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RobertDyck:
I'm guessing you are talking about the plume capture facility to test such things on Earth. Your cavern idea is just part of such a thing. Yes, you must have a gigantic volume into which you store the entire accumulated plume from the test, but without upsetting the backpressure on the nozzle, I might add. So, you don't just fire the engine inside the space. You must capture the plume real-time at ambient pressure, and pump it real-time into your storage space. That sort of thing is horribly complicated and horribly expensive. Then you have to decontaminate all that gas of radioactivity, before you can let it go. That ain't cheap, either.
So for Earthly testing, you have a huge, horribly expensive facility, and every test is a bank-breaker. Plus all the neighbors will fear you and complain, whether justifiable or not. Eventually they'll shut you down, the way our politics-of-money has been working the last 40-some years.
On the moon, it's just an open thrust stand (steel on concrete, with piping and cables out of the crater over to the base proper, where all the tanks are). Very expensive to build a simple base and even simpler facility there, because of launch costs to the moon. But every test is essentially dirt-cheap in comparison, and there are no neighbors to complain. In the long run, almost regardless of the lifetime you expect out the base, this is cheaper, and politically far easier to maintain.
GW
Not much gimbal is needed for a near-center cluster, when running on 3 out of 4. You just gimbal the odd one to thrust through the center of gravity. For a near-center cluster location, that's not a big angle. For a periphery location, it's really big.
But, with 4 engines, and you lose one, just run on 2 diametrically opposed, instead of 3. Differential thrust provides in-plane vector. Attitude control thruster provide out-of-plane control. No gimballing required.
But, my landing boat was just a rough-out design sizing. I never worked out the numbers in detail for engine-out.
GW
For a ballistic suborbital flight, the delta-vee is about like that of an ICBM for halfway-around-the-world. That would be something like 5.3 km/s as you leave the sensible air, at something near a 45-degree path angle. That's uncorrected for gravity and drag losses. In comparison, the uncorrected orbit speed is near 7.7 km/s.
For skip glide, you only need about Mach 3 to 5 as you leave the sensible air, but you will need it a bunch of times to go halfway around the world. Your path angle will be quite shallow, and your apogee just barely exoatmospheric. This isn't launch rocket stuff, this is airplane-that-goes-ballistic for a few hundred miles a jump. Apples and oranges, doesn't compare on delta-vee. Sorry, I have no data for such a vehicle.
GW
Airship? Two serious troubles: (1) very difficult to get that high a service ceiling out of buoyant lift, the density ratio to SL std is about 0.37 or so; and (2) even worse, mountain regions have very high, very turbulent winds, which are usually fatal to airships even at SL.
GW
If takeoff from the high-altitude rescue site is too difficult because of the payload increase with all the rescued people, there is always the venerable old JATO bottle.
I'm not sure that high-mountain rescue will ever be a for-profit business, so return-on-investment may not apply. Most rescue services are operated by governments. I do think any such rescue craft will be a VTOL airplane or helicopter. There's likely to be only a tiny handful of staging bases to cover the entire range of the Himalayas, same for other mountain ranges elsewhere. The craft will need hundreds of miles worth of useful range, plus considerable on-site hover time. \
GW
Rocket MHD might never make sense as an electric power plant here on Earth. Simple fuel-fired boilers will likely continue to be less expensive and more reliable for power generation. They have been, for the last half a century. I first saw rocket MHD mentioned in a propulsion textbook from the 1960's. But, the idea might have real merit off-Earth, where an oxidizing atmosphere is unavailable to run the normal furnace and boiler rig. As far as I know, no one has looked at this in a very long time.
The notion that rockets excel at speed is quite correct. Personally, I am a disbeliever in high supersonic-to-hypersonic cruising aircraft in Earth's atmosphere. The heating and the drag are just overwhelming technically and financially. It's OK for missiles. But for manned aircraft, exoatmospheric skip-glide makes a lot more sense for suborbital travel. If this were to become some sort of civil travel similar to airlines, it would be fairly easy by communications to ensure that such craft are not mistaken for missile attacks.
Whether that or suborbital ballistic is more financially feasible is an open question. Both will be a lot more expensive than subsonic transport aircraft, but for some applications where time really is money, the much faster speed could pay off. Going exoatmospheric is how you avoid the heating and the drag over most of your trajectory. There is a hypersonic entry heating transient at every dip back into the atmosphere, but the point is, it's a short transient. The ascent heating is similar, but usually much less severe.
