RocketLab Discussion Thread

[Silavite]

5 minutes ago, sevenperforce said:

I know he just means that CF is 4X lighter than SS in terms of weight per unit area, not that the mass ratio will be 4X better than Centaur, but still…wow.

I’m sure the upper stage engine will have to have decent throttling. You need that for precision orbital insertion, anyway. Unthrottleable or limited-throttle liquid engines are mostly the regime of disposable first-stage engines. And throttling isn’t too hard to do with a GG engine.

The additional density of methane as opposed to LH₂ should also help with the mass fraction.

On another note, after looking at Archimedes claimed sea level(?) I_sp, I decided to look at Rutherford's sea level I_sp. The Wikipedia page claims 311 seconds at sea level, but that number is equivalent to the sea level I_sp of the RD-180! As good as Rutherford is, I have a very hard time believing that it matches the RD-180 in that regard. Electron's user guide says that the specific impulse for the sea level Rutherford engine is 311 seconds, but it does not explicitly say that 311 seconds is the value for the sea level Rutherfords at sea level. It seems likely that the I_sp numbers for the Rutherford and Archimedes are both some kind of average between sea level and vacuum.

[sevenperforce]

18 hours ago, RyanRising said:

I suspect it's just the sea level engine running in a vacuum. The Merlin 1D (sea-level) should be roughly comparable, same cycle, modern tech, and that has ~310 s in a vacuum. That means between equivalent engines, you gain ~10, 15 seconds of specific impulse by going with methane over kerosene. Raptor is such a stupidly high-pressure engine, and is positioned so early on in the timeline of methalox launch vehicles, that it makes the performance benefit of methane over kerosene seem a lot higher than it actually is.

I did more math and I’m not so sure that makes sense. Or at least, there’s something fishy with the thrust numbers given on RocketLab’s site. The seven Archimedes engines together are supposed to have 5960 kN at liftoff and 7530 kN at burnout. The vacuum Archimedes produces 1110 kN.

But if that difference in liftoff thrust and burnout thrust is attributable to SL vs vacuum performance, and the vacuum isp of the SL engines is 320 seconds, then that would put the sea level isp at only 253 seconds, which is impossibly low. It would also suggest that the specific impulse of the vacuum Archimedes is a paltry 330 seconds, far lower than the Merlin 1D Vacuum. So that can’t be right, either.

[sevenperforce]

29 minutes ago, Silavite said:

On another note, after looking at Archimedes claimed sea level(?) I_sp, I decided to look at Rutherford's sea level I_sp. The Wikipedia page claims 311 seconds at sea level, but that number is equivalent to the sea level I_sp of the RD-180! As good as Rutherford is, I have a very hard time believing that it matches the RD-180 in that regard. Electron's user guide says that the specific impulse for the sea level Rutherford engine is 311 seconds, but it does not explicitly say that 311 seconds is the value for the sea level Rutherfords at sea level. It seems likely that the I_sp numbers for the Rutherford and Archimedes are both some kind of average between sea level and vacuum.

Curiouser and curiouser. The Rutherford has a big whopping advantage over the RD-180, of course, because 100% of its propellant goes directly to thrust, while the ORSC cycle of the RD-180 uses up a significant amount of chemical energy in the preburner. That is why it has a fairly low TWR: the electric motor and batteries are hella heavy. But the RD-180 has a chamber pressure more than double that of the Rutherford, too.

Generally, what is the relationship between chamber pressure and specific impulse for a given propellant? Is it quadratic? Exponential? 

[StrandedonEarth]

33 minutes ago, Silavite said:

As good as Rutherford is, I have a very hard time believing that it matches the RD-180 in that regard.

Doesn’t the Rutherford use electric turbopumps, IIRC? Would that help?

[sevenperforce]

6 minutes ago, StrandedonEarth said:

Doesn’t the Rutherford use electric turbopumps, IIRC? Would that help?

The RD-0124 produces 359 seconds of specific impulse thanks to its chamber pressure of 2280 psi. The Rutherford vacuum engine, at 1400 psi, can only manage 343 seconds of specific impulse. So even though electric turbopumps are more efficient from a chemical perspective, chamber pressure is going to drive specific impulse much more aggressively.

I will also note that the Merlin 1D vacuum engine is able to push 348 seconds of specific impulse despite having a chamber pressure only slightly higher than the Rutherford, and with a much less efficient power cycle.

[FleshJeb]

@sevenperforce In addition to your recent calcs here, I've been running your numbers from:

https://forum.kerbalspaceprogram.com/index.php?/topic/156515-rocketlab-discussion-thread/&do=findComment&comment=4061314

and coming out with a negative dry mass of the first stage.

