Best propellant & stuff for a Mars spaceship and lander

Quaoar · 2018-07-01 04:30:41

I would like to imagine a modular, completely reusable orbit-to-orbit spaceship, also able to spin for artificial gravity.
She does all the maneuvers propulsively, so the delta-V budget is almost 5.7 km/s for the outward trip and 5.7 km/s for the inward one. Let's take 12 km/s, considering plane changes and course corrections.
The lander must be reusable, must be also used as an habitat, and can be send one-way with propellant tanks: an orbit-surface-orbit two-way trip costs almost 5 km/s of delta-V (please correct me if I'm wrong)

1) NTO-MMH
pros: self-igniting and storable
cons: expensive, highly toxic, the exhaust velocity is only about 3.3 km/s.
Mass ratio for 12 km/s of delta-V: 38
Propellant needed for a 100 ton spaceship: 3700 tons
Mass ratio for a two-way Mars lander: 4.6
Propellant needed for a 100 ton lander: 360 tons

2) LOX-RP1
pros: better exhaust velocity than NTO-MMH, cheap, non-toxic, RP1 is storable.
cons: not self-igniting, cryocooler needed for oxidizer, limited exhaust velocity of about 3.4 km/s.
Mass ratio for 12 km/s of delta-V: 34.1
Propellant needed for a 100 ton spaceship: 3310 tons
Mass ratio for a two-way Mars lander: 4.4
Propellant needed for a 100 ton lander: 340 tons

3) LOX-CH4
pros: better exhaust velocity than NTO-MMH, cheep, non toxic, possibility of production in situ.
cons: not self-igniting, cryocooler needed for both oxidizer and propellant, limited exhaust velocity of about 3.65 km/s.
Mass ratio for 12 km/s of delta-V: 26.8
Propellant needed for a 100 ton spaceship: 2580 tons
Mass ratio for a two-way Mars lander: 3.94
Propellant needed for a 100 ton lander: 294 tons

3) LOX-LH2
pros: it has the best exhaust velocity obtainable from a realistic chemical propellant (4.6 km/s), possibility of production in situ from a buried glacier or from the ice of Phobos (if there is any)
cons: not self-igniting, cryocooler needed for both oxidizer and propellant. LH2 has a lower temperature than other cyogenic propellants and is more difficult to maintain. LH2 has a very low density and needs a bigger tank that weights almost 10% of the propellant.
Mass ratio for 12 km/s of delta-V: 13.6
Propellant needed for a 100 ton spaceship: 1260 tons
Mass ratio for a two-way Mars lander: 3
Propellant needed for a 100 ton lander: 200 tons

I don't want to talk about nukes in this post, but just out of curiosity, I'd like to report the data for a solid core NTR
4) LH2 NTR
pros: exhaust velocity of about 9.2 km/s, it may be also used as a power generator in a bimodal rocket
cons: it would be surely boycotted by public opinion, cyocooler and huge tanks needed for LH2, further R&D needed to resume the studies of 60 years ago.
A nuclear lander is unlikely the best option: a lander must be compact so it might be difficult to shield the astronauts. But just out of curiosity:
Mass ratio for a two-way Mars lander: 1.73
Propellant needed for a 100 ton lander: 73 tons

Then I would like to talk about the rocket type: which kind of cycle? Gas-generator, expander-cycle, staged-combustion, or pressure-feed?
Which one may be the best suited for this kind of spaceship?

Last edited by Quaoar (2018-07-01 06:14:13)

kbd512 · 2018-07-01 07:52:37

Quaoar,

The most practical orbit-to-orbit interplanetary transportation system is the SEP-enabled ITV that I've described elsewhere. This ITV concept is modular. The engineering and propulsion module (ITV-E) would have BFS (ITV-C configuration) or a rotating wheel artificial gravity habitat module (ITV-P configuration) strapped to the front of the ITV-E. By virtue of the ITV-E's megawatt class solar power array and efficient ion propulsion, transit duration to Mars from LLO is identical to any practical form of chemical propulsion.