GW
Precision: hi RobertDyck. Awww, don't quibble. We're talking something still sci-fi here. (ha ha).
Testing: why not test such things on the moon? No need to build an exotic and expensive facility with plume capture capability. That should almost pay for building the base right there. Put the thrust stand down in a crater, to contain the debris from a failed test (and there will be failures, that's inevitable in rocket development testing). No air and water to pollute, no neighbors to bother. It's perfect.
GW
Thanks Quaoar. Nice to know my hunch was about right.
I think the steam NTR would be even more attractive, and practical, done as an open-cycle gas core design: somewhere between 1500 and 4000+ s Isp, depending on achieved T/W. Too bad nobody ever tested such a thing, other than a couple of academic bench tests of a couple of principles, about half a century ago.
GW
The problem with a rescue craft operating on Everest isn't doing VTOL at 29,000 ft, nor is it a suitable launch site. It is a suitable touchdown spot on the mountain close enough to do any good. That's the most hostile landscape on this planet: nothing but untrustworthy rock and ice at impossible angles. In comparison, building a helo or V-22-like craft capable of hover at 29,000 ft is easy.
GW
Why not seed the flow of a rocket engine with something like salt that ionizes easily, then capture the KE of the stream down to the recombination point with an MHD generator. Then shock-down the decelerated stream to subsonic, and put it through a heat exchanger to generate steam for a turbine-driven generator. That's two DC sources of two voltages, but we ought to be able to do something with that. Might be as easy as simple seawater injection.
As for rescue atop Mount Everest, the design of a VTOL aircraft or helo for operation there is simply a matter of design for low density. It just takes a bigger prop, wing, and engine for the load. Just because no one has yet done it, does not mean it is impossible. The density ratio to SL std at 29 kft is 0.3887. Harsh but not impossible. The real problem for any aerial rescue craft on Everest is suitable touchdown sites. There are none. Coming up with a way to glom onto ice and rock at impossible angles is the true design problem. In comparison, flying an aircraft there VTOL is quite easy. Just not cheap, not ever.
GW
It should have been well-recorded in the aerospace trade magazines of the time. I'm not sure Aviation Week goes back to the 50's. But there was a magazine like that back then. I think it was called "Western Aerospace" or something very similar.
GW
Tom's suggestion of rocket troops sounds fairly ridiculous today. RobertDyck correctly pointed out the reasons why.
Yet, surprisingly enough, in the late 1950's and early 1960's, exactly that rocket troop thing was one potential mission for a new giant ICBM development being done by the US Army. They were using the ex-German scientists from Operation Paperclip for this, so you know who the chief designer was. This started before there ever was a NASA, which was created out of the older NACA in 1958, and charged with a civil space program, separate from military efforts of various sorts that were ongoing in USAF, USN, and US Army.
What you may not know is that "giant ICBM" of the US Army was the Saturn family of rockets. Those were NOT originally designed to be space launch or moon rockets, contrary to popular belief today. The Apollo moon mission had to be adapted to the biggest rockets then potentially available. That's why NASA had to accept the outside-the-agency idea of lunar orbit rendezvous in order to get down to one Saturn-5 launch per moon mission. NASA has never accepted an outside-the-agency idea since, but now may have to, since Spacex is showing them up, so very adroitly.
The mid-50's concept of the rocket troop mission was a pre-emptive strike on the Soviet Union. Saturn-1's dispensing "bombs" of LSD-25 would put everybody out on their backs, seeing pretty colors for three days, followed by Saturn-5's as a 2-stage vehicle bringing over 100-man battle units in big parachute/rocket brake landing canisters (a one-way trip). The idea was to be standing there with bayonets at their throats when everybody woke up from the LSD trip. (Of course, this would have been a fleet of giant rockets that would have made the 1970's ICBM fleet look like a bathtub toy flotilla.)
I know this sounds like the plot to a very bad sci fi story, but it was very, very real in the late 1950's. This is the real genesis of the Saturn-5 moon rocket. It is what Werner von Braun was working on, before he got to do what he really wanted: go to the moon with giant rockets.
If there is anything to be learned from a story like that, it is that government agencies are not renowned for their common sense.