Expendable first stage: 5630m/s = 320s * 9.81m/s^2 * ln (480 / (4.6 + 79.1 + x))

It doesn't even cross zero until ISP = 329s

This suggests a problem with the assumed ISPs and/or the dry mass fraction of the second stage.

[Silavite]

1 hour ago, sevenperforce said:

Curiouser and curiouser. The Rutherford has a big whopping advantage over the RD-180, of course, because 100% of its propellant goes directly to thrust, while the ORSC cycle of the RD-180 uses up a significant amount of chemical energy in the preburner. That is why it has a fairly low TWR: the electric motor and batteries are hella heavy. But the RD-180 has a chamber pressure more than double that of the Rutherford, too.

Generally, what is the relationship between chamber pressure and specific impulse for a given propellant? Is it quadratic? Exponential? 

That said, staged combustion engines have gas-liquid (or in the case of FFSC, gas-gas) injectors, which affords higher combustion efficiency in the main combustion chamber. It's not big enough to make up the difference, but the factor is there.

The relevant (approximate) relationship is this one from Sutton:

 

Wherein,

• k - Heat capacity ratio
• R - Specific gas constant
• R' - Universal gas constant
• M - Molecular weight
• T₁ - Chamber static temperature
• T₀ - Chamber stagnation temperature (assumed to be approximately equal to T₁)
• P₁ - Chamber pressure
• P₂ - Exit pressure (equal to the environment if optimally expanded)

The biggest assumption here is that of a constant heat capacity ratio, but there are others (such as no friction or heat transfer to the walls, no Rayleigh losses (all of these go together in the assumption of isentropic flow), single phase homogeneous ideal gas, velocity at the exit is purely uniaxial).

Plotting performance curves for RP-1/LOX (O/F = 2.7) and CH4/LOX (O/F = 3.2) via RPA gives the following, respectively:

Edited December 4, 2021 by Silavite
Assumptions

[sevenperforce]

1 hour ago, FleshJeb said:

@sevenperforce In addition to your recent calcs here, I've been running your numbers from:

https://forum.kerbalspaceprogram.com/index.php?/topic/156515-rocketlab-discussion-thread/&do=findComment&comment=4061314

and coming out with a negative dry mass of the first stage.

Expendable first stage: 5630m/s = 320s * 9.81m/s^2 * ln (480 / (4.6 + 79.1 + x))

It doesn't even cross zero until ISP = 329s

This suggests a problem with the assumed ISPs and/or the dry mass fraction of the second stage.

Yeah, something's tricky here.

Also, there are some baked-in assumptions with the mass of the second stage, too. There is volumetric margin for the second stage to be larger and heavier.

[FleshJeb]

10 minutes ago, sevenperforce said:

There is volumetric margin for the second stage to be larger and heavier.

You made a mistake on that post. The fairing is only 5m, not 7m.

[sevenperforce]

1 minute ago, FleshJeb said:

You made a mistake on that post. The fairing is only 5m, not 7m.

It looks like RocketLab made the mistake. Today, their website says the fairing is only 5 meters.

But this was the screenshot I took yesterday:

[Rakaydos]

7m max diamiter, but Peter Beck called it a "5m class fairing", so it only fits payloads that fit in "normal" 5m fairings.

 

[RyanRising]

16 hours ago, Silavite said:

That said, staged combustion engines have gas-liquid (or in the case of FFSC, gas-gas) injectors, which affords higher combustion efficiency in the main combustion chamber. It's not big enough to make up the difference, but the factor is there.

The relevant (approximate) relationship is this one from Sutton:

 

Wherein,

• k - Heat capacity ratio
• R - Specific gas constant
• R' - Universal gas constant
• M - Molecular weight
• T₁ - Chamber static temperature
• T₀ - Chamber stagnation temperature (assumed to be approximately equal to T₁)
• P₁ - Chamber pressure
• P₂ - Exit pressure (equal to the environment if optimally expanded)

The biggest assumption here is that of a constant heat capacity ratio, but there are others (such as no friction or heat transfer to the walls, no Rayleigh losses (all of these go together in the assumption of isentropic flow), single phase homogeneous ideal gas, velocity at the exit is purely uniaxial).

Plotting performance curves for RP-1/LOX (O/F = 2.7) and CH4/LOX (O/F = 3.2) via RPA gives the following, respectively:

It's pretty satisfying to know my estimate of "a little more than 10 extra seconds for an equivalent engine" has a sound grounding in the math, despite me doing absolutely none of the work to verify it. Thanks for actually checking!