The best upper stage engine for the BFS is the RS-25. Apart from the J-2, no other high performance rocket engine has been as thoroughly tested as the RS-25. Aerojet-Rocketdyne is restarting the RS-25 production line and that's basically a done deal.

All of our advanced chemical fuels R&D revolves around H2. There's not nearly as much work done on CH4 because we simply pump it out of the ground and burn it. It's a legacy liquid hydrocarbon. CH4 is readily available on Earth, but other potential sources are much more complicated to use. H2O is available everywhere that humans could potentially live. If there's no water sources wherever we're going, then there's little possibility of colonization in the near term.

The most useful area of research for NASA and Aerojet-Rocketdyne to expend funds on would be aerospike nozzles. If 3D printed aerospike nozzles were available, then precise engine throttling becomes practical through mixture flow control to the injector plates. At that point, the RS-25 won't resemble the RS-25 of STS fame, but it becomes incredibly useful for landing in various atmospheric environments. AJ went so far as to begin design work on a RP-1 fueled aerospike center section with a LH2 fueled outer section to combine boost and sustainment phases of launch in a single engine to enable the use of drop tanks. Apart from highly experimental combined cycle engines like SABRE, after we have working aerospike engines, we've more or less reached the zenith of pure chemical rocket engine propulsion technology.

BFS should be the last component that SpaceX develops, not the first. Develop, test, and fly the booster first. The entire SLS core stage could sit atop the BFR booster and deliver a 250t payload to orbit. If everyone standardizes 10m LCH4 fueled booster cores and LH2 fueled upper stages, then it becomes possible to substitute technology sets. The boosters and upper stages get shorter and wider and this whole business of vertical landing on other planets becomes a lot more feasible.

SpaceNut · 2018-07-01 08:19:32

Thanks for the run down of each fuel type for what it is capable of. The list of #1-3 have been proven for moon or mars journeys but so far that was one way for mars and the lunar mission durations were not long enough to prove viability after a mars cycle less insitu manufacturing.

We have talked about the need for a safe guard in either preload of landing site with all of what we might need just incase on that first mission we can not get any of the insitu manufacturing or food production going.

To which only #1 is possible for this caching approach for a mars cycle and if insitu Lox is possible then (#2) RP1 is another plausible after a first mission as a preload fuel for mars.

On mission #1 or 2 we prove making insitu CH4 (methane) is possible and we do liquify it with no problem then we can stop sending fuels for a safe guard and use them as backups for power generation or other applications for mars.

As you closed in your post something that GW is more capable of, it is those different rocket engine enhancements change only the performance a bit but has a mass trade off that is not always equal.

As for the rockets either we are using what we have with slight modification or we are using paper rockets still in planning.

kbd512 · 2018-07-01 08:38:16

SpaceNut,

On that note, RL-10, J-2, and RS-25 are all real man-rated rocket engines. If I had to design a man-rated upper stage for a vehicle, then I'd pick an engine technology that's already exceptionally well proven. That means using one of those engine types or some combination of them.

SpaceNut · 2018-07-01 09:10:08

Yes and all have been recently reworked for performance to which all you need is the remaining part of the rocket to be designed to go with them plus all other pieces to make it a complete mission to mars.

Quaoar · 2018-07-01 12:35:31

kbd512 wrote:

SpaceNut,
On that note, RL-10, J-2, and RS-25 are all real man-rated rocket engines. If I had to design a man-rated upper stage for a vehicle, then I'd pick an engine technology that's already exceptionally well proven. That means using one of those engine types or some combination of them.

RL-10 uses an expander cycle, while RS-25 uses a stage-combustion cycle. RS-25 needed a lot of maintenance and were disassembled after each Space Shuttle fly, so it might be not the best choice for an orbit-to-orbit spaceship that uses it many times during her trip. I don't know if pressure-feed rockets or expander-cycle rockets are more reliable or not.

