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18 hours ago, tater said:

7 min and counting

 

When I watch the 'in-space' flight, prior to SECO - I see little puffs of color (orange flares in the exhaust).  

I presume this is indication of turbulence or incomplete mixing of the fuels (including oxidizer in this)?

Given that it is clearly a successful rocket, would there be any reason for tweaking the combustion chamber based on the little bit flaring we do see? 

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If I'm not mistaken, due to the electric pumps they can vary the combustion ratio on the fly during the flight.  I wonder if they use that ability during the end of the flight as a method of throttling right before separation.

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So they're going with oxygen-rich staged combustion rather than the gas generator we had originally anticipated.

Looking at 303 seconds of sea level specific impulse, 329 seconds of vacuum specific impulse for the sea level engines, and 367 seconds of specific impulse for the vacuum engine.

The "not a capsule announcement" is about what I would have expected for not really having much of a plan yet. But it would be cool if they were thinking about making the service module integrated with the second stage.

Very interesting that the staging structure is NOT what I had originally thought.

cursed.png

I had imagined that the upper stage and the payload was all contained within the moveable fairing, but in actuality the upper stage is contained within the whole fixed upper half of the launch vehicle.

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It's also very interesting that they provided the expendable, reusable, and RTLS payloads specifically (15 tonnes, 13 tonnes, and 8 tonnes respectively). The difference between the downrange recovery and expended recovery is remarkably small. Definitely enough information that we can play around with it and maybe come up with some dry mass numbers.

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18 minutes ago, sevenperforce said:

It's also very interesting that they provided the expendable, reusable, and RTLS payloads specifically (15 tonnes, 13 tonnes, and 8 tonnes respectively). The difference between the downrange recovery and expended recovery is remarkably small. Definitely enough information that we can play around with it and maybe come up with some dry mass numbers.

Not too dissimilar from F9. Slightly better reusable (ASDS type landing, I presume) at a ~13% hit, and slightly worse RTLS (46% hit vs ~40%—though that's a SpaceX claim, not sure RTLS has ever been pushed to enough of a limit we have good data).

32 minutes ago, sevenperforce said:

The "not a capsule announcement" is about what I would have expected for not really having much of a plan yet. But it would be cool if they were thinking about making the service module integrated with the second stage.

Yeah, and I have to say, I love the look of  the capsule. It's like playing KSP with one set of mods plus stock, then there's a fancy new capsule mod that does not blend in, and looks super sci fi, lol.

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Interesting that there going closed cycle on the engine. Peter Beck was previously adamant a simple gas gen engine was all the performance needed, in combination with lower structural mass fractions from using carbon fibre construction.

The fact the performance seems about the same despite uprated engines makes me think they've had mass creep somewhere.

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6 hours ago, RCgothic said:

Interesting that there going closed cycle on the engine. Peter Beck was previously adamant a simple gas gen engine was all the performance needed, in combination with lower structural mass fractions from using carbon fibre construction.

The fact the performance seems about the same despite uprated engines makes me think they've had mass creep somewhere.

That seems also confirmed by the fact that despite the previous attachment to RTLS only they've since switched to a Spacex-like droneship

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14 hours ago, sevenperforce said:

It's also very interesting that they provided the expendable, reusable, and RTLS payloads specifically (15 tonnes, 13 tonnes, and 8 tonnes respectively). The difference between the downrange recovery and expended recovery is remarkably small. Definitely enough information that we can play around with it and maybe come up with some dry mass numbers.

I assume they designed it so it don't need an upper atmosphere braking burn like falcon 9 does,  also assume they stage later than falcon 9 say there falcon 9 drops fairing, this require an longer return burn requiring more fuel. 

I wonder why they go oxygen rich with methane? would not fuel rich be easier to work with? 

Edited by magnemoe
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3 hours ago, magnemoe said:

I assume they designed it so it don't need an upper atmosphere braking burn like falcon 9 does,  also assume they stage later than falcon 9 say there falcon 9 drops fairing, this require an longer return burn requiring more fuel. 

That’s counter-intuitive, though, because a longer first stage burn (to enable the fairing drop and staging to be simultaneous) means a hotter re-entry, which you’d think would require a braking burn.

Maybe if they leave the clamshell open for initial shuttlecock entry, like Stoke is planning, that will help. 

3 hours ago, magnemoe said:

I wonder why they go oxygen rich with methane? would not fuel rich be easier to work with? 

Well oxygen is denser so you can extract more torque in a smaller turbine. That’s probably the first reason. Using ORSC also gives you a higher overall O:F ratio which reduces both  your tank volume and your propellant costs. Finally, while coking isn’t typically a problem with methane, ORSC is much better-studied and has a longer legacy, and it certainly obviates any possibility of coking from any impurities that end up in the CH4.

