send two Personnel Ascent Vehicles (PAVs) but only one Fuel Ascent Vehicle (FAV), there is still plenty of risk. Especially if you're talking about reuse. Your comparison of how a nonreusable Earth Return Vehicle (ERV; As in Mars Direct) could be inspected before launch to how your proposed launchers could be checked is ludicrous: You suggest that equal levels of inspection (which is to say, none for both because there is very little checking that an any six-person crew can do beyond that done by onboard sensors) result in different confidence levels in the performance of the different rockets. For the sake of argument, the technology in the ERV and FAV/PAV is more or less the same.

Hydrogen can be stored, there is simply the question of boiloff. A refrigeration unit and a sunshade, in combination with Hydrogen's relatively large thermal inertia, can reduce boiloff to fairly low levels.

But given that Carbon monoxide/oxygen fuel seems to promise more-or-less equivalent performance, why bring Hydrogen at all?

RobS · 2012-12-28 17:07:53

I haven't followed this thread, so maybe this has already been dealt with, but the reasons Zubrin preferred methane/oxygen over carbon monoxide/oxygen are two fold: (1) the process of converting CO2 to CO and O2 requires high temperatures and is a more difficult chemical process to manage and sustain than the Sabatier process; and (2) an oxygen/carbon monoxide rocket burns at a much higher temperature than other propellants, so is much harder to develop and test, plus we lack the years of experience with it that we have for other propellants. For these reasons, and the higher Isp, Zubrin preferred hydrogen from Earth and the sabatier process for converting it into methane.

JoshNH4H · 2012-12-30 07:54:03

Those are very legitimate points. However, high temperature for Carbon Dioxide electrolysis is not necessary problematic unless it leads to increased unreliability relative to the sabatier process. Therefore the real question is whether it's easier to design a rocket engine capable of firing on CO/LOX with reasonable efficiency or design a Hydrogen storage system with low mass and low boiloff. Put in those terms it seems likely that Methlox is a better choice, but I think Carbon Monoxide also worth considering as a low Isp, higher density fuel that is very easy to make with ISPP.

GW Johnson · 2012-12-30 12:47:21

In the paper I presented at the Dallas convention two years ago, I posited a transport vehicle from Earth to Mars and back, to be based in low orbit at each end of the voyage. Some or all of the tankage has to be jettisoned, depending upon selected transfer propulsion, but the engines and habitat module are recovered and reused on subsequent missions. That helps cut down costs.

Based from orbit like that, it is possible to send a single vehicle to multiple sites on Mars in the single mission, depending upon how many landers you send, and exactly how you send them there. That increases "bang-for-the-buck". Sending a single landing vehicle to any given site presents the “standard” risks that we already well-understand from Apollo. Given sufficiently powerful propulsion, these landers can be single stage reusable, even without refueling while on the surface. Otherwise, these are two-stage one-shot chemical vehicles, unless you can refuel them on the surface. Anyone can prove that by plugging in realistic numbers into the rocket equation.

If you add refueling while on the surface, so as to make single-stage reusable chemical propulsion feasible, there are two choices: (1) carry the fuel-making equipment with you, or (2) send it down separately to the same site. If you carry it with you, there are two issues to address: (1) your lander is necessarily much bigger and heavier, and (2) the fuel-making devices must work very fast, within the time frame of the surface stay, which is limited by the men, for any of a variety of very good reasons. (Long surface stays are not very realistic for a first mission, due to all the life support uncertainties.)

Plus, you are betting lives on the fuel-making gear working correctly, at that particular site, which might be quite different from “typical” Mars. Although, that last risk can be effectively eliminated by suitable development testing, which of course takes calendar time.

If you send it (the fuel-maker) down separately, that opens a whole host of other safety issues that I have not yet seen discussed very well. The most obvious one is the capability of actually landing multiple vehicles close together at the same site, not too far out of range of each other. This takes a radar transponder and a vehicle that is steerable during entry. These are things we already have (even capsules have been steerable since Gemini), but we have never actually carried out such a homed-in landing before. That’s another issue that can be effectively eliminated by suitable development testing, which again takes calendar time.

