VASIMR - Solar Powered?

GW Johnson · 2015-05-27 09:36:40

I don't know a lot about electric propulsion. Just enough to know there's more than one kind, and all of them are very low thrust for the weight, no matter how you define the weight. But Isp is in the 5000 sec class.

The discussions here have taught me that the ones with the lowest power requirements are the ones that use the most expensive propellant (xenon). Lower power reduces weight. But there is interest in using something besides xenon.

I have previously suggested that a vehicle with both conventional and electric engines might be the best choice for manned travel. You can get some extra delta-vee efficiency at high thrust for escape and capture with conventional, and you can use the slow buildup of delta-vee at low thrust but super-high Isp during transits with electric. Does not that idea make sense?

If you use LH2-LOX for conventional propulsion, could you not store your propellant as frozen water and avoid cryogenic boil-off issues? Use solar thermal to melt it, and solar photovoltaic to electrolyze it, then more solar PV to liquify it. There's 8-9 months in transit to Mars to make from water the LH2-LOX for the capture burn into LMO. There's similar time scales available to make more in LMO for departure, and during the transit home for a capture burn into LEO.

The ratio from electrolysis is wrong, we have too much oxygen. The excess oxygen could go to life support. And maybe toward electric propulsion propellant? Could this be done? Has it been done? Is an oxygen electric thruster feasible? If so, could one be made spaceworthy for application ion about a decade?

I wonder about these concepts, but know too little about electric propulsion and solar PV to evaluate them.

GW

Terraformer · 2015-05-28 06:22:53

Well, there's a suggested microwave based thruster that uses water. Isp in the 800s area.

SpaceNut · 2015-05-28 18:58:34

I saw that quite some time ago and the reason is the energy to create the force no matter how much you apply will not increase the force beyond that point and water will react with the materials erroding the piping and more as its processed on its way to an exit.

From wikipedia:

The noble gases make a group of chemical elements with similar properties: under standard conditions, they are all odorless, colorless, monatomic gases with very low chemical reactivity. The six noble gases that occur naturally are helium (He), neon (Ne), argon (Ar), krypton (Kr), xenon (Xe), and the radioactive radon (Rn).
For the first six periods of the periodic table, the noble gases are exactly the members of group 18 of the periodic table. It is possible that due to relativistic effects, the group 14 element flerovium exhibits some noble-gas-like properties, instead of the group 18 element ununoctium. Noble gases are typically highly unreactive except when under particular extreme conditions. The inertness of noble gases makes them very suitable in applications where reactions are not wanted.

So while there might be the possibility of a noble gas with more mass to make more force with the cost to make it as its short lived does not seem to be the answer....

Terraformer · 2015-05-29 12:11:13

As far as noble gases go, Argon has the major advantage that it's very abundant.

RGClark · 2015-06-03 16:06:13

JoshNH4H wrote:

RGClark-
Please note that those numbers were given as examples and don't represent real values. The real values can be found in the following table:
https://gammafactor.files.wordpress.com … table1.jpg
I note that these nested hall thrusters are every bit as theoretical as a high power VASIMR thruster, and those numbers don't necessarily represent the mass of any engine that will ever be built.

I think GW's suggestion to use also chemical propulsion first before turning on the electric propulsion would be useful. For instance the low thrust of EP means you have to take a spiraling trajectory out of Earth's orbit. This takes more time and therefore the high speeds as shown in your table would be needed to make up for that to achieve the short travel time.

However, if you use chemical propulsion first to get out of Earth's deep gravity well then you would be just adding on the speed attained by the EP onto the velocity you would have in the Earth-Mars Hohmann transfer trajectory you attained from the chemical propulsion.

Quite likely also you would not want to depart from Earth's orbit but from a station at L2. Note that NASA just announced it wants to set a station at L2 to serve as a staging point for further BEO missions. I think the delta-v to be put on a Mars transfer trajectory in that case would only be about .9 km/s, which could easily be supplied by chemical propulsion, compared to about 3.8 km/s when departing from Earth.

You might also want to use chemical propulsion to make the trajectory closer to straight-line rather than the elliptical Hohmann trajectory. From memory I don't think this would need to be very large, about 6.5 km/s. Then the one-way delta-v that would need to be supplied by the EP would be much less than in your table.