Last edited by Quaoar (2018-07-01 12:36:37)

Quaoar · 2018-07-01 12:41:00

kbd512 wrote:

Quaoar,
The most practical orbit-to-orbit interplanetary transportation system is the SEP-enabled ITV that I've described elsewhere. This ITV concept is modular. The engineering and propulsion module (ITV-E) would have BFS (ITV-C configuration) or a rotating wheel artificial gravity habitat module (ITV-P configuration) strapped to the front of the ITV-E. By virtue of the ITV-E's megawatt class solar power array and efficient ion propulsion, transit duration to Mars from LLO is identical to any practical form of chemical propulsion.

It may even be more cheap using iodine instead of xenon. I've only some concern about spiraling inside the Van Allen Belt, but the astronauts can reach the ship with a capsule when she is out, and however she may also be very useful to send landers and other payload one-way to low Mars orbit. Please can you post some link to your spaceship, so I can make some comparison?

SpaceNut · 2018-07-01 14:14:29

Here is a topic search for BFR

will search for ITV topics next and will add them in.

All of which are Space X future rockets on the design plate. Some prototype hardware is in the process chain as well as implementation concepts.

Update on ITV and MCT
The ITV is an interplanetary Transport Vehicle
The MCT is the Mars colony transport

These are both intertweened in many discusions for space x moving targets on their plans.

These are external links for each of the Space x ships.
https://en.wikipedia.org/wiki/Interplan … ort_System
http://spaceflight101.com/spx/its-booster/
http://www.spaceflightinsider.com/organ … onization/

https://en.wikipedia.org/wiki/Mars_Colonial_Transporter
http://www.spaceflightinsider.com/organ … realities/
https://infogalactic.com/info/Mars_Colonial_Transporter

kbd512 · 2018-07-02 13:15:13

Quaoar,

Any high operating pressure pump-fed engine will require a lot of maintenance. SpaceX will discover that in due time, if they haven't already. The expanders are simpler, but high pressure turbopumps are maintenance intensive. A stock of replacement parts or replacement engines is mandatory.

After a bit more research into the topic, I agree about the use of Iodine to reduce propellant costs and simplify storage requirements. At 4,600kg/m^3, it's plenty dense for our uses.

Here's the Iodine Hall Thruster presentation from Busek Co. Inc. being developed on NASA's behalf (all ion thruster technology actually used by the US, along with the "green" AF-M315E monopropellant, was developed based upon work done by this company; a handful of advanced technologies such as MPD thrusters, like ELF, have been developed independently by universities and the like for research purposes):

Iodine RF Ion Thruster Development

GW Johnson · 2018-07-02 16:24:55

I beg to differ with Kbd512, but only a little bit.

Production pressure-fed engines have not existed in a long time, other than small thrusters. But with no moving parts at all, they can be very reliable, at the cost of extremely heavy propellant tankage.

Now-bankrupt XCOR Aerospace developed and flew in manned airplanes a unique piston-pumped rocket engine technology, including rating it for harsh cryogens to include liquid hydrogen. What they found was that this technology worked in smaller engines only, not seemingly being scalable to very large thrusts per engine. The biggest they looked at was the RL-10 of the Centaur upper stage.

They never did demonstrate anything but 10 times smaller thrust than an RL-10, though. Uniquely, in that smaller scale, they did finally demonstrate driving that piston pump assembly off the waste heat from the engine, via a third cooling fluid through some sort of heat engine. This is a technology I'd really like to see someone take on, and take to the next level. It has great promise, but no backing that I am aware of.

The promise is many orders of magnitude better reliability life than turbopumps of any kind whatsoever! Would you believe 2000-4000 hours of operation before overhaul? XCOR made claims like that, supported by at least a little data, before they finally went belly-up in a top management dispute of some kind. They went belly-up after expelling Jeff Greason, and abandoning the Lynx suborbital spaceplane. Maybe those decisions were not so good.