The other issues involve the achieved range between landed vehicles. If the return vehicle is too far from the fuel-maker, how does one transport the fuel from one to the other, when there is no fuel transportation infrastructure on Mars? By truck? By pipeline? By strung hoses? That last requires a very close range indeed between landed vehicles. The other two require equipment that raises lander vehicle size considerably; if you do that, you might as well carry the ascent propellant down with you.

Landing really close together (so that strung hoses are feasible) brings into play another very serious risk: rocket blast effects. Even a chemical rocket produces a very high velocity stream whose stagnation temperature is very high. These are very destructive plumes, and the forces they impose on impacted structures are very high (in effect the same size as the thrust force produced on the vehicle). You run the risks of puncturing the propellant tanks on your fuel-maker, and/or cooking-off the propellants with the heating of the jet blast washing all over it (that plume spreads widely at low backpressures).

BTW, the supersonic expansion that reduces gas temperatures is not “permanent”: as soon as the gas flow shocks down subsonic, its temperature is very nearly stagnation, and that’s the rocket chamber temperature. That’s what happens as soon as the plume strikes anything solid. The source temperature for heat transfer across the boundary layer is the recovery temperature, which is only a little lower than stagnation.

The other risk with close-range landings is the obvious collision risk. That can be handled by a human pilot taking manual control, as it was on the moon with Apollo. But, you have to budget descent propellant to handle that contingency. You cannot trim margins “to the bone” and still do that effectively. Minimalist mission plans never address things like this. That’s one big reason why I rarely believe the claims of practicality regarding anybody’s minimalist mission design approaches.

What I proposed in my Dallas convention paper was powering single-stage reusable landers with solid-core nuclear thermal engines (basically a resurrected NERVA), and avoiding entirely the making-return-propellant issue, by simply carrying it with you in a bigger, more capable vehicle. These same nuclear landers could push the entire landing propellant supply to Mars, separately from the manned ship.

Resurrecting NERVA for this purpose might well be a faster development time than any of the in-situ propellant technologies. I think NERVA could be resurrected by the “right team” in 5 years. I’d bet any of the in-situ propellant things will take longer. Fundamentally, it all boils down to how soon do we really want to go?

That brings up a very serious c caveat: I think the US government goal of “sometime in the 2030’s” is really code for “never”. Long development times mean this landing will not happen in our lifetimes, if it is to be done by them. If we humans are going to do this any time sooner, it has to be with tinkertoys we already have or can obtain very quickly. And it needs to be done by somebody non-governmental like a Spacex. Somebody actually motivated to go, and free enough of bureaucratic chains to go.

Pessimistic, I know. Sorry, but I’m a practical realist.

GW

Russel · 2013-01-05 03:01:12

Josh,

I think Zubrin also proposed the idea of a mixed CO and methane engine which would to some extent preserve the Isp of methane while keeping the density advantage of CO. I saw it somewhere but I can't find the link right now. I think he did this in part as an answer to the purity issues that arise from the methane synthesis, plus wanting to limit the amount of H2 needed. Considering that H2 can be indefinitely stored in the form of H2O I'm not giving up entirely on a methane mixture.

GW,

Regarding a nuclear lander I'd love to see actual numbers regarding the shielding problem. All I have to cross check against is DRA 5 where they placed the engine a fair distance from the crew and on the crewed version also added 10 tonnes of shielding. Were they being over the top? Or is it a different problem because of the duration of burn? Just want to see the numbers there.

There does come a point where the mass of shielding and the density of hydrogen and its storage issues tends to take the fun out of things.

Speaking of lander/ascent vehicles in general. Has anyone given any thought to the absolute minimum mass you really need? Space suits and seats?

Here's another question. What if you really did consider a fully propulsive landing. Presumably you'd need to keep the speed at any point within the limits of very basic thermal protection. But the worst case is you slow yourself to zero at 150Km altitude. If you did that and went into free fall you'd potentially reach 1100m/s by the time you reached the surface. But that assumes no air resistance so in reality it would be less. And it assumes no parachute - which in this case would be perfectly feasible and would probably cut the final delta v in half. So if you budgeted for 5Km/s you'd be in the right ballpark.

All of that of course only makes sense if you go as light as possible.