For instance let's say chemical propulsion made the trajectory nearly straight-line and you wanted the one-way travel time to be 39 days using EP. Depending on the year, the closest Mars gets to Earth is about 60 million km. A 39 day flight is 39*24*3600 = 3,369,600 seconds. So this would require a speed of 17.8 km/s. But 6.5 km/s was already supplied by the chemical propulsion so it would actually be 11.3 km/s supplied by the EP. However, it would take some days also for the EP to build up to this speed so you would need somewhat higher speed than this to make the travel time be 39 days.

Note though this is not including slowdown time or the extra delta-v that would require. You would need then a high efficiency heat shield that would work for the high arrival speed in such a scenario. The inflatable heat shields NASA is testing now for Mars missions might work or the magnetoshell heat shields formed by magnetic fields might do it:

Magnetic bubble may give space probes a soft landing.
03 July 2014 by Robin Hague
http://www.newscientist.com/article/mg2 … nding.html

Bob Clark

Last edited by RGClark (2015-06-03 17:13:35)

RGClark · 2015-06-05 12:18:39

Interesting report here:

Solar Electric Propulsion.
Technology Development.
www.nasa.gov/sites/default/files/files/CTaylor_SEP.pdf

It discusses the idea of using chemical propulsion for getting out of Earth's gravity well then using SEP to reach a BEO destination. Note this is how most interplanetary probes that used electric propulsion worked. The report has this surprising chart near the end:

It includes the interesting fact that using the same propulsion system at 50 kW power, a 2,000 kg dry mass spacecraft would take 91 days to make the flight to Mars when running at 4,000 s Isp, but only 23 days(!) when running at 2,000 s Isp. This make sense because the thrust drops off for EP thrusters as you increase the Isp. But the thing is we already have a 50 kW Hall effect thruster in the NASA-457m:

Performance Test Results of the NASA-457M v2
Hall Thruster.
http://ntrs.nasa.gov/archive/nasa/casi. … 014613.pdf

I was puzzled at first by the graphic on the side listing delta-v's to reach various interplanetary destinations. For Mars, it's listed as 5.6 km/s. Since the report is discussing using EP after leaving Earth's vicinity, this must mean the delta-v just to leave Earth's position in the Solar System to Mars position, so not including the delta-v for escape velocity or slowdown delta-v at Mars.

Support for this is suggested by this:

Hohmann transfer orbit

Low-thrust transfer.
It can be shown that going from one circular orbit to another by gradually changing the radius costs a delta-v of simply the absolute value of the difference between the two speeds. Thus for the geostationary transfer orbit 7.7 − 3.07 = 4.66 km/s, the same as, in the absence of gravity, the deceleration would cost. In fact, acceleration is applied to compensate half of the deceleration due to moving outward. Therefore the acceleration due to thrust is equal to the deceleration due to the combined effect of thrust and gravity.

http://en.wikipedia.org/wiki/Hohmann_tr … t_transfer

Bob Clark

Last edited by RGClark (2015-06-05 14:35:03)

RGClark · 2015-06-05 14:37:17

RGClark wrote:

Interesting report here:
Solar Electric Propulsion.
Technology Development.
www.nasa.gov/sites/default/files/files/CTaylor_SEP.pdf
It discusses the idea of using chemical propulsion for getting out of Earth's gravity well then using SEP to reach a BEO destination. Note this is how most interplanetary probes that used electric propulsion worked. The report has this surprising chart near the end:
http://oi58.tinypic.com/zt99qr.jpg
It includes the interesting fact that using the same propulsion system at 50 kW power, a 2,000 kg dry mass spacecraft would take 91 days to make the flight to Mars when running at 4,000 s Isp, but only 23 days(!) when running at 2,000 s Isp. This make sense because the thrust drops off for EP thrusters as you increase the Isp.

Hmmm. After running some numbers I wonder if that 23 day number is just the length of time it takes for the vehicle to get up to speed, i.e., it does not include the coast time.

Bob Clark

RGClark · 2015-08-17 09:05:45

I was interested to read the report, "NUCLEAR THERMAL ROCKET/VEHICLE CHARACTERISTICS AND SENSITIVITY TRADES FOR NASA’s MARS DESIGN REFERENCE ARCHITECTURE (DRA) 5.0 STUDY", http://ntrs.nasa.gov/archive/nasa/casi. … 012928.pdf. But, it seems to imply specific power, power to weight, far above what has been achieved for known space nuclear reactors.

