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#1260
by
QuantumG
on 31 Jan, 2018 22:32
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Nice... Looking at the 2018-01-21 launch I get:
T+02:33 MECO
T+02:39 Stage-2 Ignition
T+04:50 3 km/s callout
T+08:15 SECO
So a 153 second burn duration for the first stage, and 336 second burn duration for the second stage.
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#1261
by
FutureSpaceTourist
on 31 Jan, 2018 23:10
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#1262
by
vaporcobra
on 31 Jan, 2018 23:51
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and they responded to me!
They've produced 40 flight-ready Rutherfords up to this point, and intend to produce another 100 in 2018. That at least caps the number of flights they can conduct in 2018 at around 12, basically a monthly cadence.
https://twitter.com/RocketLab/status/958853552784330752
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#1263
by
ArbitraryConstant
on 01 Feb, 2018 00:38
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It's interesting to compare the upper stage Rutherford engine to Kestrel and AJ-10, since those are a similar thrust class, and I think the battery pumped rocket cycle would be most often traded against pressure fed. The higher ISP is exciting.
I think Rutherford must be getting a significantly higher chamber pressure than pressure fed engines, due to not having the constraint of tank pressure > chamber pressure and batteries having greater specific energy than COPV helium.
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#1264
by
vaporcobra
on 01 Feb, 2018 00:58
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FWIW, at 35kg and 24,000N of thrust, the TWR of Rutherford is about 69.9. Battery powered turbopumps take a big toll. The impressive ISPs are clearly a necessity for Electron's efficiency to be practical!
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#1265
by
ArbitraryConstant
on 01 Feb, 2018 02:06
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FWIW, at 35kg and 24,000N of thrust, the TWR of Rutherford is about 69.9. Battery powered turbopumps take a big toll. The impressive ISPs are clearly a necessity for Electron's efficiency to be practical!
Well, pressure fed and electric pumped both have mass penalties. Electric has pump motors and battery mass, pressure fed has gas storage and a heavier tank. Electric can mitigate by battery jettison, pressure fed can mitigate with autogenous pressurization. Solid is also relevant at this size class and its tradeoff is relatively low ISP which it usually mitigates with an extra stage.
What I infer from looking at engines in roughly the same thrust class is that pressure fed ends up with a much lower chamber pressure, hence worse ISP especially in atmosphere with the first stage. Electric could easily use a lower chamber pressure and save battery mass at the cost of extra propellant but it seems like they ended up going to a higher pressure, and these numbers have actually been revised upwards from previously so this seems like a gift that keeps on giving.
We can compare other engines like FRE-1, Firefly's upper stage engine, that was going to be autogenous pressure fed with methane which has an intrinsically higher ISP than kerolox, and it was a paper engine, and they still ended up with lower ISP. That must have been because even with no pressurant constraint it still wasn't worth it for them to go to a higher pressure. So electric ends up looking really good here IMO.
All of these have penalties compared to advanced cycles but scaling those down is challenging. Russia has a number of advanced cycle engines in a similar thrust class for fourth stages like Briz-M and they have extremely good ISP for hypergolics but the T/W and dry mass of the engine isn't amazing. Similarly there's American kick stages at similar sizes, but these are usually solids. Still looks like a good deal for RL because they turn a bunch of complex and expensive problems into the relatively simple task of 3D printing engines and connecting them to batteries.
I don't expect to see EELV class launchers using electric pumped engines, but they look well suited for small launchers, and I wouldn't be surprised to see them in the future for things like advanced kick stages and landers.
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#1266
by
LtWigglesworth
on 01 Feb, 2018 02:07
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I think Rutherford must be getting a significantly higher chamber pressure than pressure fed engines, due to not having the constraint of tank pressure > chamber pressure and batteries having greater specific energy than COPV helium.
They have given us data on the thrust (5500 lb/2500kg) and isp (311s), which gives a mass flow of ~ 8kg/s.
At a 2.6 ox/fuel ratio, the density is about ~1050 kg/m
3, so volumetric flow through the pumps on each engine is about 0.0076m
3/s.
They have also said that the batteries on the first stage provide 1MW of power, or 111kW per engine.
So, assuming a pump efficiency of 80%, the pressure rise across the pumps can be estimated at about ~12 MPa.
Assuming an injector pressure drop of 20% of chamber pressure that would suggest a chamber pressure of ~10MPa.
That's my guesstimate from the pump side.
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#1267
by
ArbitraryConstant
on 01 Feb, 2018 04:18
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I think Rutherford must be getting a significantly higher chamber pressure than pressure fed engines, due to not having the constraint of tank pressure > chamber pressure and batteries having greater specific energy than COPV helium.
They have given us data on the thrust (5500 lb/2500kg) and isp (311s), which gives a mass flow of ~ 8kg/s.
At a 2.6 ox/fuel ratio, the density is about ~1050 kg/m3, so volumetric flow through the pumps on each engine is about 0.0076m3/s.
They have also said that the batteries on the first stage provide 1MW of power, or 111kW per engine.
So, assuming a pump efficiency of 80%, the pressure rise across the pumps can be estimated at about ~12 MPa.
Assuming an injector pressure drop of 20% of chamber pressure that would suggest a chamber pressure of ~10MPa.
That's my guesstimate from the pump side.
Nice, cheers. I think I guessed about 5.5 MPa from the area ratio but that was a while ago and their ISP and thrust numbers have increased from previous statements so it sounds like they've gone through a few iterations.