I really like the iodine RF thruster. It has very great advantages over any imaginable xenon system. However, this is potential currently restricted to very small scale. This will take a major scale-up and demonstration effort before it could ever be considered for moving large objects!

Not that that can't be done, but small businesses usually do not get funded to do that, not by NASA or any other US government agency. The good-old-boy/favored-contractor network and politics are just too strong, as is the not-invented-here attitude so prevalent at NASA and all of the old space contractors.

I could not read the charts in the briefing for thrusts at the 3 cm size, but I did notice the scales were in milli-Newtons for something on the order of a kilogram of thruster hardware, at dozens to a few hundred watts of supplied electricity. You're looking at something like 10^-3 Newtons/kg, and something like 10^-5 Newtons per watt (10^-2 Newtons/KW).

For a 10-kg class cubesat spacecraft, the briefing talked about 258 days to spiral out from geosynchronous orbit to near-enough escape to reach the moon. That would be at vehicle thrust/weight ratios on the order of 10^-4. A thrust/weight nearer 10^-3 would shorten that, somewhat. This is not something that would be suitable for pushing manned craft, given the growth of life support mass with mission time, unless you also buy a high-thrust chemical rocket to carry the crew to the spacecraft out near escape. Kinda negates the savings.

If you could get the vehicle design thrust/weight nearer 10^-2, the spiral-out times would be significantly shorter. For a 100-ton spacecraft, we are talking about required thrusts on the order of 10 Newtons (some TEN THOUSAND milliNewtons!!!). Such an array of thrusters might mass around a ton (like 1 percent of the total spacecraft).

The power required would be close to 1 MW. That is one big solar array, or one big nuclear reactor power system. I seriously doubt there would be much spacecraft weight left over for the crew, their habitat, and their supplies.

I might be wrong, and I certainly hope so, but this looks to me like a very immature technology, but with great future promise.

GW

Quaoar · 2018-07-03 08:15:39

GW Johnson wrote:

I beg to differ with Kbd512, but only a little bit.
Production pressure-fed engines have not existed in a long time, other than small thrusters. But with no moving parts at all, they can be very reliable, at the cost of extremely heavy propellant tankage.
Now-bankrupt XCOR Aerospace developed and flew in manned airplanes a unique piston-pumped rocket engine technology, including rating it for harsh cryogens to include liquid hydrogen. What they found was that this technology worked in smaller engines only, not seemingly being scalable to very large thrusts per engine. The biggest they looked at was the RL-10 of the Centaur upper stage.
They never did demonstrate anything but 10 times smaller thrust than an RL-10, though. Uniquely, in that smaller scale, they did finally demonstrate driving that piston pump assembly off the waste heat from the engine, via a third cooling fluid through some sort of heat engine. This is a technology I'd really like to see someone take on, and take to the next level. It has great promise, but no backing that I am aware of.

Thank, GW
So, if we want to build our orbit-to-orbit spaceship (something like your modular Johnson-express) in the next 10-15 years, which kind of feed-system would you use for the rocket?

P.S.
What about pistonless pump rocket?

http://www.flometrics.com/wp-content/up … PC2003.pdf

http://www.flometrics.com/wp-content/up … 31-314.pdf

https://tfaws.nasa.gov/TFAWS04/Website/ … ulsion.ppt

Last edited by Quaoar (2018-07-03 09:33:35)

GW Johnson · 2018-07-03 11:27:28

Hi Quaoar:

This pistonless pump idea is intriguing. If you ever get a leaky check valve, it will most likely lead to a thrust-oscillation problem. I noticed two things. First: like all propellant feed systems, there is an ullage problem in zero-gee flight. You have to establish a little acceleration such that the liquids settle to the bottoms of the tanks, or else you won’t pull any liquids through the tank drains. It has to be enough to overcome surface tension effects for the globules to become one big liquid pool.