The engines described in the report have a thrust of 15,000 lbs and an Isp of 900 s. The thrust, or jet, power generated by an engine can be calculated as: (1/2)*thrust*exhaust_velocity. The thrust in Newtons for one engine is 15,000 lbs*4.46 N/lbs = 67,000 N. The exhaust velocity in m/s is 900 s * 9.81 m/s2 = 8,800 m/s. So the power generated by one engine is .5*67,000*8,800 = 295,000,000 watts. Now the weight for this engine, which I assume includes the reactor weight, can be calculated from the thrust/weight ratio cited in the report of 3.43 to be 19,500 N, 1,990 kg. But this correspond to a specific power of 148,000 watts/kg (!) Actually since there would not be perfect efficiency in turning the reactor output to thrust, the real number is even higher than this. But this is orders of magnitude beyond what has been done with other space nuclear reactors. See for example the numbers here:

Nuclear Reactors and Radioisotopes for Space.
(Updated July 2015)

http://www.world-nuclear.org/info/Non-P … for-Space/

Perhaps they are not including the weight of the reactor when quoting the T/W ratio? If so, then I've seen nowhere where this weight for the space nuclear power for this particular engine is specified separately.

Bob Clark

Last edited by RGClark (2015-08-17 09:07:06)

GW Johnson · 2015-08-17 11:17:33

Hi Bob:

Hope you're enjoying the convention.

That T/W includes the reactor. That's the best of the NERVA designs from the 1970's.

I'm not sure about calculating jet "power" the way you did, or how that might relate to the fundamentally thermal power generated by the reactor. But energy is conserved, so you should be able to relate reactor thermal power to the KE-rate of the exhaust stream: 0.5 massflow rate * jet velocity squared.

Typically, these engine cores run very much hotter than those in any of the space power reactor systems. So getting far higher power per unit mass out of a nuclear rocket engine than out of an electric power system should not be all that unexpected. That also shows up in expected useful life: very limited for an engine, quite long for an electric power device.

GW

RGClark · 2015-08-17 11:54:18

GW Johnson wrote:

Hi Bob:
Hope you're enjoying the convention.
That T/W includes the reactor. That's the best of the NERVA designs from the 1970's.
I'm not sure about calculating jet "power" the way you did, or how that might relate to the fundamentally thermal power generated by the reactor. But energy is conserved, so you should be able to relate reactor thermal power to the KE-rate of the exhaust stream: 0.5 massflow rate * jet velocity squared.
Typically, these engine cores run very much hotter than those in any of the space power reactor systems. So getting far higher power per unit mass out of a nuclear rocket engine than out of an electric power system should not be all that unexpected. That also shows up in expected useful life: very limited for an engine, quite long for an electric power device.
GW

Thanks. The conference was a lot of fun. The formula I used is equivalent to the one you cited. I just wanted to save having to calculate the massflow rate. That the two are equivalent comes from the fact that thrust = (massflow rate)*(exhaust velocity). So:
power = .5*(massflow rate)*(exhaust velocity)*(exhaust velocity) = .5*(thrust)*(exhaust velocity).

Bob Clark

RGClark · 2015-08-24 11:03:20

RGClark wrote:

The engines described in the report have a thrust of 15,000 lbs and an Isp of 900 s. The thrust, or jet, power generated by an engine can be calculated as: (1/2)*thrust*exhaust_velocity. The thrust in Newtons for one engine is 15,000 lbs*4.46 N/lbs = 67,000 N. The exhaust velocity in m/s is 900 s * 9.81 m/s2 = 8,800 m/s. So the power generated by one engine is .5*67,000*8,800 = 295,000,000 watts. Now the weight for this engine, which I assume includes the reactor weight, can be calculated from the thrust/weight ratio cited in the report of 3.43 to be 19,500 N, 1,990 kg. But this correspond to a specific power of 148,000 watts/kg (!) Actually since there would not be perfect efficiency in turning the reactor output to thrust, the real number is even higher than this. But this is orders of magnitude beyond what has been done with other space nuclear reactors.

I'm informed by a NASA nuclear rocket engineer that the high number of 148,000 watts/kg really is in the range of the power-to-weight ratio of nuclear rocket engines. But this means when you do the conversion to electrical power you reduce the efficiency by three orders of magnitude! In other words the lack of efficiency really does not have to do with the nuclear space reactor itself.
One problem with the conversion to electric power is poor efficiency methods are used. For instance for the American systems thermoelectric conversion is used, which typically is only 5% to 8% efficient.

