Best propellant & stuff for a Mars spaceship and lander

Oldfart1939 · 2018-07-06 07:57:44

kbd512-

Once again the green eyed maiden of H2 with high Isp is seducing thoughts towards use in the Falcon 9/Falcon Heavy upper stages. In reality, the Id, or density corrected value is more relevant. This points directly towards either MMH or Aerozine 50. Neither of these has the cryogen handling issues associated with their use, along with a very high Id and corresponding Isp. Use of either option (H2 or MMH), requires construction and qualification of an entirely new engine by SpaceX, which in turn, would slow the entire program. I am of the opinion that SpaceX is stuck with either RP-1 or L CH4 for the foreseeable future.

GW Johnson · 2018-07-06 08:22:08

Strictly speaking, Isp is to be maximized if you are weight-limited only, no volume constraint. Density-Isp is maximized if you are volume-limited, with no weight constraint.

Real launch vehicles are both weight- and volume-limited, usually more strongly volume-limited than weight-limited, although not as strongly volume-limited as a tactical weapon design. And once upon a time, I sized a lot of those tactical weapon designs for a living (and I still can). I know exactly how performance optimization under severe constraints actually works.

I got started doing that in the slide rule days. It's easier with software, but more easily led astray by the garbage-in/garbage-out risk. (The biggest risk is not also satisfying kinematic thrust constraints all along the trajectory.) Been there and done that, many times.

That last (volume-limited) has to do with ascent aerodynamics, and not just brute drag. Stability and attitude control force-level requirements get very strongly involved, particularly when the payload shroud diameter exceeds the basic vehicle diameter-over-the-majority-of-its-length. So it's not a clear-cut choice, but density-impulse really is important.

However, technical issues are not the only determinant of propellant selections in stages. There are very strong practicality and economic issues, as I tried to indicate just above. The development and logistical simplifications of using the same engine in both stages (which means the same propellant selection), are enormous. So also the economic implications and consequences are enormous.

There's enough experience out there across many companies to indicate that the denser LOX-hydrocarbon/hydrocarbon-like propellants are far better in the first stage than the higher-Isp LOX-LH2 choice. Brute frontal thrust density is far more important early in the flight when weight and drag are so high. That's also exactly why solid strapons are so popular. You're initially flying straight up against the weight, and the drag. The gravity turn is mostly completed during the first stage burn in a two-stage launcher.

By the time of the second stage burn, you're flying nearly horizontal, and pretty much at zero drag. That's where Isp is far more important for generating lots of delta-vee in a far longer burn at lower frontal thrust density. This is where LOX-LH2 in the second stage is the better choice, as long as the fatter tankage volume can be tolerated in some way, such as a longer stage.

If your company happens to have a LOX-LH2 engine in its stable, or is willing to buy them from an outside source, then you are very likely to use LOX-hydrocarbon-of-some-kind in your first stage and LOX-LH2 in your second. This is exactly ULA's workhorse Atlas-5, and in its proposed Vulcan replacement.

If on the other hand, you only have LOX-hydrocarbon-of-some-kind engines within your company stable, and for non-technical competitive reasons you don't want to buy LOX-LH2 engines from companies you compete with, then you will show a strong tendency to use the same engines (and propellants) in both stages. THAT is exactly where (and why) Spacex is with Falcon-9/Falcon-Heavy, and where they are already going with the proposed BFR/BFS. I doubt very seriously that decision is going to change.

Government agencies like NASA and USAF are not constrained by competitive issues, other than protecting vendor trade secrets (which is not fairly implemented, by the way), so their designs most often feature LOX-LH2 in the second or final stage, and LOX-hydrocarbon in the first (or lower) stage(s). Examples: original Atlas-Centaur, and the LOX-LH2 third stage of the Saturn-5.

Just sayin', there's a lot more to it than just the Isp or density-Isp numbers.

GW

Last edited by GW Johnson (2018-07-06 08:44:06)

kbd512 · 2018-07-07 07:59:36

GW,

I'm rather fond of the green eyed maiden since she enabled our little forays to the moon and could also enable trips to Mars and Venus. The technology for long duration cryogen storage is already at an advanced state of development. In a few more years, it'll be ready. In the mean time, it's time to design some worthy payloads for these super heavy lift launch vehicles. By the time the confluence of technologies are ready for prime time, it'll be time to go.

My contention is that if there was ever a place where the increase in Isp was useful, it's in the upper stages. BFR is some ways down the road, so the time to test HIAD and upper stage reusability is right now.