IIRC, the RS-25’s FRSC cycle actually uses dual turbines running off a single fuel-rich preburner because trying to build a single-shaft FRSC engine is just too complicated. 

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20 minutes ago, sevenperforce said:

That’s counter-intuitive, though, because a longer first stage burn (to enable the fairing drop and staging to be simultaneous) means a hotter re-entry, which you’d think would require a braking burn.

Maybe if they leave the clamshell open for initial shuttlecock entry, like Stoke is planning, that will help. 

Well oxygen is denser so you can extract more torque in a smaller turbine. That’s probably the first reason. Using ORSC also gives you a higher overall O:F ratio which reduces both  your tank volume and your propellant costs. Finally, while coking isn’t typically a problem with methane, ORSC is much better-studied and has a longer legacy, and it certainly obviates any possibility of coking from any impurities that end up in the CH4.

IIRC, the RS-25’s FRSC cycle actually uses dual turbines running off a single fuel-rich preburner because trying to build a single-shaft FRSC engine is just too complicated. 

I think they can shield against the heat, for falcon 9 I assume this require lots of redesigns of rocket and engine but its easier with an new design superheavy will also not require an braking burn before entering the atmosphere.  The engine nozzles can handle some heat and the skirt give decent protection itself.

Buying your arguments for oxygen rich, yes its an known technology, RS-25 burned hydrogen and the huge difference in density and volume pumped made it hard to run it all on one shaft. 
Now their benefit is that controlling this engine will be much simpler than an raptor as its just one preburner. 

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21 hours ago, sevenperforce said:

IIRC, the RS-25’s FRSC cycle actually uses dual turbines running off a single fuel-rich preburner because trying to build a single-shaft FRSC engine is just too complicated. 

I know this is kinda off topic, but while you remember right about the dual turbines, each also had their own separate preburner: SSME-Powerhead-with-preburners-turbopump

That said, you are onto something. Pratt and Whitney also thought it would be a fine idea to stick both turbines onto a common preburner: D32A535C-3A22-411D-A60A-6E176799E038.jpg

This didn’t end up actually being used as an SSME powerhead, but I have it on good authority that the XLR-129 which led to this proposal also heavily influenced the block upgrades that did fly on later engines, and which are being used as SLSME*s today!

*Space Launch System Main Engine, which I just made up. They actually call them RS-25s nowadays.

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On 9/22/2022 at 9:00 AM, sevenperforce said:

nitial shuttlecock entry

Any idea how deep into the atmosphere the ship would have to get for the fairings to 'shuttlecock' the fat end around? 

Although - rereading, the image is wrong.  Fat end of shuttlecock is fluffy, not heavy... Except Fluffy now means 'big boned'... And dense means stupid...

What is happening with the language? 

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4 hours ago, RyanRising said:
On 9/22/2022 at 9:00 AM, sevenperforce said:

IIRC, the RS-25’s FRSC cycle actually uses dual turbines running off a single fuel-rich preburner because trying to build a single-shaft FRSC engine is just too complicated. 

I know this is kinda off topic, but while you remember right about the dual turbines, each also had their own separate preburner: SSME-Powerhead-with-preburners-turbopump

Oh, dang, that's even worse than I remembered.

You'd think that with completely separate preburners, they would have worked on varying the mixture ratios to optimize thrust and Isp since the pumps are completely independent, but evidently the RS-25 maintains a single mixture ratio at all times. Weird.

And technically the RS-25 is both a FRSC engine *and* a closed expander cycle engine, since the fuel is pumped around the chamber and engine bell to cool it, then that supercritical fuel expands through a low-pressure turbopump to step up the propellant pressure before it enters the main turbopump:

Spoiler

1280px-Ssme_schematic_(updated).svg.png

The oxidizer flow also uses a low-pressure step-up pump, but rather than using a LOX expander cycle for this, they just use a tapoff from the main oxygen turbopump output and route it back into the low-pressure pump. I suppose that avoiding a dual expander cycle reduces the possibility of the fuel and oxidizer mixing if some of the cooling loops ruptured.

Since the RS-25 uses autogenous pressurization, it has a heat exchanger around the oxygen turbopump to boil the LOX to provide a gas source.

Because the RS-25 uses a fuel-rich preburner to operate the oxidizer turbopump, it had to have a separate helium tank to provide a constant flow of helium to continually purge the cavity between the fuel and the oxidizer. The only thing that doesn't need the constant helium purge would be a proper FFSC engine like Raptor.