[tater]

Live in ~3 hours

[Beccab]

3 minutes ago, tater said:

Live in ~3 hours

catch?

[tater]

8 minutes ago, Beccab said:

catch?

No, don't think so, the vehicle is black and white. The recovery boosters are now red I think.

Yeah:

 

 

[tater]

Swap, heading to SECO.

[tater]

 

 

[tater]

 

[FleshJeb]

Here’s a work-in-progress Neutron for RSS/RO. He’s soliciting input on the specs.

 

[JoeSchmuckatelli]

1 hour ago, FleshJeb said:

Here’s a work-in-progress Neutron for RSS/RO. He’s soliciting input on the specs.

 

So... what you're saying is that, "It works in KSP"!

[magnemoe]

On 12/4/2021 at 10:56 AM, Rakaydos said:

7m max diamiter, but Peter Beck called it a "5m class fairing", so it only fits payloads that fit in "normal" 5m fairings.

 

as it tappers, and just realized its not 4 way symmetric, only 2 forward fins and two strakes, the other two legs don't have them. 

Overall like the design, one question however. Assuming first stage follow an falcon 9 RTLS trajectory will exposing the satellite at separation be an problem? 
I assume its an reason falcon 9 drops the fairing some time into the second stage burn at an higher attitude. 
Yes the centaur is also an second stage inside fairing but its dropped later as the delta has SRB. 

[sevenperforce]

8 minutes ago, magnemoe said:

Overall like the design, one question however. Assuming first stage follow an falcon 9 RTLS trajectory will exposing the satellite at separation be an problem? 

I feel like there are some inefficiencies there, but since they are using only RTLS trajectory they will probably be more lofted generally.

[Silavite]

Continuing with the engine stuff, the Scott Manley and NSF interviews with Beck reveal that the chamber pressure is 1,500 psi. Plugging in the known data:

• Vacuum thrust: 1075 kN *
• Sea level thrust: 851.4 kN *
• Chamber pressure: 10.34 MPa (1,500 psi)

Along with some educated guesses:

• Contraction ratio: 1.67 (same as the H-1, which is in a similar thrust class)
• Combustion efficiency: 0.99 (Beck talked a bit about the injector in Tim Dodd's interview, and I'm guessing that they're using a coaxial swirl configuration from his somewhat oblique comments. That type supports excellent mixing/atomization characteristics and also has some throttling ability which would be needed for the upper stage.)
• Mixture ratio: 3.2 (guess which is near the I_SP maxima for this chamber pressure)
• Relative gas generator flow rate: 0.02 of the main thrust chamber mass flow rate (guess based on averages in Sutton and SP-125)

And a not-so-educated guess:

• Freezing area ratio: 2.5 (flow is assumed to be in chemical equilibrium until it reaches the point in the nozzle at which the area is 2.5 times the throat area; if somebody has a better guess for this, I would appreciate it!)

We can arrive at the following I_SP figures for sea level and vacuum:

• Sea level: 272.0 sec
• Vacuum: 344.4 sec

* These I_SP numbers seem to be unusually far apart for a first stage engine, but you have to consider that the ratio of vacuum thrust to sea level thrust (and thus vacuum I_SP to sea level I_SP, since mass flow rate is constant/choked at the throat) is really high for a first stage engine (1.26, whereas the Merlin 1D is 1.08, the RD-180 is also 1.08, and the H-1 is 1.13). This implies a high expansion ratio, but I was unprepared for how high; 37 according to RPA.

According to Sutton, an expansion ratio of 37 (for γ ~ 1.2 and pressure ratio of 100 at sea level) is within the incipient flow separation region, so I'm curious as to what kind of tricks Rocket Lab may be employing in their nozzle design:

Spoiler

Of course, trying to predict thrust/I_SP levels in an incipient flow-separation regime may not be the most reliable for a (relatively) simple program like RPA, so I'd take the exact results with a grain of salt, but the general idea still stands.

Edit: I forgot to include I_SP loss due to the gas generator in my original post (d'oh!)

Edited December 24, 2021 by Silavite
I forgot to include ISP loss due to the gas generator in my original post (d'oh!)

[Silavite]

Also... one other thought. There was a lot of talk about simplifying operations, but Neutron is planning to balloon tanks in its upper stage (which are not exactly easy to handle). Admittedly having the launch site right next to the production facility should simplify operations in regards to using such a structure (no transportation), but I'm still wondering if they have any other special procedures for working with balloon tanks.

[tater]

 

 

Edited February 15, 2022 by tater