GW
Hi Quaoar:
On my landing boat design, I had 4 engines near the center, inside a room sealed to the backside of the heat shield structure. That design would only need to gimbal a little in the engine-out condition, where only two diametrically opposite are running. Or you could rely on attitude control thrusters for steering (as in the 1960's-vintage Scout launcher).
With all 4 engines running, differential thrust among the engines will provide steering about as well as gimballing would. For engines nearer the periphery, the differential thrust moment is even larger, but so is the disturbing moment when one engine quits, but before you can shut the diametrically-opposite one down to compensate. So, maybe 8 smaller engines is even better. I don't know, really.
As for off-center but axially-oriented engines having stable plumes, I don't know. I'd cant around 10 degrees in a first design, just to be sure. Cosine(10 deg) is a number fairly close to unity, so there's not any really significant loss of thrust and Isp at low cant like that. It'll take some flight experience with this, in more than one vehicle, to really know the art of doing it.
GW
Folks have made a big deal out of lighting a rocket engine whose nozzle faces into a supersonic or low-hypersonic slipstream. That might be an issue with Earth entry (not proven either way yet), but I doubt it's that much of an issue on Mars. That's because the densities and velocities are so much lower during entry there. That means dynamic and pitot ram pressures are also a lot lower.
The cant angle solves the reversing-stream stability issue that otherwise could potentially lead to attitude control loss, if your attitude thrusters are weak. Spacex's 45 degree cant for the (not weak at all) Super Dracos is driven by geometry, not plume stability. They are firing from pods mounted on the afterbody, not anywhere on the windward-facing heat shield. It takes about 45 degrees from a low afterbody position to clear the rounded corner of the heatshield without a hot plume strike on that heat shield.
There's no data to support or deny this, but my engineering intuition suggests that 10 to 15 degrees of cant would more than suffice to ensure plume stability for supersonic retropropulsion. That remains to be seen in testing, but could be verified in flight at about 100,000 feet right here at home.
The remaining question is how many engines and where in the heat shield do you put them? Opinions vary. I'd guess 4 to 8, nearer the center to restrict engine-out torque transients that might destabilize attitude. Others prefer a ring nearer the outer perimeter of the heat shield. My guess is that either, or anything in between, could be made to work. Something else will constrain that choice.
Do these need ports through the heat shield with doors that close when the engines are not firing? Again, opinions vary. I'd say that no, doors are not necessary, if-and-only-if you can reliably seal the engine compartment to prevent all throughflow. There is no better insulator than a static gas column, even against entry plasma.
All of this is easily tested for basic feasibility by models shot to 100,000 ft on balloons and/or sounding rockets. You can shoot higher and turn nose-down and thrust again to achieve very hypersonic speeds as you hit 100,000 feet. We first did that with the X-17 ca. 1960. It works fine. Viking's chutes were tested that way, too, in the early 1970's.
Once you have a design known to be feasible, you build it, stick it on a bigger rocket, and test it "for real" all at once, at that same 100,000 feet where the density matches lower-altitude densities on Mars.
This is all very doable. Any outfit serious about landing big things on Mars would already be working on these things. NASA is generally not very serious about it yet, but Spacex is. Their Dragon with the Super Dracos and some legs is already proposed for one-way unmanned payloads over a ton for Mars, including the sample return mission.
GW
News reports today show (1) successful launch, (2) successful injection of Dragon onto orbit, targeted for rendezvous with ISS Sunday, and (3) data received from 1st stage of Falcon-9R for 8 seconds after (after !!!!) landing in the Atlantic.
Congrats to Spacex!!! Very, very, very well done!!!
GW
I talked to a Spacex McGregor technician today at the local auto parts store. He told me the test I heard yesterday was their Grasshopper test vehicle flying with the production design composite landing leg assembly that is on the Falcon-9 at the Cape. That flight verified structural integrity of the flightweight legs for an actual touchdown on solid land, which is not what will happen on the upcoming launch. They're just trying to get a splashdown survivable enough to recover hardware for evaluation.
I saw some words somewhere that indicated there was something "special" about this Dragon, too. Although, I have been unable to locate those words again. The photos do not show anything I recognize as the Super Draco thruster assemblies, which bulge outward a lot more than the standard Draco thruster units. They've been testing those Super Dracos as separate items for some time now. I'd think they'd want to pre-test them on a flight they're getting paid to make anyway, but maybe the NASA contract specifies otherwise. I did not ask the tech about that.