Centaur Upper Stage
engine: 1 RL-10A or 1 RL-10C
dimensions: 3.05m D by 12.68m L
inert mass: 2,316kg
wet mass: 23,077kg

Falcon 9 / Falcon Heavy Upper Stage
engine: 1 Merlin 1D Vac
dimensions: 3.66m D by 13m L
inert mass: 3,900kg
wet mass: 96,750kg

Falcon 9 / Falcon Heavy Payload Fairing
diameter: 5.2m D

I want USAF to commission a 5.2m diameter RL-10 powered upper stage for Falcon 9 / Falcon Heavy. It can use 2 to 4 RL-10's and IVF. ULA's ACES is supposed to be 5m in diameter. I just want a slightly fatter ACES 41, or maybe just use ACES 41 without modification since it's almost the same mass as the Falcon upper stage, almost the same length as the Falcon upper stage, and almost the same diameter as the Falcon 9 / Falcon Heavy payload shroud. There should be no loss of payload performance, little increased drag, and stage recovery is the cherry on top for us to collect and reuse those expensive LOX/LH2 engines. We can snag it while it parachutes into the ocean using an USMC CH-53K King Stallion. It's well within that bird's payload capability.

Falcon 9 can deliver 22,800kg to LEO
Atlas V can deliver 20,520kg to LEO

I think we can bump that up a bit with the 140-ish increase in Isp / similar mass / similar dimensions / similar volume, prove recovery of upper stages, and demonstrate small propellant depots for lunar missions while we're at it.

That'd make Falcon 9 a 25t+ rocket and Falcon Heavy a 50t to 75t rocket. New Glenn can focus on being a 75t to 100t rocket. SLS can focus on being a 100t to 125t rocket. BFR can focus on being a 150t to 250t rocket (with LOX/LH2, of course). I think that'd round out our payload delivery capabilities rather nicely.

SpaceNut · 2018-07-07 11:30:24

https://en.wikipedia.org/wiki/Centaur_(rocket_stage)

https://en.wikipedia.org/wiki/Exploration_Upper_Stage

https://en.wikipedia.org/wiki/Earth_Departure_Stage

http://selenianboondocks.com/2013/06/ra … nt-page-1/

https://www.ulalaunch.com/docs/default- … ehicle.pdf

What make a centaur stage important is its already man rated and the other thing is its alrady optomized for the hydrogen boiloff issue where in that a fuel cell is used on the vented hydrogen to make power as part of the boiloff.
It is a tank within a tank used to control the rate of boiloff.

Aerojet Rocketdyne’s RL10 engine
https://www.nasaspaceflight.com/2018/05 … ve-vulcan/

GW Johnson · 2018-07-07 11:59:32

I think there's another acronym here to add to the acronym sticky under "human missions". Kbd512 uses "IVF" to mean something I am not familiar with, regarding somehow using boiloff gases from LH2-LOX propellant systems as ullage thrusters for ignition.

Most of the science and medical journals use "IVF" to mean "in vitro fertilization", and I know THAT'S not what he talking about. Like Musk, I don't like acronyms very much. They too often get misused to hide what one is really talking about. That's not the case here, because I really do have a vague idea what Kbd512 is talking about. But it's very vague.

GW

SpaceNut · 2018-07-07 12:56:37

IVF = ? Internal Vehicle Fuel ?

I did find it meantioned
https://www.reddit.com/r/SpaceXLounge/c … per_stage/

IVF Integrated Vehicle Fluids http://www.ulalaunch.com/uploads/docs/P … y_2012.pdf

https://www.reddit.com/r/spacex/comment … is_of_bfs/

That aside, you could manage with a VERY small population (a few hundred?) females, if you have a sperm bank and artificial insemination technology and/or IVF. The IVF may be especially useful on Mars because you could store the eggs and sperm in a rad-hardened container, fertilize and implant the zygote in a radiation-proof (underground, most likely) "nursery" where the mother stays for about the first 15 weeks of gestation.

kbd512 · 2018-07-08 13:13:26

GW,

IVF means Integrated Vehicle Fluids

IVF enabling technology is a de-rated / de-tuned 6 cylinder piston engine from a race car that has an electric generator and pump attached to it. The rest of the system includes accumulation tanks used to store compressed gaseous O2/H2 boil-off / siphon off for reaction control system or piston engine consumption and the O2/H2 reaction control thrusters that burn the compressed O2/H2 for attitude control. The piston engine heats the propellant to re-pressurize the tanks or siphons propellant to de-pressurize the tanks and store it in the gaseous O2/H2 for the IVF system's usage. There are no separate batteries, solar panels, reaction control systems, and helium tanks to pressurize the main tanks. Apart from substantial mass savings, the system permits extended duration storage of the propellant. Running the engine consumes less propellant than boil-off. ULA did the math on the mass savings after the entire system was engineered (after the actual hardware built and tested) and it's roughly equivalent to adding another solid rocket booster to the Atlas V core, with respect to payload performance achievable with Centaur.

