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The Tech Behind SpaceX’s New Engine

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Re: The Tech Behind SpaceX’s New Engine

#121

Post by newnewpdro »

avmich wrote: post #116
newnewpdro wrote: post #106

Earlier quoted context omitted.

I'm confused; wouldn't the oxidizer-rich half run hotter? Isn't that challenge the Russians overcame but the US punted on back in the days of the space shuttle? Your post and the referenced wikipedia article claims the turbines run cooler. Wouldn't that only be true for the fuel-rich side?

> wouldn't the oxidizer-rich half run hotter? You try running it as hot as possible, but without burning metal. You have about twice as much liquid oxygen by volume than kerosene in a typical engine, but you can heat kerosene to higher temperatures without burning. So... I don't think you'll get higher temperatures with oxygen gas generator?

The hackaday article explicitly states the oxygen-rich preburner exhaust was extremely hot and required metallurgical advances before turbines could survive it.

How is the Raptor's oxygen-rich side any different from what the Russian's were dealing with?

The article left me with the impression that the Raptor is using similarly resilient materials for the oxygen-rich side, with its language describing the US and Russian designs solving the two critical halves of the problem.

There just seems to be a contradiction in the statements surrounding the turbines running cooler.

Re: The Tech Behind SpaceX’s New Engine

#122

Post by erobbins »

avmich wrote: post #117
erobbins wrote: post #114

Earlier quoted context omitted.

I'm not clear what the efficiency gains are here at all over a SSME style staged combustion cycle. In both designs all propellants go through the chamber. Comparing to a gas generator cycle there are gains, sure. It seems like the major win here is running oxygen rich on the oxygen side and fuel rich on the fuel side to reduce the chance of a leak destroying the whole vehicle. Am I missing something?

> I'm not clear what the efficiency gains are here at all over a SSME style staged combustion cycle. You have higher pressure in the chamber with full-flow approach. Since the mass flow is the same, and Isp is higher, thanks to higher pressure, you're winning. And you have higher pressure because you supply more power to pumps. Reducing risks to destroy the vehicle due to leaks is, I agree, very secondary benefit. Ma…

But... you don't.

The turbines in any kind of staged or full flow system run at a pressure higher than the chamber is at... I don't remember the numbers offhand but I want to say the SSME's preburners were at 8000ish psi, with a drop to 4000ish through the turbines, and finally to 3000 in the main combustion chamber. The SSME was only partially preburned, and so the flow through the turbine has to expand over a higher pressure ratio, but I still don't see a fundamental difference.

Re: The Tech Behind SpaceX’s New Engine

#123

Post by avmich »

newnewpdro wrote: post #121
avmich wrote: post #116

Earlier quoted context omitted.

> wouldn't the oxidizer-rich half run hotter? You try running it as hot as possible, but without burning metal. You have about twice as much liquid oxygen by volume than kerosene in a typical engine, but you can heat kerosene to higher temperatures without burning. So... I don't think you'll get higher temperatures with oxygen gas generator?

The hackaday article explicitly states the oxygen-rich preburner exhaust was extremely hot and required metallurgical advances before turbines could survive it. How is the Raptor's oxygen-rich side any different from what the Russian's were dealing with? The article left me with the impression that the Raptor is using similarly resilient materials for the oxygen-rich side, with its language describing the US and Russ…

> The hackaday article explicitly states the oxygen-rich preburner exhaust was extremely hot

I didn't find this exact quote, but the idea is the following: oxygen-rich flow can get hot enough to ignite metal. What Russians did was to find the materials withstanding ignition best and find temperatures which are still safely avoiding ignition. This combination turned out good enough to use oxygen-rich gas generators.

Raptor may be using similar approach. "Two sides of the problem" language is, in my opinion, about the tradition to have fuel-rich turbines in America and oxygen-rich in Russia, and for full flow engine you need both.

The turbine can run cooler, because you don't need to have gases from gas generator that hot in order to get enough power to the pumps. You're getting that power from the fact that you use full flow of both fuel and oxidizer, and since you have that additional source of power, you can relax requirement of the gas generator exhaust temperature.

Re: The Tech Behind SpaceX’s New Engine

#124

Post by avmich »

erobbins wrote: post #122
avmich wrote: post #117

Earlier quoted context omitted.