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#1268
by
Lar
on 01 Feb, 2018 04:31
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#1269
by
Zed_Noir
on 01 Feb, 2018 09:13
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If they use the Rutherford engine in some sort of multi burn orbital adjustment stage. Would it make sense to add a solar array along with a small battery to power the turbopump and keep the propellants from boiling off or freezing?
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#1270
by
edzieba
on 01 Feb, 2018 13:33
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If they use the Rutherford engine in some sort of multi burn orbital adjustment stage. Would it make sense to add a solar array along with a small battery to power the turbopump and keep the propellants from boiling off or freezing?
At 111kW/engine, that's a panel array similar in size to the ISS' (up to 120kW). If you add an accumulator to allow you to use a smaller array, you've now added the battery you were trying to eliminate back in again.
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#1271
by
Zed_Noir
on 01 Feb, 2018 17:23
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If they use the Rutherford engine in some sort of multi burn orbital adjustment stage. Would it make sense to add a solar array along with a small battery to power the turbopump and keep the propellants from boiling off or freezing?
At 111kW/engine, that's a panel array similar in size to the ISS' (up to 120kW). If you add an accumulator to allow you to use a smaller array, you've now added the battery you were trying to eliminate back in again.
What is that kwh battery storage for the Electron core and how many kg does it take?
Only need one Rutherford with about 90 seconds bursts of burn time at the most during orbital adjustments at perigee. So need about 120 seconds of battery storage time to operate the turbopump and ancillary functions. Multiple orbits can be use to charge up the battery to full.
So the question is how much battery storage is needed for 120 seconds of turbopump operation.
This is a multi burn orbital adjustment stage not an upper or departure stage.
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#1272
by
Kansan52
on 01 Feb, 2018 17:43
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How much or do they power the system through umblicals until launch and umbilical disconnect? In other words, battery don't begin to deplete until no ground power is available. Certainly would helped that initial current surge that normally happens with motor starts.
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#1273
by
ArbitraryConstant
on 01 Feb, 2018 18:32
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What is that kwh battery storage for the Electron core and how many kg does it take?
If it's a megawatt for 2.5 minutes then about 42 kwh. But it is almost certainly more. The "over a megawatt" number was reported years back and reported thrust/ISP numbers have increased since then, a greater flow rate and higher pressure implies higher power.
For the battery mass, who knows. It depends on how much cooling is necessary and what their cooling system looks like. The first stage could potentially use propellant flow for cooling but probably not the upper stage given the battery jettison.
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#1274
by
TrevorMonty
on 01 Feb, 2018 21:07
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The better batteries are 240wh/kg so 175kg for 42kwh. But these are for normal discharge rates of 2-3C. Electron would need 20-30C, so expect lot heavier battery with lower wh/kg.
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#1275
by
sanman
on 03 Feb, 2018 10:26
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What is the future evolutionary roadmap for this rocket?
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#1276
by
MATTBLAK
on 03 Feb, 2018 10:44
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Some minor upgrades, no doubt. If their business should really catch on, then we might expect a scaled-up version capable of placing at least a couple tons in various orbits.
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#1277
by
ArbitraryConstant
on 03 Feb, 2018 16:44
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What is the future evolutionary roadmap for this rocket?
Well, we've seen increased thrust and ISP from the Rutherford engine. This suggests they may already have improved on the advertised performance. I assume they will make other incremental performance improvements over time as any rocket would. One nice aspect of using batteries is that RL can incorporate improved batteries over time without much development work on their part. RL's batteries must be quite a high power chemistry since they are discharged in only a few minutes, and improvements in density vs power are always happening.
Another improvement would be qualifying the upper stage for multiple burns. They clearly did not regard this as part of their minimum viable product but it would enable them to eliminate the kick stage for circularization burns and may enable some neat BEO applications.
For performance improvements, it wouldn't be that surprising to see incremental performance improvements up to perhaps several times the current payload to LEO, as many of the technologies Rocket Lab draws on are improving quite rapidly, such as carbon composites, batteries, 3D printing, etc.
The trend (Orbital ATK, SpaceX) has been for smallsat launchers to end up at larger sizes but I tend to think Rocket Lab experiences the pressure to do so less strongly, as SpaceX's presence as well as others like Blue Origin reduces the potential benefit. RL's technology likely has trouble scaling up that far, turbopumped rocket cycles are still going to win above a certain size.
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#1278
by
speedevil
on 03 Feb, 2018 17:02
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The trend (Orbital ATK, SpaceX) has been for smallsat launchers to end up at larger sizes but I tend to think Rocket Lab experiences the pressure to do so less strongly, as SpaceX's presence as well as others like Blue Origin reduces the potential benefit. RL's technology likely has trouble scaling up that far, turbopumped rocket cycles are still going to win above a certain size.
My knowledge of turbomachinery is small, but I wonder if a significant part of the hardness of it on the low end is getting startup to work well.
If this is true, a small (relatively) electric motor of a few tens of percent of nominal might help lots in getting a marginal due to low mass design stable.
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#1279
by
ArbitraryConstant
on 03 Feb, 2018 17:56
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My knowledge of turbomachinery is small, but I wonder if a significant part of the hardness of it on the low end is getting startup to work well.
The startup sequence is challenging at any size.
My understanding is that as you scale turbines down the mechanical tolerances for good efficiency are extremely challenging. The result is that the turbine can't scale down as much as the rest of the engine and the thrust to weight ratio ends up being quite poor. Hence an electric pumped engine has advantages because while there's a mass penalty, it's at least a cheap mass penalty with low development costs. Also as we see with Electron's upper stage, it's possible to jettison some of the battery mass.