But for the power to weight ratio to drop so low it must be the additional weight of the electrical equipment that contributes to it. And indeed the specific power of electric motors is at most in the 10,000 watts/kg range:

Power-to-weight ratio.
2.1.2 Electric motors/Electromotive generators
https://en.wikipedia.org/wiki/Power-to- … generators

Still if we used an electric motor at this efficiency level, run in reverse to generate electricity, then it should give the needed specific power.

Bob Clark

Last edited by RGClark (2015-08-24 11:54:30)

RGClark · 2015-08-28 12:20:01

39 days to Mars possible now with nuclear-powered VASIMR:

Nuclear powered VASIMR and plasma propulsion doable now.
http://exoscientist.blogspot.com/2015/0 … lasma.html

A criticism of VASIMR plasma drive was that space nuclear power did not have sufficient power at the needed lightweight. However, it turns out that this is due to the heavy electrical generating equipment, not the nuclear reactors themselves.

Then note that recent research has produced electrical generators at the needed lightweight, thus making nuclear-powered VASIMR viable.

Bob Clark

Void · 2015-08-28 13:12:47

Nasa seems interested in the VASIMR
http://nasawatch.com/archives/2015/08/n … d-ast.html

For what it is worth;
I know that Bigelow has been waiting for LEO launch vehicles, otherwise they were ready quite some time ago.
Because of this and also because of the potential advances in propulsion that you have brought up,
and because of Musks intentions for Mars,

I do not think NASA's hardware is as non-adapted as it might seem. They were in a crippling situation, but the whole collection of developments might actually get the human race moving to Mars and elsewhere.

Antius · 2017-04-04 04:14:46

From what I have read on VASIMR, the thruster uses microwaves to produce cold plasma, which is then accelerated magnetohydrodynamically.

This concept could work extremely well if the propellant were wastes or unprocessed natural materials, such as lunar or asteroid regolith or material gathered from the Martian moons. This would make VASIMR a usable propulsion system for early missions to establish bases on the moon, asteroids or Mars and should result in greatly superior mission mass ratios.

A return mission to the moon could begin by launching a single VASIMR tug, which would function as a lunar transfer vehicle and a reusable SSTO lander. Upon return from the lunar surface, the lander would carry enough lunar regolith to propel the transfer vehicle back to LEO, where it could rendezvous with the ISS and for a return trip to the moon. The lander would remain in lunar orbit awaiting the next mission.

Subsequent missions would only need to lift the crew, food and water, propellant for the lander, spare parts and any lunar surface payload. This would improve the economics of maintaining a lunar base.

One question would appear to be: Can VASIMR function using refractory oxides as a propellant?

RGClark · 2017-04-04 06:16:16

Antius wrote:

From what I have read on VASIMR, the thruster uses microwaves to produce cold plasma, which is then accelerated magnetohydrodynamically.
This concept could work extremely well if the propellant were wastes or unprocessed natural materials, such as lunar or asteroid regolith or material gathered from the Martian moons. This would make VASIMR a usable propulsion system for early missions to establish bases on the moon, asteroids or Mars and should result in greatly superior mission mass ratios.
A return mission to the moon could begin by launching a single VASIMR tug, which would function as a lunar transfer vehicle and a reusable SSTO lander. Upon return from the lunar surface, the lander would carry enough lunar regolith to propel the transfer vehicle back to LEO, where it could rendezvous with the ISS and for a return trip to the moon. The lander would remain in lunar orbit awaiting the next mission.
Subsequent missions would only need to lift the crew, food and water, propellant for the lander, spare parts and any lunar surface payload. This would improve the economics of maintaining a lunar base.
One question would appear to be: Can VASIMR function using refractory oxides as a propellant?

I like the idea. It could open up the entire Solar System to human exploration. I believe any material can be used as the propellant, though some materials are more efficient than others.

Bob Clark

Last edited by RGClark (2017-04-04 06:19:43)

Antius · 2017-04-05 09:39:47

I carried out some calcs assuming the use of a 10te dry mass, 100te full mass reusable lander and a 100te dry mass transfer vehicle, with Ve = 50km/s. The lander is equipped with a 3,285 kN thrust raptor engine, with a vacuum exhaust velocity of 3541.4m/s. Delta-V between lunar surface and LLO is taken to be 1870m/s. An approximate 130m/s fuel margin takes this up to 2000m/s.