If we can secure three Falcon Heavy flights, then we can ship my ITV concept to ISS for assembly. If we get LOX/LH2 upper stages with IVF, then we can prove HIAD recovery of the upper stages and IVF at the same time. Whenever BFR is ready to fly, the entire thing can be shipped to ISS. Anyway, that's the plan.

GW Johnson · 2018-07-11 10:41:30

Thanks, Kbd512. Between this and what you put in the acronym sticky under "human missions", I understand much better what this IVF thing is really all about. Odd how an internal combustion engine plus some sorts of compressors or pumps, and an electrical generator, is lighter than the things it replaces.

I understand "internal combustion engine", but I noticed that "spark ignition" was not specified. I assume this derated engine is spark ignition since both hydrogen and methane are high-octane materials. I did not see anything about how the higher flame temperatures of fuel plus oxygen versus fuel plus air get handled, but I assume that "derating" is probably compression reduction and retarded timing, both of which reduce flame temperatures.

To use this with low-octane/high-cetane RP-1 instead of methane or hydrogen, the derated compression ratio must be under about 3 or at most 3.5, to avoid detonation knock. Makes me wonder if ULA has anybody on their staff at either Boeing or Lockheed-Martin who really understands piston engines. In its original form, my farm tractor would burn either kerosene or gasoline in a spark ignition engine, but only with some serious extra intake heat on kerosene, and then only starting on gasoline. In that form, compression ratio was 3.5.

My old tractor is no longer in that form, it's current compression ratio is about 5.5 or 6. The extra heat shutter is corroded immobile, and has been for decades. This sort of thing was last produced about 1950 or 1955. I'm pretty sure ULA would have nobody on their staffs at their respective locations old enough to have any sort of experience with this kind of technology. I rather doubt they would have anybody with much piston experience of any kind, since the government pretty much made the transition to turbines and jet fuel in the 1970's.

Myself, I am old, and I really do know about piston engines, both spark-ignition and compression-ignition. That same old farm tractor I converted to run on E-85 alcohol fuel about a decade ago. The compression is too low to take advantage of the really high octane number, but it runs so cleanly now compared to gasoline, that my maintenance activities have drastically reduced.

Anyhow, thanks for telling us about IVF. That is an intriguing new use of an old technology.

GW

Last edited by GW Johnson (2018-07-11 10:43:29)

elderflower · 2018-07-12 03:05:24

You can control the combustion temperature and ameliorate the knocking problem by recirculating exhaust gas (steam in the case of a hydrogen/oxygen engine). You would need to preheat before startup.
Effective lubrication might be very difficult.
You probably need to expand the unrecirculated exhaust gas through a turbine to extract more power and to reduce the thrust on your vehicle from the exhaust stream.

kbd512 · 2018-07-12 07:08:33

GW,

The engine is substantially de-rated and simplified. The original in-line six cylinder engine with header / valves / etc, was making 100's of kilowatts of power in racing form, but the ULA engine is reduced to 21kW peak output with 3kW nominal output since that's what's necessary for normal operations. ULA's variant is a low-RPM flat head more suitable for a small tractor than a race car. The high heat carrying capacity of Hydrogen doesn't hurt, either.

GW Johnson · 2018-07-12 12:20:39

Between greatly-lowered compression ratio, strongly-retarded timing, and reduced max throughput massflow, the flame temperature might well be reduced around 1000 F. That would be very consistent with the low power levels Kbd512 talks about. If the compression is low enough, you could do this spark ignition with LOX-RP-1, although the RP-1 will be liquid injection, not vapor like CH2 and H2. It's a vastly different fuel control, with the same oxygen control for all.

GW

SpaceNut · 2018-07-12 17:39:41

Auxilary power system to power the ship through out its flight and is simple as it a recharge to battery power and gives primary power while in operation.

Quaoar · 2018-07-13 06:31:33

GW Johnson wrote:

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.