GW Johnson · 2013-01-05 12:32:48

I'm not a nuclear rocket engineer, or even a plain reactor engineer. I don't have hard numbers to throw out. But here are some thoughts about reactor radiation safety, coming from an old mechanical/aeronautical engineer with some knowledge of the subject coming from wide reading.

Using solid-core nuclear rockets as propulsion in anything resembling a safe manner is not a trivial issue, to be sure. I do think the applications of orbit-orbit transport, and planetary landing, end up addressing this risk entirely differently, though.

For the orbit-orbit transport, issues of artificial gravity should interact constructively with the need to provide radiation shielding from your reactor. There is a need to stage-off emptied propellant tanks after every “burn”, but if the design comprises a set of docked modules, it is easy to reconfigure into the same length “slender baton” shape at each stage-off. Spin the “baton” end-over-end for gravity: 56 m radius at 4 rpm is 1 full gee at a tolerable spin rate. (This is true even for chemically-powered designs.)

This reconfigurable “slender baton” shape not only maintains radius for artificial gravity at low rpm, it also maintains the much longer distance that is so very necessary for getting shielding benefits out of your remaining propellant tanks. Somewhere around 40 meters of propellant tank fluids and structures should be quite effective at shielding the crew from nuclear radiation, during or between “burns”.

The lander is a vastly different proposition. Compact as it has to be for landing stability, shielding “steady state” by distance with tanks and fluids is impossible. The alternative is tons of lead or concrete, etc, also very undesirable. But since the descent and ascent “burns” are brief (minutes only), there is no need to shield “steady state”. Using what little tankage and structures shielding benefit that there is, the crew need only endure brief intense exposures, integrating to a very modest accumulated dose, actually. But this does require that the crew evacuate to a surface shelter remote from the lander during the surface stay. And you keep your distance from them in orbit, too, except when in use.

The intensity of the exposures might be mitigated slightly by a shift to thorium reactors instead of uranium. This gets the worst-offender plutonium-239 out of the picture. But, fission leaks neutrons, no matter what. I like the thorium approach better than uranium, in part because of the slightly-reduced danger from shut-down cores, but mostly because it is a more plentiful fuel. But, I recognize that we have to start with what we know: uranium.

Dangers of radiation from a contained core after engine shutdown could be mitigated greatly by the open-cycle gas core concept, which is essentially an “empty steel can” between “burns”. Only induced radioactivity is still a problem, and that is far less intense, and decays far quicker. That’s why I’m such a fan of developing the gas core technology ASAP, although we still have to start with what we know, that being HEU solid core.

GW

Russel · 2013-01-08 03:57:19

Has anyone worked on the idea of a rover specialized in refueling - perhaps also including the "lab" ?

RobS · 2013-01-08 07:17:49

I wish they would land something on Mars that would test the ISRU system of making methane from Hydrogen and CO2. But I suppose mass and energy limitations have been the problem. Maybe the Falcon Heavy will change that. Indeed, the Falcon Heavy is probably the key to launching a sample return mission, because it's the only launch vehicle with enough throw weight to send the entire sample return in one launch. As yo may have heard, NASA recently proposed a sample return missin that would require I think 3 launches, and Zubrin criticized it in print.

Russel · 2013-01-09 10:45:28

GW,

I'm wondering if you or anyone else who can do the math or the simulation can give me an answer to this.

Firstly, what order of lift to drag would be needed to keep a Mars lander high enough for long enough that one can keep the temperatures down to a modest level. I leave the definition of modest open there. It could mean low enough to be effectively non-ablative with conventional (but thinner) heat shield. Or it could mean low enough (a few hundred C) that you can afford to be relaxed about light weight materials.

Secondly, If we start from a low orbit and then a de-orbit burn, but then also just before interface apply increasingly more retro thrust, so your speed at interface is proportionally lower, what effects would this have in terms of peak heating and temperature?

As I said before, the extreme case is you simply hit the retros until you reach interface at zero, and then free fall the rest of the way. From 150Km up under Martian gravity, assuming no air and no parachute you'd reach 1100m/s by the time you hit the ground. Presumably then with ordinary ballistic drag you'd be travelling much slower. Question, what does the simulator say about this? Now, what if parachutes were used? Is it even necessary?