The mass ratio for the lander is 1.759. So a 100te lander would need to consume 43.15te of methane/O2 propellant to make the landing. Of this some 9 tonnes will be methane and 34.16 tonnes are oxygen (Musk suggests a mixture ratio of 3.8 to 1). For a reusable SSTO, the oxygen can be manufactured on the moon from indigenous resources, for all but the first mission. That is important, as it effectively doubles the amount of payload that an individual lander can deliver to the surface.

The lander mass balance prior to landing is: Dry mass & crew: 10te; methane: 18te (9te for landing + 9te for return launch); 34.16te oxygen; 37.85te payload. Total = 100te. For return to LLO, the lander requires 9te of methane but 68.32te of lunar oxygen. Twice as much lunar oxygen is needed, because the SSTO must carry enough for the subsequent landing. The mass balance on takeoff is: Dry mass & crew: 10te; methane: 9te; 68.32te oxygen; 12.68te payload. In this case, the payload is lunar material.

The following mission, delivers another 37.85te of payload to the SSTO in LLO and another 18te of methane. Note that the 12.68te of lunar material is not enough to provide propellant to the transfer vehicle VASIMR engine, which must provide a delta V of 8km/s for itself, the crew and payload. However, if that 12.68te of material is lunar oxygen, it could be used to support high thrust chemical burns at optmimum points along the vehicle trajectory, taking advantage of the Oberth effect and dramatically reducing delta-V of the transfer vehicle between LEO and LLO.

The use of a reusable lunar SSTO supported by lunar oxygen production and a reusable VASIMR transfer vehicle, means that for repeat missions, some 66% of material launched into LEO can be delivered as payload to the lunar surface. Let’s say that a lander and transfer vehicle have a lifetime of 100 missions with minor refurbishment and each weigh 100te fully loaded. Over a 100 mission lifetime, some 3785te of payload would be delivered to the lunar surface for a total LEO payload investment of 5935te. That is a 64% payload efficiency.

For lunar missions, it doesn’t make sense using lunar material as VASIMR reaction mass under this scenario. For asteroids, with much lower surface gravity, it may make more sense.

Last edited by Antius (2017-04-05 09:41:47)

SpaceNut · 2017-04-05 17:29:45

The trouble with speed is how do you slow down quickly as the penalty for fuel is we are not getting there......

GW Johnson · 2017-04-06 16:54:58

A fast search on the internet for VASIMR turned up some performance figures for the 200 KW version that was to have been tested at ISS. That is various suffixes on VF-200 as a model number. Its optimum specific impulse power level is 200 KW, which is the rating, you should not exceed it. All of that is based on the preferred propellant: argon gas.

It produces 4980-5010 sec Isp at that 200 KW power, and 5.8 Newtons of thrust. The actual hardware core of the engine is about a meter long and half a meter wide, but there is a lot more gear needed to make it work.

Extra gear, such as a thermal radiator to dump waste heat, and a source of cryogenic coolant for the superconducting magnets (and no, I could not find out how much). Not to mention 200 KW of electricity.

The waste heat radiator appeared to be two folding panels maybe 2 m x 2 m each in size.

Excluding the electricity source, my uneducated guess is a single-digit number of tons for a 5 N thruster. This thing is only useful for delta-vee when you are already in vacuum and weightless, like all ion thrusters. I could not find any traceable weights for any of this gear.

GW

Last edited by GW Johnson (2017-04-06 16:56:42)

Oldfart1939 · 2017-04-06 18:09:16

I've been quiet here, but everyone should read Robert Zubrin's comments, entitled the VASIMR Hoax, in his small volume "Mars Direct." Currently available in Kindle format at Amazon.com.

"VASIMR...is neither revolutionary nor particularly promising. Rather, it is just another addition to the family of electric thrusters, which convert electric power to jet thrust, but markedly inferior to those we already have."
Robert Zubrin, in Mars Direct.

Antius · 2017-04-07 09:13:20

Setting aside VASIMR for the moment, I found some useful information for the NSTAR engine used on the Deep Space 1 (DS-1) comet and asteroid rendezvous probe.

http://www.astronautix.com/n/nstar.html

For a thrust of 0.092N and ISP 3100s, propulsion system mass minus fuel was 49.4kg, some 40% of which was the propellant storage tank (minus propellant). That gives a weight-thrust ratio of 537kg/N.