I did ask him about the new thrust stand for Falcon-Heavy. He said the concrete, much (but not all) the steel, and the water supply system for the sound-deadening water curtain, are in place. They hope to test a Heavy on it by year's end, but there's a lot of thrust stand steel still to install. They gotta get it done soon, as USAF wanted them to launch a Heavy from Vandenburg sometime this year. I think that launch date will slip, although the web site did not reflect that, last time I looked several days ago.
GW
I quite agree with Louis: self-sufficiency isn't possible in the first mission or two, it will have to be developed over a long time.
My point is that there will be one (and only one) government-funded manned mission to Mars. That's just politics-of-money, which is the controlling factor for government efforts, as was so clearly demonstrated by the truncated Apollo program.
What that one-and-only government mission accomplishes has to be attractive enough to draw visionary private concerns into doing missions two-and-on. If you don't do that, expect at best a half-century hiatus, just like the moon. Or worse, we never go back.
Just landing a hab (however that is accomplished) won't be enough to do that job of placing something "attractive" on Mars. Just-a-hab is tantamount to an Apollo-style flag-and-footprints stunt. To do this "right", you have to try out every imaginable ISRU technique while you're there, and leave the best ones running on automatic when you go home. You have to set up multiple greenhouse approaches while you're there, and leave the best one "running" when you go home, to see if it really balances and self-sustains.
The equipment to do all of that is voluminous and it is heavy. You don't get that from landing one-to-three 10-ton hab modules shot direct from Earth. It'll take a whale of a lot more than that. That's why I think the total payload to be landed is nearer a minimum of 100-to-300 tons than it is 10-to-30 tons. You can land it in small chunks, of course, but I think it would be safest to do that with a reusable landing boat from Mars orbit (however it is fueled).
GW
It all depends upon what you are trying to accomplish. For one-way deliveries of habs and supplies, there's little need for anything over around 10 tones.
For a two-way reusable "Mars ferry" or landing boat, something needed when you delivering a permanent base or a real colony, vehicles like that start somewhere around 30 tons, and go up. It doesn't matter whether you fuel it from orbit or the surface, you still have a big velocity requirement for the flight up, plus a smaller one for the landing.
I think that the "minimalist" notions of just landing habs and supplies will lead to another Apollo-like problem. The government or consortium of government agencies that sends the first mission will not send another. It's the "been there and done that" excuse, used to cancel Apollo in the middle of the planned landings.
If you don't establish a first base (to be left running on automatic) on that very first trip, then it is unlikely that anyone will do that job for many decades, just like what happened with the moon. But if you do leave an operable base, a visionary private concern might go back sooner, and put it to good use.
Landing 10 ton habs and supply modules shot direct from Earth is probably not the best way to do that job. For one thing, there is the "we missed the aim point" problem (too far away is simply wasted). For another, there is the landing area collision problem: you might hit something already there, or damage it with your rocket blast. And, there is the difficulty of assembling the base from payloads you have to transport over rough ground.
I may be wrong, but it seems more sensible to me to use a bigger landing boat to make fewer trips down with bigger cargoes, from a bunch of stuff you sent to Mars orbit. Fewer trips with bigger cargoes reduces the probability of incurring the two problems cited above, and it reduces the effort needed to transport stuff over rough ground for base assembly.
That base needs to be more than just a hab. It needs to be a place to experiment (manned and unmanned) with all sorts of ISRU and crop greenhouse stuff. Otherwise, why go back?
There is a longer-term danger here if you think too small. It already hurt us with the moon.
GW
Oh, I quite agree with you, Tom, that we ought to stop the Hitler wannabees.
But, it helps if we ourselves behave in way that doesn't make us look just as bad. And the problem is, we have been doing exactly that for about 3 decades at least. That's why the rest of the world no longer thinks very well of us.
The last time the US truly had the moral high ground was WW2, and for just a few weeks after 9-11. Too many of our leaders have cared more about polls and elections than actually doing-the-right-thing. They have screwed it up and cost us our good reputation, little by little, year after year.
This is a big, complex, rapidly-industrializing world. The US is one of the big, complex, and powerful countries in it. Trouble with "big" is, after a while, you start to act like it. Like I said before, nations tend to behave exactly like 5-year-olds misbehaving on the playground. They all do, us included.
