New Mars Forums

Here on earth the aggregate in concrete is small stones.   Dirt as we know it here does not work.  Too fine-grained,  and the organics decompose,  leaving behind voids.  The key to concrete as a building material is the sand and stones:  two widely disparate particle size distributions as reinforcement. 

Here,  concrete is cement plus sand (around 0.5 mm diameter) plus gravel/small stones (around 50 mm diameter).  And,  both the sand and the stones are washed very clean,  and thoroughly dried.  Not only that,  they are carefully sieved for their size distributions.  The chemistry that sets the cement is a hydration reaction of the calcium oxide with water.  It is slightly exothermic,  and requires the water to be in the liquid phase. 

Any concrete variant on Mars will require a source of calcium oxide (unslaked lime) for the cement.  Sand and stones should be no problem,  but the lime is a problem.  That means we need something widely available that is peculiar to the physical chemistry of Mars,  to serve as the source material for cement. 

I suggested ice because it is widely available there and here,  and because the same sort of particle reinforcement would provide similar properties.  Plus,  on Mars,  temperatures are well below the freezepoint of water. 

There are fiber-reinforced versions of concrete in use here.  So a basalt fiber reinforced material would make a lot of sense there.  With ice there is no chemical reaction,  only a phase change.  But,  the result is a strong composite material anyway. 

Building materials (like concrete) are generally particle-reinforced composites,  and they have to be very inexpensive,  and widely available.  We may not be able to identify viable candidates until men have actually landed on Mars and played with the local materials.  It is difficult in the extreme to program a robot to look for stuff like this,  precisely because you won't know it until you see it. 

That's why I keep saying don't plan on successful ISRU on the first mission.  Do the detailed exploration and the ISRU experiments on the first mission,  and use that new knowledge to plan an ISRU scheme that actually will work,  on the second and subsequent missions. 

GW

To reinforce something with basalt fibers means you can't heat to basalt meltpoint,  or the fibers melt.  Use those fibers as the reinforcement in some other matrix instead. 

How about ice?  I once froze a wet dishrag,  and used it as a bat to hit a ball a very long way indeed.  Mars is a cold place.  Basalt fiber-reinforced ice might be a pretty good building material.  Bricks and panels. 

Or a blend of materials: 

Compression members could be ice with a regolith aggregate in it,  analogous to concrete.  For bending and tension situations,  imbed some ropes made of basalt fiber to act as the rebar. 

Ice and basalt seem plentiful on Mars,  although individual sites most likely vary a lot.  There just might be a real future for ice-regolith-basalt fiber as the analog to reinforced concrete. 

Toughest part is melting the basalt.  It would be nice if someone spotted an active volcano with basaltic lava. 

GW

RobertDyck:

Is Nextel 440 the replacement for the older Nextel 312?  Silicate fire curtain cloth.  Your parasol idea is intriguing. 

Given a low enough ballistic coefficient,  I think it might work for aerocapture,  or possibly even re-entry to a landing. 

GW

I have my doubts that concrete as we know it would ever cure in intense cold,  unless the water in it remained liquid.  The hydration reaction requires liquid water. 

I could cast slabs down to about 29 F (-2 C),  but only because the cure is very slightly exothermic,  and the slab never cooled below 32 F (0 C) until the "cure" was finished. 

And,  that only gets it to "walk-on-it" strength.  Typically,  concrete slowly cures further over about 30 days to its full strength.  It's not exothermic enough to get a difference from ambient during this period. 

I rather suspect we'd need a new kind of concrete on Mars.  Something will work,  just not what we do here. 

GW

Long-duration LH2 storage can be problematical because of leakage,  and because of boil-off.  I know gaseous hydrogen will leak right through the steel welding gas bottle over time.  Not so bad in liquid phase,  but still,  we're looking at about 2 years for a round trip,  unless one bites the bullet and finds some way to fly much faster. 

Here's a really nutty idea:  for hydrogen that won't be used until many months later,  why not ship it as ice?  On-orbit solar thermal to melt,  on-orbit solar PV to electrolyze.  So it takes a while with a low-powered system.  It'll be a while before you need it.  It's a very damage-tolerant and minimal-maintenance way to ship your NERVA propellant and your oxygen supply. 