If mass scales with thrust, a 20N thruster would mass 10.74 tonnes. The NSTAR thruster consumed 2.3kW. Scaling the power requirements gives a requirement of 500kW for a 20N thruster. Let us assume that solar panels can be manufactured with 20% efficiency, mass 1kg/m2 and allow a 30% power loss due to alignment errors. Specific power would be 200W/kg. Total power supply mass for a 500kW supply would be 2.5tonnes. That takes total propulsion system mass to 13.24 tonnes. The delta-V for a round trip between LEO and LLO is about 16km/s using low-thrust propulsion. Assuming constant vehicle mass, at ISP 3100 the mass ratio would be 1.69. Some 41% of the starting mass of the vehicle must be propellant. Let’s say non-propulsion structure takes the dry mass to 20 tonnes and some 30 tonnes of payload are added. Vehicle starting mass with propellant would be 84.6tonnes. On the outbound journey, the vehicle would expend some 19.6 tonnes of propellant, so average mass over the outbound journey is 75.82 tonnes.

With a thrust of 20N, it would take 351days to make the outbound trip and 266 days to make the inbound trip. Round trip time is 20.3 months.

If we halve the payload mass, it makes little difference. I have simplified the calculation somewhat in that I have assumed a constant dry weight, including payload. In reality, the vehicle will make the return journey twice as quickly, giving a round trip time of perhaps 16 months. But the conclusion remains clear. Electric propulsion may be suitable for non-living cargo delivered to the lunar surface. But it is unsuitable for human passengers.

For repeat missions, excluding spares for the cycler and oxygen plant on the moon, each 30 tonnes of payload delivered to the lunar surface would require a total of 80.3 tonnes delivered to LEO, which includes 34.6 tonnes of cycler propellant and 15.7tonnes of methane for the lander. A payload efficiency of 37%, perhaps 33% accounting for the mass of the cycler and lander over 20 or so missions.

Amortising over 20 missions, with a lander and cycler cost of perhaps $100million and cost of payload delivery to LEO of $100/kg, the cost of payload delivery to the moon would be ~$500/kg, minus the purchase cost of the payload itself.

Last edited by Antius (2017-04-07 09:24:35)

GW Johnson · 2017-04-07 11:34:12

As I said above, I don't know much about electric thrusters.

From what I read about Ad Astra's VASIMR, it has the advantage of not having eroding electrodes, which could portend longer service life. It has the disadvantage of requiring large amounts of electricity for tiny whispers of thrust. That last is in common with the other electric designs, I think. Although it seems to offer more thrust than a Hall effect thruster at the same power level. At least, that's what I read.

As near as I can tell, Zubrin's real complaint about VASIMR is that fast trips to Mars with it are a hoax until there are nuclear electric power supplies both more powerful and lighter-weight than those contemplated today. Also from what I read, Chang-Diaz at Ad Astra says essentially the same thing in different words: fast ion flight is impossible until the high power/weight power supplies exist. That being the case, I do not understand why there is such acrimony between the two of them. Egos perhaps? Not an unknown phenomenon.

Not enough money is being spent on electric to make much of any difference to sending (or not) men to Mars or the moon. The real money pits are (1) SLS/Orion and (2) ISS. Of those two, I see ISS actually serving a useful purpose. SLS/Orion, maybe not so much, except as pork for certain congressional districts.

According to AIAA's "Daily Launch" email newsletter, word on the street is that USAF is getting seriously interested in the lower prices that reusability offers, specifically Spacex now, and the possibility with Blue Origin's New Glenn. They already demonstrated this by letting Spacex break into the military satellite business not long ago. They are speaking publicly now about the attraction of lower prices with reusability, when they didn't before.

The same newsletter says that ULA is looking at reusing the BE-4 engines in their new Vulcan launcher design that is supposed to replace Atlas-V. It also hints that the LOX-methane BE-4 from Blue Origin is ahead of, and favored over, Aerojet Rocketdyne's AR-1 LOX-kerosene engine for the same application. I have to wonder if methane offers fewer coking problems than kerosene. Kerosene coking ranks right up there with turbopump life among the limiting factors for engine reusability.

According to that same newsletter, Orbital ATK has decided not to pursue reusability in their next launch offering, whatever that turns out to be. They are still trying to get a launcher going for their Cygnus, and the imminent demise of Atlas-V over Russian engines makes that imperative.