That's the extreme case. What I'm saying is there might be a happy optimum where you bring enough fuel for (say) 1 or 2Km/s delta v slowdown before interface and for this you're rewarded with a lighter heat shield, and more time to play with. Clearly that first 1 or 2 Km/s would have a huge effect on the total energy. And I presume also on the temperature.

Would love it if I had some numbers to play with.

Now combining the above two issues. Suppose I came up with a lander that was a bit more like a delta winged affair (it might prove totally unnecessary btw) and the said lander was able to use thrust at all stages both for stability and to control its angle and thus lift. Imagine for the moment that weight is distributed more towards the tail which is where the bulk of the thrust is, but there's also some nose thrusters.

So, from the deorbit you approach interface. As you do you apply some measure of retro thrust, losing (I don't know) 0.5 - 2Km/s. As you head into interface you use a smaller amount of thrust to optimise lift, mostly from the nose. You keep yourself as high as possible for as long as possible, losing energy as slow as possible. Then the inevitable loss of altitude as you cruise on past (again I don;t know) maybe 2Km/s. But you're still keeping the temps way down.

Finally you do a shuttle-spiral and with a bit of technique and good code you land on your tail. Well maybe. It might (just) be possible to land vertically on your belly. Whether its possible to take off (fully fueled) like that, is another matter.

Anyhow, if anyone can venture some numbers, it'd be appreciated

GW Johnson · 2013-01-10 20:26:41

Russel:

I don't have the tools to do all of what you describe. I did look at deorbit vs angle at entry (in a crude but decent pencil-and-paper way) by just calculating derivatives (slopes) off the elliptical vacuum path, evaluated at the "entry interface" altitude. The lower your orbit from which you deorbit, and the minimum deorbit delta-vee you think you can get away with, then the shallower your flight path angle at interface.

My numbers were picked in ignorance, but as educated guesses. I picked 200 km altitude from a circular orbit. I picked a surface-grazing transfer ellipse, which requires a 50 m/s deorbit delta-vee. At the "entry interface altitude" of 135 km as recommended by Justus and Braun, I got 1.6 degree depression angle of the flight path below horizontal. That really shallow angle is quite safe (no bounce-off) because all velocities are well below escape. And it gets you survivable end-of-hypersonics altitudes, even with ballistic capsules of very high ballistic coefficient. Best of all, you get this inherently, it is not a matter of careful precision control.

My entry model is a simple pencil-and-paper model that they used about 1956 to estimate non-lifting warhead reentry. In effect, figured as 2-D Cartesian (even though it's definitely not), the flight path that the warheads actually flew is very nearly constant flight path angle. The model equations give you a velocity vs altitude profile based upon a density scale height approximation, and a constant ballistic coefficient approximation. The model was given in the same Justus and Braun report I cited in my postings over on "exrocketman". The dynamics were pretty good, but the convection model as presented was pretty bad. I fixed that, and posted what I did.

That's what I've been using for the Mars lander problems. I ran some trade studies, then investigated a couple of ballistic coefficient extrapolations to high masses, before I was satisfied enough to run some lander designs. All this stuff is posted in a bunch of "exrocketman" postings.

I don't have any way to run a lifting model. But qualitatively, it's the same as a ballistic capsule with extra forces applied. The idea of decelerating earlier higher up to limit skin temperatures is real. You could run my 1956 warhead model at unrealistically-low ballistic coefficients and see the same phenomenon. If the pulse of deceleration gees is earlier in the entry trajectory, the convective heat rates are lower in the thinner air. The heat integrals over time are about the same though.

It's the lower heating rates that reduce skin temperatures. You still have to have a sink to store all the absorbed heat, however. Details like the actual thermal gradients achievable with this or that material, are very tough, extremely-compressible, heat transfer analyses, best done with appropriate correlations and models. NASA and the big aerospace companies have the best ones. Not me. All I can do is get "into the ballpark".

Flying a winged vehicle gets you a lot more hypersonic L/D potential (on the order of 1) than a tilting-attitude capsule (on the order of 0.1). Lifting bodies are in-between. The place where lift capability seems to get you more benefit is later in the trajectory, after peak deceleration, where you can get some cross range capability the capsule cannot get you. That's what the NASA (and related) reports seem to indicate. On the other hand, a winged vehicle with a very low wing loading gets you that early deceleration, and technologically seems easier to actually build than a low-ballistic coefficient capsule (for which ablative inflatables seem to be the only potentially-feasible route).