Second, the longer you intend to burn, the bigger your supply of high pressure gas must be. The XCOR piston-pump design I mentioned avoids this issue by recirculating its heat engine drive fluid (which is the rocket engine cooling fluid), taking advantage of phase changes liquid vs gas. I don’t know the details, but I presume there is an electrically-driven pump that starts the recirculating process, until the heat engine can take over that duty.

Third, and common to all designs, if the gravity acceleration vector is pointed the wrong way, your ullage problem gets quickly insoluble. If the tanks were upside down here on Earth, likely none of these schemes would work. Could be a problem for a tank just laying on its side.

It might be possible to replace the high-pressure gas source with a much smaller high-pressure gas tank fed by a solid gas generator, or a whole series of them. These would be propellants designed to burn cleanly, and at modest chamber temperatures. Such things already exist, such as the solid starter cartridges for starting aircraft engines with no battery, since WW2. Automotive air bag cartridges are another example.

The tank ullage problem has been solved since at least the Saturns of the 1960’s by solid propellant ullage motors. These are small solid motors that provide just enough thrust to settle the propellant liquids into a pool, for just long enough to have that settling occur, plus get a liquid engine start. On the Saturn 5 second stage, there were 3 of these, each a small pancake motor about 6-7 inches diameter and 2-2.5 inches thick. These were made at the plant where I once worked, where Spacex tests its rockets now.

GW

Quaoar · 2018-07-03 12:29:53

GW Johnson wrote:

Second, the longer you intend to burn, the bigger your supply of high pressure gas must be. The XCOR piston-pump design I mentioned avoids this issue by recirculating its heat engine drive fluid (which is the rocket engine cooling fluid), taking advantage of phase changes liquid vs gas. I don’t know the details, but I presume there is an electrically-driven pump that starts the recirculating process, until the heat engine can take over that duty.

Do you think it's possible to build a bigger version of the piston rocket, something like 6-8 MN of thrust and 70-100 bar of chamber pressure?

GW Johnson wrote:

It might be possible to replace the high-pressure gas source with a much smaller high-pressure gas tank fed by a solid gas generator, or a whole series of them. These would be propellants designed to burn cleanly, and at modest chamber temperatures. Such things already exist, such as the solid starter cartridges for starting aircraft engines with no battery, since WW2. Automotive air bag cartridges are another example.

What kind compound would you use for this solid gas generator?

GW Johnson wrote:

The tank ullage problem has been solved since at least the Saturns of the 1960’s by solid propellant ullage motors. These are small solid motors that provide just enough thrust to settle the propellant liquids into a pool, for just long enough to have that settling occur, plus get a liquid engine start. On the Saturn 5 second stage, there were 3 of these, each a small pancake motor about 6-7 inches diameter and 2-2.5 inches thick. These were made at the plant where I once worked, where Spacex tests its rockets now.

It is possible to use a miniaturized version of the Space Ship One hybrid rocket as a multiple-restart ullage rocket?

GW Johnson · 2018-07-03 13:16:28

Well, the design that XCOR was pursuing was the RL-10 as it exists on the Centaur upper stage. That was something like 25,000 lb thrust per engine, which is crudely 0.1 MN. They thought they might double it. What they demonstrated in test was 2500 lb thrust, or about 0.01 MN.

I don't know what the XCOR piston-pumped engine chamber pressure really was, but a value near 1000-1500 psia (67 to 100 atm or 67 to 100 bar) seems likely.

As for gas generator propellant formulations, I never worked with things suitable for this application. I worked with real thruster-rockets, and with fuel-rich formulations to be used as ramjet fuel supplies in lieu of liquid fuels. Can't advise you there, but I know such formulations exist, including solid propellant oxygen generators used on MIR.

I am sure that a hybrid could be used as an ullage motor. Be aware that the state of hybrid technology is not yet as advanced as solid or liquid rocket technology. But it could be, given the effort to do so. I'd recommend a storable oxidizer like a strong acid, to get hypergolic ignition of a fuel grain that is a simple particulate-loaded rubber. Such storables can be pressure-expelled from bladders in tanks, eliminating their ullage problem, something we still cannot yet do with cryogens, even "soft" cryogens like liquid oxygen.