The engine I was thinking of was not actually the RS-25, but the RS-68. It has a single burner (although it's a gas generator, not a preburner) that exhausts into separate turbines which run the pumps independently:

Spoiler

JwpV8.png

The two gas generator exhausts also provide single-axis roll control.

The Chinese YF-77, which is essentially a clone of the RS-68 (except that it uses active dump cooling on the nozzle instead of ablative cooling like the RS-68) does the same thing, with a single gas generator and two separate turbines with two separate exhausts. They fly the YF-77 in pairs on the Long March 5 core, though, so I don't know whether they use the nozzles for roll control at all.

Spoiler

YF-77_.jpg

As always, I highly recommend reading this amazing blog post about the many possible complexities of just one type of engine cycle.

1 hour ago, JoeSchmuckatelli said:

Any idea how deep into the atmosphere the ship would have to get for the fairings to 'shuttlecock' the fat end around? 

Although - rereading, the image is wrong.  Fat end of shuttlecock is fluffy, not heavy... Except Fluffy now means 'big boned'... And dense means stupid...

The engines on Neutron's first stage will always keep the center of gravity extremely low, just like with Falcon 9's first stage. Opening the fairings would definitely pull the center of lift/drag further back, but that's not going to be a major issue since the center of gravity will already be low enough. I was just thinking in terms of additional drag.

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On 9/21/2022 at 2:31 PM, tater said:

FdMunoaXgAMLxus?format=jpg&name=large

This is clearly just preliminary concept art, but I spy four reasonably large engine bells peeking out from what appears to be the service module.

Given that there is no apparent intention to make the upper stage reusable, it might make sense to give the service module a ring of simple, high-thrust, low-efficiency methalox engines (preferably ignited with a solid pyro) plumbed through the PAF to the upper stage itself. Then, in a low-altitude abort, those engines would drain the upper stage in a matter of moments, pulling it and the capsule free of the booster:

possible.png

This would eliminate the need for carrying the abort propellant on the service module like Starliner (which means less total mass to orbit), and also avoids plumbing the capsule itself with the same propellant used for abort like Crew Dragon. The abort motors would be pretty thoroughly oversized but they'd certainly have no lack of propellant.

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For anyone who might find it useful, I went ahead and did a pixel count with known values (total vehicle height and second stage height) to estimate the dimensions of everything in Neutron so far:

Neutron-Pixel-Count.png

Interesting that the "5 meter diameter" fairing appears to be the true internal envelope diameter (as this is the diameter of the upper stage) rather than a measurement of the outside of the fairing, which is notably larger. The 7 meter diameter does not include the landing legs.

I'm sure someone can drum up tank volumes from this.

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21 hours ago, sevenperforce said:

This is clearly just preliminary concept art, but I spy four reasonably large engine bells peeking out from what appears to be the service module.

Given that there is no apparent intention to make the upper stage reusable, it might make sense to give the service module a ring of simple, high-thrust, low-efficiency methalox engines (preferably ignited with a solid pyro) plumbed through the PAF to the upper stage itself. Then, in a low-altitude abort, those engines would drain the upper stage in a matter of moments, pulling it and the capsule free of the booster:

possible.png

This would eliminate the need for carrying the abort propellant on the service module like Starliner (which means less total mass to orbit), and also avoids plumbing the capsule itself with the same propellant used for abort like Crew Dragon. The abort motors would be pretty thoroughly oversized but they'd certainly have no lack of propellant.

That sounds problematic you have to take the second stage with you? Benefit of the falcon 9 system is that the abort system fuel can be used for the rcs system. Now you could have pressurized methane and oxygen tanks for abort and then pipe them back into second stage once second stage is well under way. 

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I've made a few slight modifications to my pixel counting approach from Friday, now with added tank volumes (I assumed that the spherical caps were geometrically perfect and that the volumetric split on the upper stage was equal to the measurable volumetric split on the booster):

index.php?action=dlattach;topic=53194.0;

Based on these numbers (and using standard density for LOX and LCH4) we would be looking at an O:F ratio of 2.99 which seems shockingly fuel-rich. Of course, it is possible that the engineering in this image is inexact with respect to the positioning of the common bulkhead and the O:F ratio is closer to 3.4-3.6 or even higher.

If the noted O:F ratio holds, however, then we are looking at 363 tonnes of propellant on the booster and 84 tonnes of propellant on the upper stage.