No one has ever done it, but a steel truss frame covered by a ceramic fabric, perhaps even a straight-wing design, might be that very-low-wing-loading reentry vehicle. It would resemble how we built Piper Cubs in the 1920's, adapted to a silica-fiber ceramic fabric covering. There are lots of technological problems, and I'm not sure I know what they all are, but I haven't come up with any show-stoppers yet. I do know that the skin has to be stood-off from the steel frame by low-conductivity stand-offs that are not ablatives (or you will have to recover the airframe for every flight).

I also know the interior must be fed with a sacrificial coolant flow (most likely steam) to absorb the heat absorbed into the vehicle, and to prevent hot gas through-flow into the interior, because the fabric is porous. Somewhere around 25-50% porosity, depending upon the weave. And structural strength (as always) is the killer. There cannot be high L/D early in the entry without tearing the wings off; plus a low wing-loading vehicle is inherently more fragile than a high wing-loading vehicle (more mass is more material available to be stronger).

A small model test article, perhaps thrown out from a de-orbiting capsule, might be a good final proof-of-concept test. I find this concept very intriguing. It might lead to a very simple, very reusable, and very inexpensive design for reentry vehicles, including really nice cross-range and subsonic flight characteristics.

GW

Last edited by GW Johnson (2013-01-10 20:29:03)

GW Johnson · 2013-01-10 21:48:58

Russel:

I took a quick look at the idea you mentioned (kill the orbital velocity while at orbit altitude, and free fall vertically). For a 200 km altitude circular orbit about Mars, the orbital velocity is 3455 m/s. That would be your delta-vee required to initiate the vertical fall.

A ballpark estimate of velocity at the surface would be a vertical fall at Mars gee constant through that same 200 km. That’s crude, because gravity isn’t constant, but it is ballpark since the drop is small compared to a Mars radius. I get 1220 m/s. You said about 1100, so I know we’re both in the ballpark.

There will be some retardation due to atmospheric drag as you approach the Martian surface, but not a lot, since the air is really thin there. So, you have to rocket-brake to kill the bulk of that terminal velocity. Just for a number, let’s say that delta vee will be somewhere around 1000 m/s. The total delta-vee required of the descent propulsion is then about 4455 m/s, for that scenario.

On the other hand, you could do a near-minimum deorbit burn, and see how much aerobraking you can get, which is the conventional reentry approach. From the same circular 200 km orbit, the surface-grazing transfer ellipse has an apoapsis velocity of 3405 m/s, thus requiring only a 50 m/s deorbit burn to put you on that trajectory.

As you descend along the transfer orbit, your velocity increases (conservation of energy). At the interface altitude of 135 km, I’m calculating 3469 m/s for the actual entry velocity, vectored 1.6 degrees below horizontal. And that’s the end of what standard vacuum orbital mechanics can tell you, because you are now inside the atmosphere. From there, I used my crude warhead entry approximation.

The final ballistic coefficient that I guessed for a 60 ton lander was 400 kg/sq.m. That plus 1.6 degrees trajectory depression below horizontal, and 3469 m/s at 135 km, together lead quickly to local Mach 3 end-of-hypersonics at around 700 m/s and 17 km altitude, still near 1.6 degree depression. That's too low to bother with a chute or ballute; you are only a minute or so from impact. That 700 m/s is the minimum velocity to be killed by rocket braking, but there’s also still a controlled descent to make to touchdown after you kill it. So, at least double it to somewhere around 1400 m/s. The min total delta vee for this kind of descent is thus somewhere around 1500 m/s.

Compare delta-vees: somewhere around 4400 to 4500 m/s for the velocity-kill / vertical drop scenario, versus somewhere in the ballpark of 1500 m/s for the conventional aero-braking entry. Shallow-entry aerobraking seems to have about half an order of magnitude advantage in reduced delta-vee requirements over the vertical-fall scenario. That same kind of effect is why all the capsules and spaceplanes we have ever flown here on Earth used shallow entry aerobraking. It’ll be no different on Mars, just the smaller numbers, similar to those we are contemplating here.