GW

Quaoar · 2018-07-03 16:33:29

GW Johnson wrote:

I am sure that a hybrid could be used as an ullage motor. Be aware that the state of hybrid technology is not yet as advanced as solid or liquid rocket technology. But it could be, given the effort to do so. I'd recommend a storable oxidizer like a strong acid, to get hypergolic ignition of a fuel grain that is a simple particulate-loaded rubber. Such storables can be pressure-expelled from bladders in tanks, eliminating their ullage problem, something we still cannot yet do with cryogens, even "soft" cryogens like liquid oxygen.
GW

So can we also use a classic pressure-feed liquid propellant NTO-MMH with bladder-tank as an ullage motor?

Last edited by Quaoar (2018-07-03 16:34:09)

SpaceNut · 2018-07-03 16:40:55

Was wondering if we had a Falcon 9 with a dragon outfited with life support of a crew of 2, what could we expect if we were capable of refueling it on mars surface.

The other part of my question would be to change the fuel to storables of course with the coirrect engines fully loaded on orbit would it have enough fuel to relaunch back to orbit after landing.

Or just point me to a simulator or the equations to use....

GW Johnson · 2018-07-04 07:55:35

To answer Quaoar:

Sure, any storable-propellant system that is expelled by gas pressure from a bladder inside the tank can be used as an ullage thruster. Most hydrazine monopropellant or hydrazine-NTO systems are easily capable of this function. And have been used in this way.

You can expel from the bladder at low pressures, and pump it up to high pressures for the thruster's chamber, or you can expel at the injection pressure at the cost of a heavier, high-pressure tank. The high pressure tank approach gets you the instant response you want in an attitude, etc., thruster.

If it's a one-time deal (like the Saturn stages), the solid cartridge is lighter and cheaper. But it's a one-shot device. You would have to have a set of them dedicated to each relight that you want. More than just one or two relights, and the bladder-expelled storables are the better deal.

To answer Spacenut:

To reach low Mars orbit from the surface, you need the delta-vee to reach orbital velocity (about 3.6 km/s) plus a tad extra for a small drag loss (very low density) and a gravity loss reduced by Mars's lower gravity. At most you would need 4 km/s deliverable capability, and probably only a little less, say 3.7 or 3.8 km/s. More if you need on-orbit maneuvering, such as to rendezvous. So, go with 4 km/s as a decent figure of merit.

At MMH-NTO Isp (~330 sec), the mass ratio required to reach 4 km/s is 3.44, something easily achievable in a single stage, even carrying payload. It corresponds to a vehicle whose propellant is about 71% of its ignition weight. If the inert mass fraction is 10-15% for something with landing legs and a heat shield, there is still 14-19% mass fraction left over for payload. Biggest problem is there's no way currently to make MMH and NTO on Mars.

The crewed Dragon capsule itself does not not come anywhere close to that mass ratio. Its empty weight is around 5-6 tons, and it has propellant tanks for only something like 1.5 tons. There is no way to refuel a Dragon on Mars's surface and reach orbit on its Super Draco's. There is more-than-enough delta-vee in either Falcon-9 stage, although not at the thrust levels you would really want, but there's no easy way to land such a thing.

But, you only need about 0.7 to 1 km/s delta-vee to land from Mars orbit, or even direct from an interplanetary trajectory, because the bulk of the deceleration delta-vee is done with a heat shield: aerobraking. What that really means is a chemically-fueled single-stage two-way reusable lander is possible. The mass ratio at MMH-NTO Isp for 5 km/s delta vee is 4.69, something still achievable in a single stage.