I do like the idea of reusable tankage on the nuke ship.  Throw nothing away.  It can fly many missions,  and not just to Mars.  Pretty much any design capable of ferrying men from LEO to Mars orbit can be used to visit Venus,  Mercury,  and some of the NEO's.  Refitted with a more powerful engine,  it might take men to the main asteroid belt.  You need a habitat module big enough for comfy living on 3-year time horizons.  I'd suggest a solar flare radiation shelter that is also the flight control deck. 

Here's another nutty idea:  build the vehicle as a long train of modules.  Put the habitat on the end away from the NERVA,  and spin it slowly end over end for artificial gravity.  That solves a whole host of life support issues associated with zero-gee.  56 m radius at 4 rpm is pretty close to the 1 gee we evolved with. 

GW

Louis:

I went and looked at the link.  Very intriguing.  Cold structural blocks and tiles I can see us mining from basalt provinces.  But for fiber,  why go to the trouble of melting it?  Why not catch it liquid coming out at the mid-ocean ridges,  and use the sea to chill the filaments solid as you extrude them?  Sounds like some sort of undersea robot materials-harvesting plant to me.  Probably self-transportable as a submarine vehicle. 

The active lava flows hitting the sea from Hawaii I think are basaltic,  too.  There ought to be a lot of volcanic basaltic lava harvestable all over,  already liquified by mother nature here on Earth.  Not all volcanoes are basaltic,  but many are.  On Mars,  that's another problem.  I don't recall any reports of active volcanism reported there. 

Intriguing idea,  though.  If it can be used as a glass fiber replacement,  then cloth can be woven from it. 

GW

UV-protected polymer-insulated electric cables:  Sorry,  I was thinking more of modified supplies brought from Earth to make the initial installation. 

Basalt fiber?  Can you weave this into a cloth?  If so,  it might be possible to sleeve metal wire with it as an insulation,  resembling the really-old fabric insulated wires no one sees anymore,  except in older-model electric stovetop cookers.  Those are asbestos-fiber fabrics.  Would basalt be an insulator,  or is it too conductive to serve?  I honestly don't know.

Also,  would basalt fiber pose the same inhalation hazard that asbestos fiber and fiberglass pose?  Does anyone know? 

GW

If you do TSTO with a rocket+ramjet 1st stage,  you have to decide fundamentally between vertical-takeoff/ballistic/wingless and horizontal-takeoff/lifting-flight/winged.  The decisions on which stage does what with wings are secondary to that first level choice.  At least,  that's what my attempts at sizing systems and components says.

If you do HTO,  I recommend a winged booster airplane carrying a second stage that could be either a small rocket plane or a ballistic rocket.  My best approach so far uses a high-speed ramjet design as the fuselage (for the biggest cross-sectional area),  with rockets and propellant tankage in the wing fillets and wings.  The ramjet has a centerbody external compression spike in the nose,  and the booster pilot rides there. 

It takes off on rocket only,  climbing and accelerating to M1.6-ish,  then shifts to ramjet propulsion only.  It climbs at M1.6-ish in ramjet,  gradually pulling over level and accelerating,  at about 60,000 feet.  There it accelerates in ramjet to about M5.5 to 6,  depending more on the drag of the configuration than ramjet thrust.  Light the rockets too,  pull up sharply,  and stage at a path angle for ballistic flight of the second stage.  The booster then cuts rocket and throttles back ramjet,  does a decelerating pull-down 180-turn,  and cruises back to base at M1.6-ish in ramjet.  At base,  it goes subsonic powerless for a glide landing,  retaining a small kitty of rocket propellant for go-around or divert. 

The 3 most important HTO staging variables,  in order of importance,  are velocity,  path angle,  and altitude.  This kind of approach could get you close to M6 (about 1.8-1.9 km/sec) at staging,  path angles 30-60 degrees,  and probably no more than 60,000 feet (18 km) altitude. 

The ramjet has an external- or mixed-compression inlet with a min takeover speed around M1.5 to 1.6,  a slightly convergent-divergent nozzle with a throat area pretty near 65% of the combustor duct area,  a sudden dump flameholder approach,  probable perforated air-cooling sleeve,  and Inconel-X skins like the X-15.  (Shock-impingement heating damage becomes intolerable above M6 with all known materials,  so I see no point to trying a scramjet for this.) 