GW

Russel · 2013-01-13 00:17:38

Just a quick response for the moment and I'll go and do a bit more research.

The thought behind losing some extra energy before interface is to achieve one or two things.

One is to reduce temperature to the point where you're not relying on ablation. Which can amount to a large amount of mass on a large heat shield. So you can trade off mass in the heat shield for mass in fuel. The hope is that in so doing (and you burn off much of the fuel mass before interface) you buy yourself a lot more margin further down.

The other is to reduce temperature to the point where you can rely upon lighter materials.

So I wasn't really thinking about slowing to zero. Rather I was thinking of losing some velocity where its most fuel efficient to do so. Say, from 3.5Km/s down to 2.5Km/s. That's half the energy gone. What that does to temperature is where I need to understand more.

Regarding structural loads. The idea is to be able to have some fine control of attitude using some thrusting during hypersonics. Maximum lift as you enter the atmosphere. Then less lift but lower structural loads further down. I may have to expand on that later.

louis · 2013-01-13 10:29:13

What is the Grasshopper intended for? Does anyone know?

http://www.spacex.com/updates.php

GW Johnson · 2013-01-13 10:51:11

Hi Russel:

Have you seen the deceleration dynamics profiles I posted over at "exrocketman"? For shallow entry trajectories, velocity vs slant range (or altitude, take your pick) is an S-shaped curve transitioning from high to relatively low. Acceleration (decel, actually) is a pulse (or narrow bell curve) situated in between two zones of relatively flat and low deceleration. Initially, the V-squared effect is high but density is incredibly low, while finally V-squared is relatively low while density is significant. In between, you get high forces acting to decelerate the vehicle. With low ballistic coefficient, this pulse occurs earlier in the trajectory, at high coefficient, later. The same thing would apply to lift forces.

The convective heating rates follow the same pulsed behavior, and at pretty much the same place in the trajectory as the deceleration pulse. Earlier is a bit lower heating peak, later is a bit higher. Peak heating rates go with peak skin temperatures. But, the heating integral over time is pretty much the same no matter where in the trajectory the pulse takes place. Down to around 100 kg/sq.m on Mars from LMO, I don't see heat rates that might excuse me from needing some sort of significant and heavy heat shield. But if I got down closer to 1 or just a few kg/sq.m, I might see the opportunity to use the lighter options. That corresponds to the wing loading of the typical Piper Cub or other light aircraft, around 6 to 15-or-so pounds per sq. ft. No one has ever designed a reentry vehicle like that.

Steepening the trajectory moves you very quickly toward hitting the surface before aero drag can decelerate you to local Mach 3 (the semi-arbitrary "end-of-hypersonics" point). From 200 km on Mars, I ran one scenario at an intermediate ballistic coefficient, and at 6-something degrees depression instead of 1.6 degrees. I had surface impact between Mach 4 and 5. Same thing happens with warheads here: impact is at two-digit Mach at typical suborbital ICBM conditions.

I haven't run the vertical drop scenario with my reentry model. It starts from near-0 velocity, not 3469 m/s on Mars. But I don't think you see much drag in that thin air until too late to do much good. And remember, we have no chute designs that work much above Mach 2.5.

Your model of killing only part of the orbital velocity would fall somewhere between the two extremes I looked at. I don't know how things vary between those extremes, but I doubt it's directly proportional to velocity. KE, perhaps. Dunno.

Heat shields can be either ablatives or refractories, and can be augmented by sweat cooling or simple heat-sinking. There's a lot of options there, especially on Mars, where entry is less demanding because the KE's at interface are so much lower. Any technique we can make work here, can be made to work easier there. At very low wing loading (ballistic coefficient), it might be possible to do refractory with ceramic fabric instead of low-density tiles. But, skin temperatures must be under 2300 F with alumino-silicate fiber to avoid embrittlement and irreversible volume change from temperature-induced solid phase change. That's a big "if".

GW

Russel · 2013-01-14 08:31:03

GW,

Yes, seen the profiles. Still drilling down to figure out how I would go about coding the model you've used.

There appear to be some problems that can only be resolved by trying out different scenarios with the model. For the moment I've given up on exact temperatures but instead I'm looking for proportional changes to rate of heating. Total heat is a different beast and it all depends on getting the heat to go somewhere else. But reducing the overall flow has to be a good thing.