Such a vehicle would be 79% propellant. If 10-15% inert (remember, landing legs plus a heat shield, plus enough enclosed volume to carry bulky payload), that leaves 7-11% payload. Smallish, but practical. If you base from low Mars orbit, design the thing for refilling on orbit, and send propellant from Earth, you can make multiple excursions to the surface and back, with a single lander vehicle. For a 10 ton payload, the vehicle you design will fall in the 91-143 ton range of ignition weight.

This kind of thing really is technically feasible right now. It even looks a tad better using LOX and LCH4, which might well be produced on Mars, given a site with exploitable ice deposits.

GW

Last edited by GW Johnson (2018-07-04 08:26:09)

SpaceNut · 2018-07-04 16:16:10

Thanks for the reply GW.
Was looking at the dragon not as a use of its draco rockets as we have had your analysis on that capability in the past just for landing but was only for a crew to make use of it as a mars lander with the cargo area under the capsule for supplies with a reuseable stage with legs under it. From the info the first stage of a Falcon 9 is over kill for mars so a shorter wider stage would need to be developed if we did go that route of using a man rated dragon with a single stage lander design for mars.

SpaceNut · 2018-07-04 16:51:30

MMH is when used with LOX another possible combination. https://en.wikipedia.org/wiki/Monomethylhydrazine

This is one that you bring half of the fuel and make Insitu oxygen for the return.

ADVANCED HIGH PRESSURE ENGINE STUDY

Fuel table thespacerace.com/forum

High Energy-Density Liquid Rocket Fuel Performance

kbd512 · 2018-07-04 22:43:51

GW,

Settling of cryogens (LN2) were experimented with aboard STS or ISS (I forget which now). They rotated the propellant tank / vehicle about its axis, that throws the blob of propellant against the side walls of the propellant tank, and then the liquid blob is in the correct position to start feeding the engines. After you start the engine, that pushes the propellant blob towards the rear of the tank, and then you're in business.

The IVF concept implemented by ULA implements ullage functionality by using accumulation of the gaseous O2/H2 boil-off in cylinders that feed O2/H2 thrusters used to settle the propellant blobs prior to main engine start. The piston-powered combustion engine uses waste heat dumped into recirculated propellant to pressurize the propellant tanks.

Both concepts have experimentally been proven to work. There's no need for entirely separate systems for reaction control, ullage, and propellant tank pressurization. The mass savings associated with not using separate systems is so great that it's almost the same payload mass benefit as adding another solid rocket booster to Atlas V.

The Russians and Chinese use NTO/MMH in some of their boosters and upper stages, but so far as I know, we don't. We'd need engines with RD-275M thrust levels to power BFS.

NPO Energomash RD-275M (SC; NTO / UDMH): 1832kN and 315.8s vac
SpaceX Raptor (FFSC; LOX / LCH4): 1900kN and 375s to 382s vac

Perhaps the propellant tanks would be a bit smaller, but perhaps not since 59s to 66s (20%?) is a significant drop in Isp. At what point does the drop in Isp start making the upper stage so heavy that there's not much useful payload left?

Quaoar · 2018-07-05 03:56:38

GW Johnson wrote:

To answer Quaoar:
Sure, any storable-propellant system that is expelled by gas pressure from a bladder inside the tank can be used as an ullage thruster. Most hydrazine monopropellant or hydrazine-NTO systems are easily capable of this function. And have been used in this way.
You can expel from the bladder at low pressures, and pump it up to high pressures for the thruster's chamber, or you can expel at the injection pressure at the cost of a heavier, high-pressure tank. The high pressure tank approach gets you the instant response you want in an attitude, etc., thruster.
If it's a one-time deal (like the Saturn stages), the solid cartridge is lighter and cheaper. But it's a one-shot device. You would have to have a set of them dedicated to each relight that you want. More than just one or two relights, and the bladder-expelled storables are the better deal.
GW

Thanks GW,

I also discovered the existence of electric-feed rockets

https://en.wikipedia.org/wiki/Electric-pump-fed_engine

https://en.wikipedia.org/wiki/Rutherfor … et_engine)

May they be useful for a Mars mission, given that a manned spaceship with life support needs battery and solar panels anyway?