Rocket T/W at TO need not exceed the [EDIT:  0.x instead of the original 1.x,  sorry,  my mistake ] 0.3-0.5 range of any other airplane.  The inert fraction of the booster will look more like an airplane because that is what it really is:  30-40%.  If you try to push the staging altitude too high,  the ramjet is unable to accelerate fast enough,  and cannot fly the return range.  Staging altitude is a tradeoff driven by that range.  I don't have an exact figure for that yet. 

The velocity requirement for the second stage is around 20,000 feet/sec (6.1 km/sec),  which is high,  but do-able with the one-shot technologies that we already have.  Its nozzle pretty much always operates in near-vacuum conditions. 

I know a lot less about the VTO case,  except that it is more of a ramjet-assisted rocket vehicle.  The idea is to shoulder some of the dead-weight loads onto as much higher-Isp / lower frontal-thrust ramjet as you can afford.  I don't think the ramjet will exceed about 20% of the first stage thrust,  maybe not even 10%. 

Vehicles like this typically leave the sensible air (60,000 feet, 18 km) fairly slow:  M 2-ish.  Thus,  a low-speed ramjet design is better:  normal-shock / pitot inlet,  convergent-only nozzle capable of unchoked operation,  and usable thrust at high subsonic flight speeds (say,  M 0.7+).  This thing most likely would be a set of strap-on pods staged off around 60,000 feet,  and perhaps flown back to launch site for recovery like a big model airplane,  but a glider,  of course.  Nose wheel and tail skids,  like X-15.  Some sort of extendable wing?  Scissor swing wing? 

To solve the dead-weight problem between launch and ramjet ignition,  I'd package a missile-type integral solid booster inside the ramjet case,  which is lined with simple ablatives.  Plain steel,  no external heat protection needed.  I'd make the ramjet pod modular,  for easy re-insulation and reload by very few folks with very little effort.  An inlet module,  a concentric fuel tank and air duct module,  a combustor/booster module,  and a nozzle module. 

Mix and match and stack,  then hang these on the core vehicle as strap-ons.  It'll shed a port cover and a booster nozzle at transition to ramjet.  The core rocket is a typical TSTO with staging around 10,000 feet/sec (3 km/sec) well outside the sensible atmosphere.  You just get to add a few % more payload than it would normally carry. 

Those are my best guesses right now. 

GW

For Bob Clark:

It makes a lot of sense for the ramjet and the rocket to burn a common fuel.  My old article that Rune referenced just above does exactly that. 

There are several kerosenes: JP-5/Jet-A/Jet-A-1,  RP-1,  and commercial K-1 kerosene for kerosene heaters and lanterns.  These all have almost identical distillation curves and properties.  K-1 is the dirtiest. 

Clean is very important:  fuel metering orifices clog easily.  Liquid methane is the very cleanest,  though.  You have to rework your fuel system to handle a cryogenic fuel, but fundamentally,  both ramjet and rocket can burn liquid methane. 

Rocket and ramjet might be made to burn diesel,  but vaporization is difficult.  Both can be made to burn gasoline quite easily.  The downside is the flammability and explosion hazards associated with gasoline.  If you use gasoline (and both rocket and ramjet will burn it),  there is no point to high octane rating in either engine:  30 MON drip gas is fine. 

Anything a ramjet can burn,  a gas turbine can be made to burn as a drop-in fuel,  FAR's and type certificates notwithstanding. 

There is an old ramjet fuel known variously as RJ-5 / Shelldyne-H.  It is an artificial synthetic,  a single substance,  not a mixture of compounds like all distillates.  It does behave exactly like kerosene in all the subsystems of a kerosene ramjet,  it's just denser:  sp.gr > 1.  (It sinks in water,  unlike all the petroleum liquid hydrocarbons.)  But,  it's expensive.  Yet,  propellant costs are but a tiny fraction of launch costs. 

For first stage use,  density x Isp is more important than Isp.  That's why kerosene beats hydrogen for first stage fuel use in practical designs.  RJ-5 is really good for that.