The biggest problem I can see is getting enough lift (not just drag) at high altitudes (above 60Km). In this regime what seems to happen is you get a lot of heating (due to velocity) before you get a lot of drag. Without enough drag, lift to drag won't help you much. But on the other hand, at high velocities the effective acceleration towards the ground is much lower (since you're still sub-orbital). So which effect(s) win out?

I can feel some code coming on.

There may still be a role for some modest level of thrust even after interface but my gut feeling says its subject to the law of conservation of difficulty. Might be just as easy to burn off some velocity before interface. Again, only the computer knows.

As I said before, the reason I like having structures that are separate to the main part of the landing craft is that they don't have to be treated as kindly as far as cooling goes. Indeed they could be sacrificial. Same goes for wings. And the more I look at the profiles above the more I realise that its very thin air and very high speeds is where the problem is.

Besides even modest (shuttle like) wings perform better at lower mach.

As for a "drag net" the reason I like it so much is that it can radiate a lot of heat - and do so away from the craft you're trying to protect.

Btw, one thought I keep getting pushed to is making the main craft relatively slender with as little drag as possible (at least until you reach lower altitudes) and instead rely upon auxiliaries (temporary wings or nets or ballutes). Not much more I can say without the numbers though.

GW Johnson · 2013-01-14 10:23:28

Hi Russel:

My bandaid / minimum-analysis approach really isn't adequate for what you want to investigate. For that you need either a 3-dof (lift-drag-pitch) or a 6-dof (lift-drag-sideforce-pitch-yaw-roll) trajectory code. The attitude controls you impose will dominate your trajectory results. And, that code cannot tell you when the airloads rip the wings off, or do the aeroheating. Those are separate analyses coupled to your trajectory results.

I suspect wings are of little use landing on Mars subsonically, due to the extremely thin air. But some sort of lifting surface might be of use hypersonically. However, gravity "drop" hypersonically is almost negligible, even here. The trajectory bends downward pretty much the same as the surface curves. That's why that idiotically-oversimplified 2-D Cartesian warhead entry analysis comes so close to predicting the actual path, even for pure ballistic objects with zero lift.

The pitfall with multi-body constellations is excessive aeroheating where shocks impinge on adjacent structures. That's really hard to avoid if the shape isn't totally simple. That's the effect that nearly cut the tail off the X-15 on the Mach 6.7 flight with the scramjet pod test article replacing the ventral fin. The pod nose shock impinged on the X-15's fuselage right under the engine. Cut it about halfway through, from the photos I've seen. Yet most of the rest of the airframe was still OK. That particular flight tested some sort of white coating that was intended to help protect the Inconel skins at speeds over Mach 6. They judged it a failure separately from the shock impingement problem.

GW

Russel · 2013-01-14 11:12:44

Might be of interest:

http://www.ssdl.gatech.edu/papers/confe … 3-0908.pdf

Russel · 2013-01-14 11:47:54

I think I need a bit more spare time but sometime I would like to do a model.

I don't think it needs too many degrees of freedom to reveal whether or not certain approaches are headed in the right direction.

Agreed that wings aren't much use in subsonics. The question I need to answer is when do I feel inclined to settle for the ugliness of disposable bits. True wings I suspect are headed that way.

In my calculator based numbers, gravity drop kicks in under 3Km/s and certainly around 2.5Km/s so there's still an open question as to whether you get enough lift first or gravity kicks in first.

I hear what you're saying about landers in multiple parts. As with all things engineering, if its ugly, heavy or expensive at least keep it small. In this case the structural elements that cross hot gas streams would have to have some pretty severe protection. Lots of ablative wrapper. Even sacrificial coolant. But its not impossible, just hard to calculate/test.

And when its all said and done, if all this complexity can be avoided by burning extra fuel, that's still the winning option.

Btw, any thoughts on how far you could get with water/steam (or even starting with ice) is your sacrificial coolant? How about even creating a little thrust (steam powered rocket, anyone?) in the right places. Now I did do the maths on this a while back and realised it was a non starter if you wanted to absorb all the energy of entry. But in the right places? Dunno.

Last edited by Russel (2013-01-14 11:51:24)

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