Last edited by Quaoar (2018-07-05 03:58:30)

GW Johnson · 2018-07-05 07:34:48

Quaoar:

I had not seen the electric-pumped designs before, been out of the industry for a long time. Looks good except for the weight of the electrical gear. For an attitude/translation thruster system, increased weight of a small-percentage system is no big deal, and such, if present, can serve the ullage function, too. For a main engine system which is a bigger percentage of the weight, any weight addition is a bigger deal.

I think any of these ideas could easily serve multiple functions as thrusters and ullage for the main engines. The higher the thruster Isp, the smaller its tank can be for a given delta-vee budget. And then there's Kbd512's idea of using propellant boil-off gases as the cold-gas ullage thruster without adding any other thruster system. Same thing could serve both ullage and attitude/translational thruster functions.

Kbd512:

I think the rotational action you describe would only work at fuller tank levels with an off-center drain. The more off-center the drain, the lower the tank level can be and still work. Intriguing idea.

I think the propellant and engine choices being made by private commercial outfits like Spacex have less to do with technical optimization and more to do with practicality. It really simplifies all the development efforts and production logistics (and testing) if you use the same engine in your first stage as your second stage, just fitting a longer vacuum bell to the second stage engine. Of course that means the same propellants in each stage, usually.

Fitting an LH2/LOX upper stage to Falcon-9 or Falcon-Heavy might increase its payload (just a hunch, haven't run the numbers because I don't know the real airframe shape and volume constraints). Maybe. But if they can do an acceptable job with the same engine and propellants in both stages (and so far they can), with all the resulting simplification and reduced development, why not? It's just practical common sense to do so. And less expensive, too.

GW

Oldfart1939 · 2018-07-05 10:45:02

Re: choice of propellants; my concept stated elsewhere for a lander with takeoff capabilities selected MMH in combination with locally produced LOX. This combination has adequate Isp and a corresponding Id which allows for bringing along the actual fuel component as part of the landed mass. Oxidizer (LOX) comprises a large proportion of the subsequent takeoff mass (weight). Zubrin always promoted bringing along a supply of H2 for subsequent conversion to CH4. The MMH/LOX system is something of an eclectic solution.

SpaceNut · 2018-07-05 17:06:57

I remembered and its one way to be sure that man can go and come back with current technology without waiting another decade for better developed technology that still will be untested for mars conditions. Then you can bring all the gear for insitu manufacturing of water and methane which would lessen the cost of a second mission as you would not need to bring the fuel for the return to orbit with you to the surface which would also increase the tons to the surface for useable payloads in science or in habitat creation tools along with more insitu tools.

kbd512 · 2018-07-05 18:27:29

GW,

SpaceX says they want to bring back the upper stage of Falcon / Falcon Heavy using the HIAD technology developed by NASA. That requires more performance to prevent additional loss of payload capability. I'm fairly certain that a LOX/LH2 upper stage of the same diameter as the payload shroud, powered by the new RL-10 CECE or J-2X, when combined with IVF for a reduction in RCS / pressurization / power subsystem masses, would restore any payload mass lost to the recovery systems. There may even be a minor improvement in performance.

If SpaceX wants to demonstrate full reusability using their existing rockets while maintaining performance, then they need to embrace LOX/LH2. It's not a green-eyed monster. LH2 is obviously more onerous to store and use than RP-1, but Mr. Musk himself said that the cost of propellants when compared to the cost of throwing away the hardware or designing an entirely new rocket to achieve the desired performance is virtually meaningless. Much as the Russian space program has mastered the use of oxygen rich combustion and storable propellants, the American space program has mastered the use of LH2 for greatly improved Isp. It's not any sort of game changer, but it's clearly an enabling technology.

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Best propellant & stuff for a Mars spaceship and lander

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