The design proposed in December of last year (which is still what shows up on their site) had a launch mass of 480 tonnes, which would mean a ~6.9% total vehicle dry mass percentage, which is pretty impressive but not out of the realm of possibility for carbon-fiber construction. But GLOW could have changed since December; the vehicle height went from 40 m to 42.8 m. Total liftoff thrust then, with 7 G/G methalox engines, was 5.96 MN; it is now 6.61 MN (165 klbf x nine engines). I'm not sure if they said exactly why they went from 7 engines to 9; if it was GLOW increase and they were adding engines to maintain the same 1.27:1 TWR at liftoff, then we'd be looking at 532 tonnes GLOW, which (with this prop load) corresponds to 16% dry mass percentage which is honestly pretty bad. More likely that the increased engine count will merely increase TWR.

The engines themselves lost 14% of their sea level thrust but of course 9 is 29% more than 7.

On 9/24/2022 at 1:07 PM, magnemoe said:
On 9/23/2022 at 3:18 PM, sevenperforce said:

Given that there is no apparent intention to make the upper stage reusable, it might make sense to give the service module a ring of simple, high-thrust, low-efficiency methalox engines (preferably ignited with a solid pyro) plumbed through the PAF to the upper stage itself.

That sounds problematic you have to take the second stage with you? Benefit of the falcon 9 system is that the abort system fuel can be used for the rcs system.

Yeah, I don't think my idea was particularly well thought through. That upper stage is far too heavy.

Both Dragon 2 and Starliner allow the abort system fuel to be used for the RCS/OMS system. The difference is that Starliner keeps the abort engines, abort propellant, and RCS/OMS engines separate from the capsule, which has only monopropellant RCS for pointing during re-entry. Dragon 2 keeps everything onboard. Advantage for Dragon 2 is being able to reuse the abort engines and prop tanks and OMS/RCS; disadvantage is that all of those systems are inside the capsule with the crew which is fundamentally more risky than having them in a separate service module.

And of course Starliner has no space for unpressed cargo.

I wonder if RocketLab would use Rutherford engines for abort. Electric turbopumps are good for quick starts, after all. But you'd probably need more than four, since all together they'd only push 100 kN which would give a notional 10-tonne crew capsule only about a gee of acceleration. And using kerolox for OMS would be...interesting. They've already got experience with appropriate RCS via Curie and HyperCurie.

On 9/16/2022 at 11:52 AM, Geonovast said:

If I'm not mistaken, due to the electric pumps they can vary the combustion ratio on the fly during the flight.  I wonder if they use that ability during the end of the flight as a method of throttling right before separation.

For Electron, yes. Neutron will be using staged combustion with a single oxygen-rich preburner and a single turbine to drive both the LOX turbopump and the CH4 turbopump. Varying the combustion ratio would require variable gearing between the shaft and the CH4 turbopump which would add an incredible amount of complexity and additional failure points.

Edited by sevenperforce
Notation about RCS/OMS
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On 9/27/2022 at 12:15 PM, sevenperforce said:

Based on these numbers (and using standard density for LOX and LCH4) we would be looking at an O:F ratio of 2.99 which seems shockingly fuel-rich. Of course, it is possible that the engineering in this image is inexact with respect to the positioning of the common bulkhead and the O:F ratio is closer to 3.4-3.6 or even higher.

If the noted O:F ratio holds, however, then we are looking at 363 tonnes of propellant on the booster and 84 tonnes of propellant on the upper stage.

Someone over on NSF pointed out that these are probably ellipsoidal caps, not spherical caps, so I recalculated.

While the lower (assumed CH4) booster tank volume somewhat intuitively remains the same, the upper (assumed LOX) booster tank volume goes up from 238.3 cubic meters to 255.4 cubic meters, bringing the apparent mixture ratio up from 2.99 to to 3.2 which is still surprisingly fuel-rich but not as severely as before.

Booster propellant mass goes up, from 363 tonnes to 382 tonnes, and upper stage prop mass goes up from 84 to 90 tonnes. If the 1.27:1 TWR holds then we are looking at a total vehicle dry mass percentage of 11.1% which makes sense.

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On 9/27/2022 at 9:15 AM, sevenperforce said:

For Electron, yes. Neutron will be using staged combustion with a single oxygen-rich preburner and a single turbine to drive both the LOX turbopump and the CH4 turbopump. Varying the combustion ratio would require variable gearing between the shaft and the CH4 turbopump which would add an incredible amount of complexity and additional failure points.

Mixture ratio control isn’t necessarily something that’s exclusive to electric or double-shaft pumps - the RD-180, a similar single shaft ORSC engine to the one proposed here, features some degree of mixture ratio control despite that, and I don’t think they were doing any variable gearing nonsense.

Edited by RyanRising
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