GW

Rune:

I know nothing about any "super Draco" thrusters for Dragon.  The ones they have are supposed to be the launch escape system for any manned Dragon,  so capsule T/W is substantially larger than 1.  For the Mars mission paper I gave at last August's convention in Dallas,  I re-engineered some data from the Spacex website for Dragon as they had it posted last summer.  I was showing 0.9 km/sec delta-vee capability with 6 suited astronauts on board.  If you put more propellants in the unpressurized module,  and connected that to the Draco system,  it adds around 1.4 km/sec more to the capsule's delta-vee capability.  I looked at that to get 2+ km/sec total delta-vee for Dragon as an emergency escape crew return vehicle from a very high-speed Mars return in an emergency,  somewhere well above 15 km/sec.  Maybe as much as 25 km/sec. 

Stock Dragon post reentry has no unpressurized module.  So the capsule's delta-vee capability is around 0.9 km/sec utter max,  post re-entry.  That's enough to land from considerable height without a parachute,  but it's nowhere near enough to make the entire landing without a chute.  Not on Earth.  Moon?  Not even there.  Any "pinpoint" to the landing is during that final burn,  but it's not all that much maneuvering.  You know advertising flacks,  they do tend to oversell stuff. 

On the other hand,  Dragon's true significance is that it is the very first capsule in history intended to be re-used,  heat shield and all.  It does need a new unpressurized module and a new nose cap for each flight.  But,  that heat shield was designed for a free return from Mars at about 50,000 feet/sec [15 km/sec] entry velocity.  You could re-use it even after a return from the moon (36,000 feet/sec [same as 11 km/sec] for comparison).  Coming back from LEO at around 26,000 feet/sec [8 km/sec] is by comparison "easy",  since the heating is proportional to velocity squared.  I'd guess they could fly the same heat shield to LEO about 4 times before replacing the pieces that make it up.  And it is segmented. 

None of Dragon's capsule-type competitors are intended to be reusable.  Not Orbital's Cygnus,  not Boeing's scaled-back Orion.  While my contention is that Spacex achieved its lower costs per unit of payload to LEO by smaller logistical support tail,  not actual reusability of anything,  having a capsule you can fly several times is still going to help significantly.  Plus,  it's simply a technical capability that we need.  And we have needed it for a long time.  It is a major accomplishment. 

I'd have to go find my original notes for that re-engineering I did on Dragon to find the Isp I used for the Dracos.  But I think it was likely closer to 270 sec than 300 sec.  There's quite a gap between what is theoretically-possible under perfect expansion,  and what can actually be had with a real engine bell,  even in vacuum. 

GW

Landings:

I quite agree with RobertDyck.  I'd add these things to the discussion:

The main problem with lifting bodies,  aside from that speed where control is "iffy",  is the high landing speed due to very high wing loading or broadside ballistic coefficient.  A vehicle designed to fly into space and back is inherently still a bit heavy on landing because of the heat shield and the control propellants.  Most of the those lifting body designs had touchdown speeds at Edwards in the 200-300 mph range.  Stability on the ground during roll-out (skids or wheels) is a serious issue,  unless the vehicle is rather long,  like X-15,  which landed fairly well at 200 mph. 

Gemini was originally a Rogallo wing-type parasail design,  with skids for dry lake bed landing.  My memory may be off,  but I think the landing speed was near 200 mph like X-15.  The Rogallo was steerable,  but the short "wheelbase" of the skid system was unconditionally unstable on the ground at any speed.  Test articles always flipped over and went bouncing down the dry lake.  Good thing they weren't manned.  It just never worked,  so they went with a parachute and water landing.  The capsule was already in production,  so that's why the chute came out the side instead of the nose on Gemini. 

I've got no real problem with MMH N2O4 residuals after landing.  Just don't go sniffing the rocket nozzles for a while.  I remember NASA did have concerns,  when the shuttle first flew,  but they got over being nervous about it after a while. 

The Draco thrusters on Dragon are rather powerful,  and it has a lot of delta-vee built into it.  If one reserves a kitty of propellants,  one could fire them last second to slow a chute landing on land to survivable levels for the crew.  It would feel like a car crash,  just like Soyuz.  If it had shock-absorbing legs,  the heat shield would survive for re-use,  too.  Of course,  they're extra weight.  Reduces payload or crew size a tad.  But the water landing version is said to be able to carry a crew of 7. 

