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posted by martyb on Thursday April 16 2015, @02:42PM   Printer-friendly
from the no-sign-of-dilithium-crystals dept.

The New Zealand based commercial space company Rocket Lab has unveiled their new rocket engine which the media is describing as battery-powered. It still uses fuel, of course, but has an entirely new propulsion cycle which uses electric motors to drive its turbopumps.

To add to the interest over the design, it uses 3D printing for all its primary components. First launch is expected this year, with commercial operations commencing in 2016.

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  • (Score: 2, Insightful) by ikanreed on Thursday April 16 2015, @03:16PM

    by ikanreed (3164) on Thursday April 16 2015, @03:16PM (#171616)

    Electric thrusters are essentially useless for launching from the ground to space because their maximum acceleration tends to be a lot less than 1 g.

    But once you're even in very low orbit, the thrust to (fuel) mass ratio of a nuclear power supply is gigantic. You can burn almost indefinitely, and get between space locations much faster.

  • (Score: 5, Informative) by kaszz on Thursday April 16 2015, @03:20PM

    by kaszz (4211) on Thursday April 16 2015, @03:20PM (#171618) Journal

    The important bits:
      * Small high-performance electric motors and lithium polymer batteries to drive the turbo pumps.
      * Height: 18 m, about a 1/3 the size of average rockets it will compete with to take satellites into space.
      * The new electric propulsion cycle could generate 20489 N of thrust.
      * Range of 500 km above the earth.
      * Uses carbon-composite.
      * Launch price for satellites at about $6.6 million, competition at $100 million. (NZ dollar or US dollar?)
      * Uses 3D printed parts for all primary components including its engine chamber, injector, pumps and main propellant valves.
      * Materials such as titanium and other alloys go through printers to create complex, lightweight structures, reducing the build time from months to days and increasing affordability.
      * Can carry 110 kg
      * 3 stages
      * Weight: 10 500 kg
      * Top speed: 7639 m/s
      * Rutherford engine that uses liquid oxygen-kerosene.
      * Stage-1 peak thrust 150 kN
      * Stage-2 peak thrust 18 kN

    It got to have some serious Li-Poly batteries! And the cost reduction seems to be about using metal 3D printing and carbon-composites? Less weight, faster production and smarter parts due 3D design = less cost?

    • (Score: 2, Interesting) by slinches on Thursday April 16 2015, @05:06PM

      by slinches (5049) on Thursday April 16 2015, @05:06PM (#171648)

      For one-off or small production lots that are common in space applications, metal 3D printing (DMLS) can be considerably cheaper and faster. It eliminates the need for expensive casting tools and hand fabrication that is expensive and difficult to reproduce.

      Although, I don't know how it comes out to be more efficient to carry fixed mass, relatively low energy density batteries on a rocket. I would have figured the added weight of the batteries and electric motor would be significantly greater than a small combustor and turbine + the fuel to drive it. In space applications you're usually trying to do everything you can to minimize the dry weight of the platform.

    • (Score: 3, Interesting) by subs on Thursday April 16 2015, @09:29PM

      by subs (4485) on Thursday April 16 2015, @09:29PM (#171739)

      I tried to run some ballpark numbers for the electrical component's weight and it didn't look very pretty. I have turbopump power requirement numbers for the RD-170, 192MW for an engine providing ~7887kN of thrust using a very efficient staged combustion cycle (309s SL Isp on RP-1/LOX!). Scaling it down to the 150kN of this rocket, an equivalent turbopump (assuming the chamber pressure stays about the same), it would need an electric motor of about 3.65MW output - comparable to a large electric locomotive. But that aside, if the turbopump were to run for ~2 minutes, the required energy storage to power this thing is on the order of 120kWh, which given current Li-Ion storage comes to about 1 ton in mass. That's about 10% of the vehicle's lift-off mass, aside from the weight of the supporting electrical conversion equipment and the motors.
      My guess is, they run a lot lower chamber pressures and a so can get away with a lot weaker pumps, trading efficiency and mass-to-orbit in the process. Maybe that's the reason why even their 3-stage design has a pitiful 1% of launch mass to orbit, vs. e.g. the Falcon 9's 2.6% at only 2 stages (meaning, despite fewer stages, it's more than 2x as capable).

      • (Score: 2, Informative) by rumata on Friday April 17 2015, @03:01AM

        by rumata (2034) on Friday April 17 2015, @03:01AM (#171859)

        Your guesstimate doesn't start from a very useful point. The RD-170 is pretty much _the_ highest performance, highest pressure lox/kero engine there is (maybe except the Merlins). If you look at the picture on this page [] you can make a reasonable guesstimate of their engine parameters.
        I came up with 4-5MPa chamber pressure (compare to 24.5 for the RD-170), based on exit-diameter estimate of 200mm, expansion ratio guesstimate 10, and thrust 16.7kN.