GW

Thanks GW,
How does exactly work the piston rocket?
A working fluid in a closed cycle cools the rocket becoming hot, then it moves the pistons of the pump, it is cooled by the incoming propellant via an heat-exchanger, and it comes back again to cool the rocket?

Last edited by Quaoar (2018-07-13 08:26:45)

GW Johnson · 2018-07-13 14:36:53

Hi Quaoar:

As best I can tell, the piston pump was driven by compressed gas in the two airplanes that XCOR flew under rocket power. These were quite small rockets, and flight durations were limited.

XOR was working on a subscale demonstrator for replacing the RL-10 in the Centaur upper stage. That engine nominally is 25,000 lb thrust, the subscale demonstrator was 2500 lb thrust. Their piston-pumped full scale design, a paper design only, was 25,000 to 50,000 lb thrust.

The subscale demonstrator piston pump was tested and qualified to handle liquid hydrogen fuel. I'm not sure if the heat engine to drive that pump was demonstrated in that same subscale engine, but I know they tested it in something and it worked.

The third fluid is the liquid coolant for the rocket engine parts. It flashes to vapor, which then drives a piston engine device providing shaft power to the piston pumps. After driving the piston engine, this fluid re-liquifies, and is pumped back up to pressure for recirculating into the coolant passages in the rocket engine. The shaft power not only operates the piston pumps for fuel and oxidizer, but also the third-fluid coolant pump.

The whole heat engine and piston pump device was all one package operating with one common crankshaft. I actually saw this hardware before they made the tests with it. It was about the size of a medium motorcycle engine.

GW

Last edited by GW Johnson (2018-07-13 14:37:51)

Quaoar · 2018-07-14 09:47:22

GW Johnson wrote:

After driving the piston engine, this fluid re-liquifies, and is pumped back up to pressure for recirculating into the coolant passages in the rocket engine. GW

To be re-liquified, the fluid has to be cooled, I suppose. Is the fluid cooled regeneratively in some kind of heat-exchanger by the incoming propellant?

GW Johnson · 2018-07-15 09:13:43

Hi Quaoar:

I honestly don't know the answer to your question. When expanding drive gas to do work in a heat engine, there is some drop in temperature as the pressure drops. Whether that is enough to reliquify in this particular design, I don't know. If not, you either need a heat exchange radiator to cool it to reliquifaction, or else you need to compress gas instead of liquid back up to the engine-cooling pressure, which requires a lot more power and involves heavier equipment.

GW

Quaoar · 2018-07-19 16:25:52

GW Johnson wrote:

Hi Quaoar:
I honestly don't know the answer to your question. When expanding drive gas to do work in a heat engine, there is some drop in temperature as the pressure drops. Whether that is enough to reliquify in this particular design, I don't know. If not, you either need a heat exchange radiator to cool it to reliquifaction, or else you need to compress gas instead of liquid back up to the engine-cooling pressure, which requires a lot more power and involves heavier equipment.
GW

So, to recap if you have to project an orbit-to-orbit spaceship to Mars, which has to be reused many times, which kind of rocket would you use: expander cycle, gas generator, pressure fed, pistonless pressure pump or piston pump?

kbd512 · 2018-07-19 19:13:31

Quaoar,

For whatever it's worth, ion engines have already operated continuously for years without maintenance of any kind. There is no chemical rocket engine that comes close to touching that reliability figure. By their very nature, absent considerable maintenance, all combustion engines are limited lifetime propulsion systems. The greater the power output level, the shorter the lifetime and the more involved the maintenance operations become. I'm not sure there's any way around that.

With 600kWe to 700kWe class ion engines, the number of days it takes to get to Mars, including spiral out / spiral in time, is no more than the number of days BFS requires. Argon may require a little more input power than Xenon, but it's significantly more affordable and no new development is required. The thin film arrays from Ascent Solar have already demonstrated 1400W/kg in LEO. If a ship required 1MWe of input electrical power, then that's about 1.5t worth of hardware using multi-stranded array tethering materials and gyroscopic deployment and stabilization of a circular array.

Quaoar · 2018-07-20 10:45:25

kbd512 wrote:

Quaoar,
For whatever it's worth, ion engines have already operated continuously for years without maintenance of any kind. There is no chemical rocket engine that comes close to touching that reliability figure. By their very nature, absent considerable maintenance, all combustion engines are limited lifetime propulsion systems. The greater the power output level, the shorter the lifetime and the more involved the maintenance operations become. I'm not sure there's any way around that.
With 600kWe to 700kWe class ion engines, the number of days it takes to get to Mars, including spiral out / spiral in time, is no more than the number of days BFS requires. Argon may require a little more input power than Xenon, but it's significantly more affordable and no new development is required. The thin film arrays from Ascent Solar have already demonstrated 1400W/kg in LEO. If a ship required 1MWe of input electrical power, then that's about 1.5t worth of hardware using multi-stranded array tethering materials and gyroscopic deployment and stabilization of a circular array.