No matter what happens in Ukraine, the US is under no real threat from Russia, because they aren't a superpower anymore. Putin is driven in part by that. He's trying to recapture the superpower that was the Soviet Union, but I don't think he (or anyone else) can pull it off, at least not for a time best measured in halves of centuries. Some of the others in this conversation choose not to believe it, but he's weighing his chances right now, for invading part of eastern Ukraine. Not because he really needs or wants it, but because that's how you look like a superpower.
The real threat to the US (or most other countries) is from within, not without. What I find to be true is that very large organizations, government or private, tend to display a bureaucratic arrogance that is inversely proportional to the level of competence they display. What that says is that smaller is better behaved. But, the huge, complex world is demanding big. Big governments generally try to make economic slaves of their citizens, sooner or later. Rock-and-a-hard-place, that is.
GW
Yeah, borane and borax are not the same. There used to be a laundry detergent with borax in it, so that form of boron is fairly plentiful. I think it was "Twenty-Mule Team", and I think Ronald Reagan was its TV pitch man. This was 1955 stuff.
I think Quaoar likes my Mars landing boat designs. He's looking for better propellants for it. I did take a look over in the other thread at that organo-metallic stuff with the lithium in it. Isp-wise, it does look pretty good. Lithium-poisoning-wise, maybe not so good. Peroxide stability-wise, not good at all.
GW
The Rube Golberg rig on Curiosity was the bandaid required to get high enough end-of-hypersonics altitude to use a chute effectively, and still land a 1-ton rover with a fixed total spacecraft mass budget. There's a lot of things coming together constraining that problem.
The ballute being discussed in this thread is an alternate bandaid required to make chutes feasible with even bigger masses to be landed. The bigger the mass, the earlier in the entry hypersonics your ballute has to be deployed, and the more it looks like its own inflatable heat shield. Whether you tow it behind or ride behind it makes no real difference, except that riding behind eliminates your heat shield at the cost of inherently-poor vehicle attitude stability.
There is a completely different approach: supersonic/hypersonic retro propulsion. Great big items will come out of hypersonics at local Mach 3 (about 0.7 km/s) somewhere near 5 km altitude on Mars, instead of the 15-30 km needed to make a chute work. But around a gee or two of retro thrust can take you from that 0.7 km/s to zero in those 5-15 km of slant path length.
Folks are afraid of it because we haven't tried flying this way yet. It's just "new". But I believe you can land objects approaching 100 tons this way. I think there are fundamental solutions to all the design issues.
GW
On the assumption that the specific heat ratio of the exhaust gases really is 1.2, I get right at an exit Mach of 5.57 for the 250:1 exit bell area ratio quoted in the article. That corresponds to a chamber/exit pressure ratio of 4767; or for 500 psia chamber, an exit static pressure near 0.1 psia. The corresponding vacuum thrust coefficient calculates as 2.0048, unpenalized for any nozzle inefficiencies. For this, c* has to be 7527 ft/sec, which is right up there close to the values quoted for LOX-LH2. So, this is very good stuff.
On the other hand, I get a peroxide/fuel ratio by mass of 3.790 at stoichiometry. The optimium ought to be only somewhat smaller than that, not smaller by a huge amount. Yet the data table in the article says the optimum O/F ratio is 0.7 by mass, which is very wildly different. In comparison, stoichiometric for LOX-LH2 is by mass O/F 7.936, while opt r = 4 to 4.5.
That says at least some of the items in the table are not what everyone would normally interpret them to be. That casts quite a bit of doubt on just using stuff out of that article without independent verification, because somebody was playing sales games with this. A second indicator pointing the same direction is the claim of non-toxicity, which doesn't square well with the medical effects of small doses of lithium we were discussing earlier.
Now I tried estimating c* from chamber temperature out of their table, and my assumption of sp.ht. ratio=1.2, plus my stoichiometric estimate of exhaust gas molecular weight. I got about 5514 ft/sec. That was WAY off the CF-based value, which casts a lot of doubt on even my CF-based estimate. My assumption of sp.ht. ratio=1.2 is quite probably wrong.
Of the two, I think the CF-based estimate of c* is a lot more reliable than the first-principles estimate. It quite probably is in the 7000+ ft/sec ballpark. This stuff really is almost as good as hydrogen. I'd rather not use it in an atmosphere because of the lithium medical risks, but out in space, this is good stuff for very short term use.
