Plasmoid Thruster Space Propulsion Designs

Plasmoid thrusters have had successful proof of principle experiments done in the lab at Princeton University. The systems should be compact and simple and based upon magnetism present in solar flares. There are designs for initial systems with ISPs in the range of ion drives. The systems should scale to higher ISPs of 50,000 and to higher thrust power levels. They hope to make the first systems for space deployment in the 2-5 year time frame. The first applications would be tugs or stages moving from low earth orbit to the moon.

It can deliver high thrust at high and variable exhaust velocity (tens to hundreds of km/s). It should therefore get unsurpassed gas mileage for longer trips (Mars and beyond).

According to computer simulations run by the Princeton Plasma Physics Laboratory and the National Energy Research Scientific Computing Center at the Lawrence Berkeley National Laboratory in Berkeley, California, her thruster concept produced exhaust velocities that are ten times greater than a traditional ion propulsion system with speeds at hundreds of kilometers per second. Exhaust velocities in the range of 20 to 500 km per second, controllable by the coil currents, are observed in the simulations.

Plasmoids with radius 10 cm and reconnecting field of 800 Gauss , the calculated thrust is about 50 Newtons. Taking into account a duty cycle of about 33 % (i.e. the distance between two consecutive plasmoids is twice the plasmoid length). The input power varies from a few to a few hundred kiloamps. In the simulations 100 kiloAamps corresponds to about 10 MW of power. For this unoptimized high-power case (with a thrust of 50–100 N), the ratio of thrust over power is thus about 5 to 10 milli-Newtons per kilowatt. Tentatively the optimal parameter range for this new thruster will be ISP (specific impulse) from 2000 to 50 000 s, power from 0.1 to 10 MW and thrust from 1 to 100 N. It would thus occupy a complementary part of parameter space with little overlap with existing thrusters.

There are various small funded projects working on Plasmoid Thrusters.

There is a US Air Force SBIT phase 1. Eagle Harbor Technology, Inc. (EHT) and the Professor Ben Jorns at the Plasmadynamics & Electric Propulsion Laboratory at the University of Michigan (UM) partnered to advance the state of the art Rotating Magnetic Field-Field Reversed Configuration thruster (RMF-FRC). This technology takes inspiration from fusion energy science. The RMF created plasmoids are accelerated out of the thruster at high speed. Using previously developed EHT Full-Bridge Modules, EHT built a solid-state power system to drive the RMF coils for the UM thruster. EHT has used a similar power systems to drive an Electrode-less Plasma Source and High Power Helicon Experiment for pulse durations up to 1 ms. In this SBIR, EHT developed a new version that can operate continuously at up to 4 kW average power while still driving peak currents of 2 kA at 500 kHz.

Rotating Magnetic Field-Field Reversed Configuration (RMF-FRC) operation. Left: Plasma fluxes into volume with biased magnetic field. Middle: RMF antenna generates azimuthal current in plasma giving rise to field reversal. Right: Lorentz force interaction with background magnetic field.

There is a new multi-turn, multi-lead design for the first generation PT-1 (Plasmoid Thruster) that produces thrust by expelling plasmas with embedded magnetic fields (plasmoids) at high velocities. This thruster is completely electrodeless, capable of using in-situ resources, and offers efficiencies as high as 70 percent at a specific impulse, Isp, of up to 8,000 s. This unit consists of drive and bias coils wound around a ceramic form, and the capacitor bank and switches are an integral part of the assembly. Multiple thrusters may be ganged to inductively recapture unused energy to boost efficiency and to increase the repetition rate, which, in turn increases the average thrust of the system.

The thruster assembly can use storable propellants such as H2O, ammonia, and NO, among others. Any available propellant gases can be used to produce an Isp in the range of 2,000 to 8,000 s with a single-stage thruster. These capabilities will allow the transport of greater payloads to outer planets, especially in the case of an Isp greater than 6,000 s.

19 thoughts on “Plasmoid Thruster Space Propulsion Designs”

  1. I read the article 3 times, and didn’t find where Brian explain whats “Plasmoid Thruster” and how it works.
    This should be the start for any such article.

  2. Um can someone simplify this for me? Put one of these on a one ton satellite and it could accelerate it at .x g’s for x hours on a gallon of which gas?

    • Less, of course, the annoying mass of the energy source(s). 10 MW of electricity is hinted to as being nuclear-provided. I’d expect that. The α of 10 kg per kW then backs into 100 tons for the reactor. That’s a lot of ‘dead weight’. Or not …

      Physics provides further guidance: 50 N given 10 MW at 50% overall efficiency (very reasonable assumption) is 5 MW of exhaust. In kinetic energy terms on a 1 second analysis window (to get rid of all the annoying energy/power conversions by proxy), the figure-of-merit is 50 N ÷ 5,000,000 J/s (W) = 10 micronewtons per watt.

      We also remember that F = mv and that E = ½mv². divide those two and ‘m’ goes away leaving F/E = 2/v. This also reduces to newtons per watt! 0.000010 Ns/W = 2/v thus v = 200,000 m/s.

      Working that back into the F = mv partial equation, and we get 50 ÷ 200,000 = Δm per second. 0.000250 kg/s. 21.6 kg/day. 7,880 kg/year. Hallelulah!

      Well, let’s see a bit further. The spacecraft need to buzz along for what 5 powered years? OK, it needs about 40 tons of exhaust mass. And a 100 ton reactor. And a whole bunch of waste-heat fluidics panels to wick away the reactor’s waste energy. The whole thing might come in at … hmm… 500 tons? How fast might it go at ¾ flight (3.7 years)?