        Additionally they use carbon fiber tanks in a small diameter vehicle. I'd take a guess that due to minimum gauge reasons they have excess strength in the tanks. One way to utilise that would be to run the tanks at higher pressure than traditional, further reducing the pressure the pump has to supply.

        This probably allowed them to minimise the pump power requirements to the point where going electrical (_much_ simpler than turbines) became attractive.


        • (Score: 3, Informative) by subs on Friday April 17 2015, @12:17PM

          by subs (4485) on Friday April 17 2015, @12:17PM (#171982)

          Point taken, but without more detailed engine data we won't know for sure. At present, their website is rather light on that, my I guess they're still working out basic stuff.
          And yeah, upon closer examination, your 4-5MPa seems more realistic, though I came up with a somewhat different nozzle area to match their figures. RP-1/LOX, ~270s at SL, ~16.6kN (implying ~6.2kg/s flow rate), chamber at ~4.5MPa I calculated around 25 sqcm throat and a 6x expansion ratio to the nozzle exit (nozzle exit diameter ~14cm). Anyhow, adjusting for this lower pressure the work required appears closer to 1MW on the pump (allowing for some electrical losses along the way), or around 30kWh for a 2 minute burn - a battery pack weighing in at ~200-300kg and that's without the compressors, motors and electrical equipment. Not a deal breaker in a 10 ton rocket, but certainly not light. By comparison a complete ~2MW turbopump can weigh less than 1/4 of that and you could feed all 9 engines with it, significantly reducing the number of components. I'd have hard time imagining this scaling up, though.
          In any case, maybe performance here was second to cost. Pure electrical systems can be quite cheap. Anyhow, thanks for setting me straight on the chamber pressure.

    • (Score: 2) by subs on Thursday April 16 2015, @09:38PM

      by subs (4485) on Thursday April 16 2015, @09:38PM (#171741)

      In fact, given the shitty performance and still requiring the complexity of liquid-fuel engines and 3 stages, why didn't they just go with a dumb-as-a-duck solid fuel design, like the Japanese Epsilon []. About similar payload-to-orbit performance and much, much simpler design. In fact, per kg to orbit, the Epsilon is about 2x as cheap as this crap heap.

      • (Score: 2) by kaszz on Thursday April 16 2015, @11:22PM

        by kaszz (4211) on Thursday April 16 2015, @11:22PM (#171778) Journal

        Perhaps this is just a step before they do something additional that will gain some very useful advantage?

      • (Score: 1) by rumata on Friday April 17 2015, @02:28AM

        by rumata (2034) on Friday April 17 2015, @02:28AM (#171851)

        In fact, per kg to orbit, the Epsilon is about 2x as cheap

        How do you figure that?

        Epsilpon (from your wikipedia link): 450kg to 500km SSO for $38M.
        --> $84.4k/kg

        Elektron : 100kg to 500km SSO for $4.9M.
        --> $49.0k/kg

        Course, Epsilon is flying, while Elektron is not.


        • (Score: 2) by subs on Friday April 17 2015, @10:38AM

          by subs (4485) on Friday April 17 2015, @10:38AM (#171969)

          Thanks for correcting me. I didn't notice that the 100kg they quoted was for a 500km SSO. I was going by the original article's infographic, which just said "110kg to 500 km orbit" and $6.6M. Assuming both were to a regular (easterly) 500km orbit, that would come to $54k/kg for the Epsilon and $60k/kg for the Elektron. Their website paints a rather different picture, as you noted. Why, however, does their price estimate differ by so much? My guess is, they don't know how much it'll cost either.
          Having had a look at the very few performance numbers they mention a vacuum Isp of 327s, which would imply a fairly high chamber pressure in excess of 10MPa (in fact, the NK-33 with its 331s runs at clost to 15MPa) and so a *very* large turbopump (on the order of 1MW). I'm very skeptical of their performance claims, since just the batteries needed to support this would be massive, but we'll see when and if this thing actually flies.

  • (Score: 1) by cyberthanasis on Thursday April 16 2015, @07:30PM

    by cyberthanasis (5212) on Thursday April 16 2015, @07:30PM (#171698)

    The man in the promotion video talks about launching to orbit as something easy and the difficult part is the economics. I am very sceptical about this, but I hope that they succeed.