The peroxide stability issue does come into play here. It's supposed to be 98+% to anhydrous, and that stuff is stable but for a handful of days at best.
GW
OK, that 469 sec Isp reported in the article for anyhydrous H2O2/Li3AlH6 is for theoretical perfect expansion all the way to vacuum at 100% nozzle efficiency, starting from a chamber pressure of 500 psia. Nothing you build will match that. It would be more helpful to report chamber c* velocity, which can be combined very quickly with a realistic thrust coefficient for the application, to produce a realistic Isp.
But, c* is a power function of selected chamber pressure: c* = K P^m. You also need an idea of specific heat ratio to get the thrust coefficient "right", but those are nearly always in the vicinity of 1.20. For LOX-LH2, at 100 psia and r = 4.5, c*= 7840 ft/sec; and at 1000 psia r = 4.0, c* = 7950 ft/sec. Those are for real combustion chamber efficiencies, not ideal thermochemical-code values.
My old Pratt & Whitney vest-pocket handbook reports a 100 psia vacuum isp for LOX-LH2 of 454. Yet at 3000 psia with finite expansion in the Shuttle engine, Isp was typically reported to be 467. So all these things do make a real difference.
I haven't got a good handle on the organometallic c* yet at 500 psia, much less its pressure dependence. But I'll see if I can get you a realistic estimate, and post it here.
The way you use it is estimate a thrust coefficient CF at application backpressure-limited expansion (and a realistic nozzle efficiency, usually near 0.983). Then CF c* / gc = Isp, where gc is the units-converting gravity constant in your F=ma/gc equation. That's the most realistic way I know to get a realistic Isp for your engine calculations.
GW
Whether or not 10 tons is appropriate depends on your mission design. If you want a reusable "Mars ferry" or "landing boat", that will be in the 30-60 ton range.
GW
Jet-A is the same product from the oil companies as JP-5, just bought to different specs. RP-1 and K-1 kerosene (for stoves and heaters) is almost the same product, differing mostly in cleanliness and additive packages.
Jet-B and JP-4 are also the same product from the oil companies, bought to different specs. The difference is that these are "wide-cut" products, meaning a wider mix of compounds with a wider range of molecular weights and boiling points than the other kerosenes. A half-and-half mix of gasoline with Jet-A is a pretty good approximation to Jet-B.
A diesel engine can be made to run on almost any fuel, as long as it has a low autoignition temperature with air, which means low octane and high cetane numbers. Diesels would run quite well on natural (or casing head, or "drip gas") gasoline, which is typically about 30 motor octane. Spark ignition requires high autoignition, which is high octane / low cetane.
Turbines don't give a tinker's damn about octane or cetane, they can be made to run on just about anything. That would include gasoline, jet fuel, diesel fuel, ethanol, methane, and hydrogen. The mods are to the fuel injection rig, not the engine. You have to be able to meter and finely-spray it, nothing more.
About the only difference between biodiesel and no. 2 petroleum diesel is heating value (just under 18,000 BTU/lb biodiesel vs almost 19,000 BTU/lb for petroleum diesel), and freezepoint (biodiesel without additives freezes about +29 F). Otherwise, biodiesel works as a drop-in fuel in diesel engines. No mods required at all. But, while biodiesel is made from vegetable oil in many cases, it is not vegetable oil. Vegetable oil is a diesel fuel, but not a drop-in fuel.
I used 20 to 30% biodiesel blends in jet fuel about 20 years before interest in this subject became widespread. It's a drop-in fuel in turbine engines. The differences are freezepoint nearer -20 F than -58 or -60 F, and a heating value loss so small you cannot find it in test data on the dynamometer or in flight. I did find a good trace anti-freeze agent that worked to -68 F or lower, in a B-30 blend with Jet-A, but that's another story.
Ethanol performance in turbines is lower and proportional to volumetric heating value. Ethanol doesn't work in diesels, because it has super-high octane and super-crappy low cetane. In spark ignition engines, you can blend it with gasoline up to about 35% ethanol, and not do any modifications at all. Beyond E-42, there are three required modifications, which if not done right, results in poor characteristics.
I have done all these things. For many years now.
The trouble on Mars isn't the fuel, it's the oxidizer. None is in the atmosphere. That's totally unlike here.
GW