      Tsiolkovsky’s Rocket Equation helps … ΔV = Ve log( Mi/Me ) where Ve is exhaust effective velocity (200,000 m/s) and Mi is initial mass (500 tons) and Me is end-mass (500 minus 30 is 470 tons). which calculates out at 12,400 m/s or 12.4 km/s.

      I dunno … it seems like a whole lot of energy-producing mass is weighing down my enthusiasm. 12.4 km/s is peanuts. Well, not peanuts: 2.6 AU/year … so Jupiter in only 2 years, (ahem… 3.7 years after starting).

      And the distance to the solar gravitational telescope point … 500 AU … assuming a more reasoned 65% thrust, 30% decelerate and 40% of rest mass as reaction-mass, gives (grind, grind, grind) 30 years of acceleration, 12 AU/year and about 30 years of drift and 15 years of deceleration for 75 total elapsed years … to get to 500 AU.

      Nope. Need much more power, and a MUCH light weight, potent nuclear source. And no-touch-no-fix reliable for decades. One heck of a proposition.

      ⋅-⋅-⋅ Just saying, ⋅-⋅-⋅
      ⋅-=≡ GoatGuy ✓ ≡=-⋅

  3. As for power, wouldn’t the Lawrenceville Plasma Physics’ Dense Plasma Focus fusion idea provide both electric power and thrust?

    IIRC, the exit speed of the helium products of the pB11 fusion is in the range of 10% of C.

    • LPP was originally a thruster. I doubt that LPP would be able to produce adequate net energy and meaningful thrust because they must convert the kinetic energy of the fuel and byproducts to electrical energy.

      You could probably use LPP in conjunction with some other ion/plasma thruster. As Brett rightly points out the problem isn’t making usable ion or plasma engines but that we can’t power them. Specifically you need a high power per kg of weight. LPP would work well in this regard as it is quite light. Solar works well but is limited to ranges near Mars.

      • LPP might be useful as a thruster; It would consume net power, but you’d get a lot more output out of it for KW in than you would for a simple plasma thruster, because the fusion taking place would amplify the available energy.

        The goal of LPP is PB fusion, which is aneutronic, but absurdly hard to reach engineering breakeven on. If they compromised on the aneutronic aspect, and used an easier reaction, they’d much more easily get net energy out, even while using it for propulsion, too.

        At the cost of their apparatus becoming radioactive, which is why they don’t just dump some deuterium into their system and, presto, net power out! They don’t have the budget to deal with that. But that might not be an issue for a rocket engine to be used for a relatively short time and then thrown away in space.

        But for rocketry purposes Helion Energy looks more promising to me.

  4. To make human settlement on Mars in a fast, simplified, scalable way the Musk’s way is great.

    Other ways would take too long, need so much other tech researched or extra structures built,…

    Things would speed up after enough people were already on Mars,…

  5. X3 ion thrusters look promising for Mars and some other missions.

    The kw range is higher up above 100 kw, if they cluster more of them together is could be used for manned missions to Mars. They would need a lot of solar panels. Solar is proven, used technology.

  6. I’m not saying this is a bad technology, just getting away from the grid is a big improvement. But the limitation on electric thrusters of all sorts isn’t the thruster. It’s the power supply.

    And this doesn’t solve that. Solve that problem, and almost any kind of electric thruster would get this sort of performance. Don’t solve that problem, and you’re stuck with low accelerations regardless of how you use the electricity to accelerate the propellant.

    • As far as I understand plasma based thrusters can be actually scaled for thrust since you can convert more mass to plasma by applying more energy, while in Hall effect thrusters there are more factors that constrain the scaling. But question stands, there is already an operating thruster which is 2 times more efficient, VASIMR. The limiting factor is indeed power to weight ratio of the power source. It is hard to believe that they would be able to achieve anything in 2-5 years when they do not have neither thruster nor the power source.

    • Hail Brett… See my reply above (to Jim Bærg … The power supply is quantitatively THE problem, not the thruster. Read and pour another strong beer.

  7. 100kW per 1 Newton of thrust is 2.5 worse than 40kW/N of VASIMR. NEXT has the power to thrust ratio of 24kW/N. What is the point?

      • The problem is power to weight ratio of the power source. Usually the higher output means increased weight. What is the point of being more efficient than VASIMR, but also have lower thrust and consequently slower and heavier at the same time?

  8. Played interesting game lately – Terra Invicta. Pretty neat, not like Kerbal space program, but anyway you can get some fun and grasp the mechanics of space travel.

    Would be pretty interesting, what would be the best.
    Elon wants to refuel Starship in Earth orbit and then go to Mars,.. But you need a lot of launches to have enough fuel for Mars and perhaps is not the best way.

    Perhaps they should build 2 space docks, platforms at the most appropriate places in Earth and Mars orbit. They would be used to refuel and move cargo with some emergency options as well. Earth- space dock, Mars – space dock launches would be used with normal Raptor thrusters, because methane on Earth and Mars won’t be a problem. Now the travel between 2 space docks could be made by more fuel efficient newer ion engines. That would cut down the need for so much refueling and so many launches for fuel, but they would need to transfer cargo 2 times and that is another problem,… I am really disappointed by Elon’s decision to postpone Mars launch to so much later date. I really can’t say if that is because political pressure or other reasons. Hope they build good version of tesla bot, because they could use it in first starship launches to Mars.

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