GW

Last edited by GW Johnson (2017-04-07 11:40:01)

Oldfart1939 · 2017-04-07 11:37:21

For these electric thrusters to become at all useful, the thrust needs to increase by several orders of magnitude, while the mass of vehicle decreases by a similar scale. 20 Newtons isn't much thrust. The mass is huge, and even though the thrust is ongoing, it will have an abysmal acceleration.

Oldfart1939 · 2017-04-07 11:55:15

GW-

In my opinion, RP-1 is at best, a compromise. It offers ready availability and a low cost. Unfortunately it's a petroleum distillate, that undoubtedly contains traces of sulfur-bearing compounds which are ultimately responsible for coking. Methane, at least on paper, would be a much better choice as a fuel until the secondary issues are considered. These being, it's (1) a modest cryogenic fuel, and (2) requires larger tankage in the vehicle due to lower density. Rocketdyne attempted to address some of these problems a few years back by experimenting with 1,2-Diethylcyclohexane. They apparently abandoned it--I suspect due to the considerably higher cost relative to RP-1. It would be better in many regards to RP-1, since it's a synthetic by necessity; the Enthalpic density is comparable to RP-1 but without some of the refined fuel's drawbacks, such as gelling at low temperatures, and undoubtedly less coking.

GW Johnson · 2017-04-07 12:39:20

Well, methane is a mild cryogen, and it has to store under some amount of pressure too. So you have pressure tanks and cold, when you can use a much lighter, cheaper uninsulated, unpressurized tank for kerosene. Methane's Isp advantage is not all that high: about 11 sec difference out of ~300-310 for a sea level engine design.

I really don't understand its popularity, except for (1) you can make it on Mars, and (2) it won't coke up the regenerative cooling passages in your engine (leading to longer service life if you can solve the turbopump life problem).

I do smell more of a fad than a real trend here. Unless the non-coking feature is worth all the other hassle.

GW

Oldfart1939 · 2017-04-07 12:55:03

Of all the available fuels possible in use, my choice would be either Aerozine-50 or MMH. High Isp, and without the need for pressure or cryogenic storage. The real problem is the cost; not so much of it's manufacture, but through the implementation of EPA regulations. It's a transportation issue--not a manufacturing one. The combo of MMH and LOX is pretty good.

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#76 2015-05-27 09:36:40

Re: VASIMR - Solar Powered?

#77 2015-05-28 06:22:53

Re: VASIMR - Solar Powered?

#78 2015-05-28 18:58:34

Re: VASIMR - Solar Powered?

#79 2015-05-29 12:11:13

Re: VASIMR - Solar Powered?

#80 2015-06-03 16:06:13

Re: VASIMR - Solar Powered?

#81 2015-06-05 12:18:39

Re: VASIMR - Solar Powered?

#82 2015-06-05 14:37:17

Re: VASIMR - Solar Powered?

#83 2015-08-17 09:05:45

Re: VASIMR - Solar Powered?

#84 2015-08-17 11:17:33

Re: VASIMR - Solar Powered?

#85 2015-08-17 11:54:18

Re: VASIMR - Solar Powered?

#86 2015-08-24 11:03:20

Re: VASIMR - Solar Powered?

#87 2015-08-28 12:20:01

Re: VASIMR - Solar Powered?

#88 2015-08-28 13:12:47

Re: VASIMR - Solar Powered?

#89 2017-04-04 04:14:46

Re: VASIMR - Solar Powered?

#90 2017-04-04 06:16:16

Re: VASIMR - Solar Powered?

#91 2017-04-05 09:39:47

Re: VASIMR - Solar Powered?

#92 2017-04-05 17:29:45

Re: VASIMR - Solar Powered?

#93 2017-04-06 16:54:58

Re: VASIMR - Solar Powered?

#94 2017-04-06 18:09:16

Re: VASIMR - Solar Powered?

#95 2017-04-07 09:13:20

Re: VASIMR - Solar Powered?

#96 2017-04-07 11:34:12

Re: VASIMR - Solar Powered?

#97 2017-04-07 11:37:21

Re: VASIMR - Solar Powered?

#98 2017-04-07 11:55:15

Re: VASIMR - Solar Powered?

#99 2017-04-07 12:39:20

Re: VASIMR - Solar Powered?

#100 2017-04-07 12:55:03

Re: VASIMR - Solar Powered?

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