It depends on the ion engine exhaust velocity and the mass of the spaceship. For example, an hypothetical NSTAR-like ion engine, with an exhaust velocity of about 30 km/s, with 1 MWe of power has a propellant flux of about 0.0022 kg/s (2E/V^2) and a thrust of about 66 N. The non impulsive delta-V transfer from Earth to Mars is about 10 km/s, so a round trip has a delta-V of about 20 km/s, so a ship with the aforementioned ion engine needs a Mass ratio of about 2.
Let's take a BFR-like spaceship with a inert mass of 85 ton and 150 ton of payload (lander + supplies): she has a dry mass of 235 ton, 235 tons of propellant and a wet mass of 470 tons.
A 470 ton spaceship with a 66 N ion-engine has an acceleration of 0.00014 m/s2, so to gain 10000 m/s of delta-V, she needs 71428571 seconds or 826 days. 826 days are more than 2 years: a much longer time than the almost 8.5 months of a Hohmann impulsive transfer.

To have a BFR-like transit time you needs at least a 3-5 MWe ion engine but I think there are no engine of this class ready to use by now.

Last edited by Quaoar (2018-07-20 11:03:21)

GW Johnson · 2018-07-20 11:30:05

My understanding is that the largest existing ion engines are KW-class devices, when 100KW to MW-class devices are needed to push large objects. They can be developed, they just haven't been, yet. There's no science or technology lack that prevents it.

As for a long life chemical rocket engine, I think I'd use the piston-pump technology XCOR tested in sizes up to 2500-5000 lb thrust. That kind of pump just lasts a whole lot longer than a hot-gas-driven turbopump, although it is a little heavier. If the propellants have boiloff issues, you can use boiloff gases in an IC engine the way Kbd512 describes, to develop the shaft horsepower to drive the piston pumps.

XCOR said they couldn't scale their design past about 25,000-50,000 lb thrust, but that's with the 3rd fluid coolant as drive medium for their heat engine. I don't know for sure, but I think their cooling/heat engine system had the scaling difficulty, not the basic piston pump. To the best of my knowledge, you can build those as big as you want.

GW

kbd512 · 2018-07-20 11:33:22

Quaoar,

We've tested MPD thrusters up to 1.5MWe at universities and NASA has tested MPD thrusters up to 1MWe. Perhaps we'd need multiple engines, but the point is that it can be done. The Ascent Solar folks have tested thin film arrays of up to 2.25kW/kg in their labs. The tonnage involved with a multi-MW-class thin film solar array is clearly not a major problem. It's an engineering problem to be sure, but the basic 1.4kW/kg thin film solar technology has already flown in LEO for more than 6 months in preparation for use on JAXA's Jupiter solar sail probe.

If the multi-MW thrusters were problematic, although clearly it's just a matter of feeding in enough power, then an array of 25 of the 200kW X3 thrusters would still work. The X3's would weigh about 5.6t, whereas 5 of the MW-class MPD thrusters would weigh than 1t, but it's still feasible.

The propellant and launch costs savings are considerable, to say the least.

kbd512 · 2018-07-20 12:04:39

GW,

I already posted a link to the testing of Aerojet-Rocketdyne's X3 thruster. That thruster is supposed to be run at 200kW before year's end. So far as I know, Aerojet-Rocketdyne doesn't make paper rocket engines. The MW class MPD thrusters I've alluded to have been tested in a variety of universities across the US and Europe.

Here's another paper on Aerojet-Rocketydne's X3:

100 kW Nested Hall Thruster System Development

Quaoar · 2018-07-20 12:19:13

kbd512 wrote:

Quaoar,
We've tested MPD thrusters up to 1.5MWe at universities and NASA has tested MPD thrusters up to 1MWe. Perhaps we'd need multiple engines, but the point is that it can be done. The Ascent Solar folks have tested thin film arrays of up to 2.25kW/kg in their labs. The tonnage involved with a multi-MW-class thin film solar array is clearly not a major problem. It's an engineering problem to be sure, but the basic 1.4kW/kg thin film solar technology has already flown in LEO for more than 6 months in preparation for use on JAXA's Jupiter solar sail probe.
If the multi-MW thrusters were problematic, although clearly it's just a matter of feeding in enough power, then an array of 25 of the 200kW X3 thrusters would still work. The X3's would weigh about 5.6t, whereas 5 of the MW-class MPD thrusters would weigh than 1t, but it's still feasible.
The propellant and launch costs savings are considerable, to say the least.

