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 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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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).
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 http://www.rocketlabusa.com/about-us/vehicle-technologies/carbon-composite-technologies/ [rocketlabusa.com] 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.
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.