GW

I may be the victim of past thinking,  but I think the main decision to make is whether you want lifting flight or ballistic flight post re-entry.  Putting the weight off-center a bit in a traditional capsule gets you side forces of some significant fraction of the drag force during re-entry hypersonics.  This allows you to shift the location of the landing ellipse by a couple of ellipse dimensions L-R,  or downrange-uprange.  But it's worthless as a directional control once the hypersonics are over.  From that point,  you're ballistic.  Control there requires real wings. 

Wings are pretty useless during re-entry because they're fragile and vulnerable to overheat,  as in the shuttle design,  the X-20 we abandoned,  and the X-15 which was considered for orbital flight using a Titan-3 booster,  but which plan was abandoned because it could not survive re-entry.  In the X-15's case,  this was because of wings burning away really fast at nose-first attitudes,  but inadequate structural strength to re-enter more nose-high ("semi-broadside") like the shuttle.  The X-20 was to have addressed that weakness with experimental ablatives,  and/or transpiration cooling,  and/or heat-sinking.  Shuttle did it with slow ablatives (carbon-carbon LE's)and refractories (low density ceramics). 

There are a couple of things we haven't tried yet:  some sort of ablative-protected inflatable structures,  and protection by an engine plume between vehicle and hypersonic shock layer.  There is great potential in both ideas,  but I see nothing at all going on around me.  Not a space agency anywhere is exploring this,  and private concerns “typically” do not invest their own money in exploring radically new technologies (that’s why there have to be government agencies,  no one else will do that job).

The aerospike nozzle is indeed what you might want for engines of any type that must function across varying backpressure going up.  Coming down,  I fear the sharp aerospike might be rather vulnerable,  if you used the engine plume as your heat shield.  But I don't know.  No one does,  not yet.  Aerospike would have a lot of application to first stages going up,  except our current methods for building them end up being heavier and inconveniently-packaged for a cylindrical shape,  and the nozzle efficiencies typically run 2 or more % lower than a traditional bell.  It only looks better at off-design backpressure.  It's a non-trivial trade-off,  and unfamiliar territory,  so no one has really done it.  X-33 was supposed to do it,  but never flew for a whole plethora of reasons,  some technical,  some not. 

Recovering a upper rocket stage in its entirety would probably be best done with a blunt ablative on the front end,  and some ablative-protected inflatables around the engine(s) to double as floats,  and as water-impact shields for the engine bells.  Take the main sea impact on the heat shield,  we already know that works fine.  You just need to add a series of drogues and chutes that deploy out the rear,  around the engine(s).  Biggest risk I see is shroud line fouling on the engine(s).  But it's doable. 

That's a lot of extra gear and equipment.  Plus,  the tankage has to take the "whack" when it hits the sea at dozens to a hundred gees.  You're just not going to build a thing like that at 5-8% inert fractions.  Steel,  aluminum,  and titanium are "only so strong".  I really don't see organic composites as a tankage option for something intended to survive Mach 25-class re-entry,  either.  It's just too thermally harsh.  Something that could be re-flown a lot of times is just going to be heavier.  10+% inert,  probably 15+%.  Just guessing. 

Lower stages are far easier.  Hypersonic heating under Mach 10 or 12 is a whole lot less intense.  Organic composites in some areas become feasible.  You could even put flyback wings on it,  perhaps.  But whatever you do,  it won't be 5-8% inert.  The way around higher inerts at fixed rocket performance is more stages.  That,  too,  is well-known and well-proven.  More stages does not necessarily mean more cost.  But you do have to keep it simple so the logistical support tail can stay small.  The rocket / system designs need to look more like battlefield weapons than our traditional launch rockets.  Trying the first steps into that model is the real secret of Spacex's lower costs.  They've re-used nothing yet.  And at 5% inerts,  they won't.  You can bet any flyback first stage with landing legs or whatever,  will be a lot heavier than the stage they're flying now. 

As for the suborbital rocket planes,  those typically leave the air and return at Mach 3 or so.  That’s actually very little transient heating,  and it is a short transient.  Not at all the same problem as sustained flight at Mach 3.  Even with organic composites as exposed structure,  it is possible to heat-sink your way through that short re-entry.  There is no radical new heat protection or structural technology there with Spaceship Two or Lynx.    The “radical” item in Spaceship Two is the folding tail.  Lynx’s “radical” technology  is the long life / low maintenance engine. 

GW