Rocket Lab Unveils "Electric" Rocket Engine

New submitter Adrian Harvey writes 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 rocket 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.

Comments

Re:Questionable engineering decisions.

[gman003]

Uh, the SSME engines ran off the same propellant as the rocket - LH2 and LOX. It's a staged-combustion rocket - some of the fuel and oxidizer flow was diverted to a preburner, which partially combusted them (the mixture was fuel-rich, limited by oxygen), ran the fuel-rich exhaust through turbines for the fuel and oxidizer pumps, then exhausted into the main combustion chamber where it was mixed with the remaining oxygen to complete combustion.

A better example for a separate propellant would have have been the V2 rocket, which burned ethanol and LOX, and had a pump powered by hydrogen peroxide.

Right on all other points, though.

Re:Meh.

[Anonymous Coward]

Because you're asking the wrong question. Aerospace engineers don't plot energy density vs. power density: they plot specific impulse (or equivalently exhaust velocity) vs. acceleration (or thrust). Which gives the same answers in the form of variables more directly useful for rocket equations.

You can get such plots in any good propulsion or mission planning text. E.g. Rocket Propulsion Elements by Sutton & Biblarz (8th ed). has one on p. 42; Space Mission Engineering: The New SMAD* by Wertz et al. has one on p. 548. (*Space Mission Analysis and Design was the name of earlier editions.)

Re:Hype pain

[brambus]

Turbine engines typically achieve around 33%-34% efficiency. Going off of Wikipedia, non-rechargeable lithium batteries are around 1.8MJ/kg, whereas kerosene is around 46MJ/kg. Now with kerosene, you need to carry around another 2.5 parts of oxygen, so gram-for-gram, the split is around 13 MJ/kg for RP-1/LOX. Accounting for engine efficiency, it comes to around 1.6MJ/kg for non-rechargeable lithium batteries driving an electric motor pump vs. 4.5MJ/kg for an RP-1/LOX turbopump. IOW, the turbopump version is around 3x more efficient. Now the dry weight of the assembly. A 1MW turbopump can be built in as little as 50kg (in fact, the Merlin 1C turbopump weights around 70kg and produces 1.86MW). A comparable DC electric motor would probably weigh in at close 2x than that. Not to mention, the dry weight of the turbopump is just the pump plus about 4-5% of the fuel weight for the tank to hold it, whereas for the electric motor pump + batteries, dry weight is essentially unchanged throughout the entire burn.
Overall for a 1MW pump system for a 120s burn, the numbers would stack up roughly like this:

• wet turbopump: 50kg + 8kg of fuel + 20kg of oxidizer + 2kg tank, total: 80kg.
• dry turbopump: 50kg + 2kg tank = 52kg
• wet & dry motor + batteries: 100kg motor with pump, 74kg batteries, total: 174kg.

From a pure performance perspective, electrically driven pumps in rocket engines are simply worse. However, considering the cost and complexity of turbopumps and the relatively small part that fuel pumping overhead contributes to overall efficiency, it may be a cost worth paying, especially on a smaller launch vehicle, where the electrical equipment is relatively cheap. I'm not convinced ti scales to multi-MN engines, though, as there the electrical requirements would be enormous (100MW+ electric motors are somewhat impractical, as is the supporting electrical equipment).