Interesting. Which is the power/weight ratio of the ion engines?

kbd512 · 2018-07-20 13:09:53

The X3 thruster weighs 227kg and produced 5.4N of thrust at that steady state power level. It's full input capability is 200kW, but that's for "low gear", like the gears in a transmission. The 200kW run is expected to produce a bit more than double that figure. Let's just call that 54N/MW. Aerojet-Rocketdyne has a design for a version of the X3 technology with more channels (gears) that weighs 320kg and accepts 1.2MW to 1.4MW of input power and projected produce 36N to 216N of thrust (output does not scale linearly with input power in these electric thrusters and also increases as input voltage increases). X3 and all variations of X3 are NHT's or Nested Hall Thrusters with concentric ring channels. No major design hurdles had to be cleared to get X3 to work.

Edit: Isp ranges between 1200s for low gear and 3800s for high gear

The internet is littered with research papers about MW-class MPD thrusters that were run by universities or NASA or other space exploration agencies. Google "MW class MPD thruster" and start reading.

MW-class MPD thrusters from:
NASA Glenn Research Center
Georgia Tech (variant closely associated with the one used at NASA GRC)
various universities in Japan, Europe, and Russia

Last edited by kbd512 (2018-07-20 13:14:53)

Quaoar · 2018-07-22 03:31:32

kbd512 wrote:

The X3 thruster weighs 227kg and produced 5.4N of thrust at that steady state power level. It's full input capability is 200kW, but that's for "low gear", like the gears in a transmission. The 200kW run is expected to produce a bit more than double that figure. Let's just call that 54N/MW. Aerojet-Rocketdyne has a design for a version of the X3 technology with more channels (gears) that weighs 320kg and accepts 1.2MW to 1.4MW of input power and projected produce 36N to 216N of thrust (output does not scale linearly with input power in these electric thrusters and also increases as input voltage increases). X3 and all variations of X3 are NHT's or Nested Hall Thrusters with concentric ring channels. No major design hurdles had to be cleared to get X3 to work.
Edit: Isp ranges between 1200s for low gear and 3800s for high gear
The internet is littered with research papers about MW-class MPD thrusters that were run by universities or NASA or other space exploration agencies. Google "MW class MPD thruster" and start reading.
MW-class MPD thrusters from:
NASA Glenn Research Center
Georgia Tech (variant closely associated with the one used at NASA GRC)
various universities in Japan, Europe, and Russia

There are some issues whit electric propulsion: a 5 MW solar panel array, considering a very optimistic triple junction panel with 40% of efficiency and a mean solar irradiance of 1000 W/m2, must have a surface of 12500 square meters, and that huge surface has to continuously face the sun during propulsion. The best way to do it, avoiding shadowing between panels, is to split the surface in two 80-meter-sided squares and mount them on a rotating platform, that can be oriented to the sun regardless of the thrust direction. This is not impossible to do, but is not very simple - given that now we have no more Space Shuttle and Canadarm - and even if the panels are light, the whole structure - panel arrays plus support and motors - can be heavier than 1400W/kg.
Another issue is the artificial gravity: a chemical rocket or a NTR-propelled space ship can be made as a GW's rigid baton and can spin while coasting, providing artificial gravity to the astronauts. A solar electric spaceship has to use the ion engines for the whole part of her trip, so if you want artificial gravity you are forced to build a ship with a separate spinning section. Event this is not impossible but is difficult, massive and expensive.
For those reasons I think a chemical or NTR spaceship is best suited for the crew, while a solar electric one may be useful as a cargo-tug.

Last edited by Quaoar (2018-07-22 03:44:14)

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Re: Best propellant & stuff for a Mars spaceship and lander

#47 2018-07-20 12:04:39

Re: Best propellant & stuff for a Mars spaceship and lander

#48 2018-07-20 12:19:13

Re: Best propellant & stuff for a Mars spaceship and lander

#49 2018-07-20 13:09:53

Re: Best propellant & stuff for a Mars spaceship and lander

#50 2018-07-22 03:31:32

Re: Best propellant & stuff for a Mars spaceship and lander

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