> I'm not clear what the efficiency gains are here at all over a SSME style staged combustion cycle. You have higher pressure in the chamber with full-flow approach. Since the mass flow is the same, and Isp is higher, thanks to higher pressure, you're winning. And you have higher pressure because you supply more power to pumps. Reducing risks to destroy the vehicle due to leaks is, I agree, very secondary benefit. Ma…

But... you don't. The turbines in any kind of staged or full flow system run at a pressure higher than the chamber is at... I don't remember the numbers offhand but I want to say the SSME's preburners were at 8000ish psi, with a drop to 4000ish through the turbines, and finally to 3000 in the main combustion chamber. The SSME was only partially preburned, and so the flow through the turbine has to expand over a highe…

Let me explain. The principal advantage of full flow system over regular staged combustion is higher Isp of the engine.

Higher Isp of the engine is achieved because the pressure in the chamber of full flow engine is higher, than chamber pressure of staged combustion engine.

The pressure in the chamber is higher because pumps produce higher pressure.

Pumps can produce higher pressure, because with full flow pumps have more power driving them than with staged combustion. Power here is roughly equal to volume flow (m^3 / s) multiplied by pressure drop (N / m^2). Full flow scheme gives you more volume flow, since all the propellant is used to turn turbines.

So even though all the propellant in both cases is burned in the chamber, Isp is higher with full flow scheme. Rocket engine maximizes momentum of the exhaust, not the energy, which is conserved.

Re: The Tech Behind SpaceX’s New Engine

#125

Post by gpm »

avmich wrote: post #123
newnewpdro wrote: post #121

Earlier quoted context omitted.

The hackaday article explicitly states the oxygen-rich preburner exhaust was extremely hot and required metallurgical advances before turbines could survive it. How is the Raptor's oxygen-rich side any different from what the Russian's were dealing with? The article left me with the impression that the Raptor is using similarly resilient materials for the oxygen-rich side, with its language describing the US and Russ…

> The hackaday article explicitly states the oxygen-rich preburner exhaust was extremely hot I didn't find this exact quote, but the idea is the following: oxygen-rich flow can get hot enough to ignite metal. What Russians did was to find the materials withstanding ignition best and find temperatures which are still safely avoiding ignition. This combination turned out good enough to use oxygen-rich gas generators. R…

> but the idea is the following: oxygen-rich flow can get hot enough to ignite metal.

A relevant example of igniting metal being the first Raptor test firing, which was slightly "engine rich". Which we could see by the green given off by burning copper (and later confirmed by Elon, also somehow fixed).

Video: https://twitter.com/elonmusk/status/1092270756715737088

Re: The Tech Behind SpaceX’s New Engine

#126

Post by InclinedPlane »

trhway wrote: post #58
debt wrote: post #56

"For example, the turbine of the V-2 rocket was spun with steam ..." Steam technology and rocket technology have a shared history. That kind of bottleneck shape that a is quintessential shape of a rocket engine, is actually a Steam-Engine-era technology called a de Laval nozzle. I didn't know they used a tiny steam engine inside a V-2. I think it's cool that the design of such an old technology, the steam engine, liv…

>I didn't know they used a tiny steam engine inside a V-2. it wasnt steam engine per.se., it was catalytic peroxide decomposition.

The Soyuz first stage uses the same system even today.

Re: The Tech Behind SpaceX’s New Engine

#127

Post by CarVac »

ballooney wrote: post #83
CarVac wrote: post #9

Earlier quoted context omitted.

A full-flow staged combustion engine has more efficiency due to both propellants being fully vaporized before they mix, rather than a liquid-liquid mixture (as in a gas generator engine) or a liquid-gas mixture (as in a normal staged combustion engine). The seals are between the turbine side (hot gas) and compressor side (liquid fuel) of the turbopump. In a full-flow staged combustion engine, you don't have to worry…

No, this is all just wrong. The efficiency is because you dont throw a load of enthalpy overboard as in a gas generator. The phase is _exremely_ second order by comparison. You’re all over this thread being wrong and confused. Source: am a rocket engine designer.

I was under the impression that the question was regarding FFSC versus non-FF staged combustion.

You say the main benefit is just increased pressure?

Re: The Tech Behind SpaceX’s New Engine

#128

Post by mannykannot »

The F1 engines of the Saturn V 1st. stage used a rich-mixture gas generator to drive the turbopump, but its exhaust was fed into the engine's nozzle about halfway down, through an annular manifold [1]. Up to that point, the combustion chamber and nozzle were cooled by circulating fuel through them, but beyond that point, the cooler turbopump exhaust layer protected the nozzle extension.

In pictures of launches [2], you can make out the brown smoky annulus of the turbopump exhaust, for a distance about equal to the length of the nozzle extension, until it either mixes with the hot exhaust, or with ambient air and then burns, at which point the smoke particles become incandescent.

[1] https://history.msfc.nasa.gov/saturn_apollo/documents/F-1_En...

[2] https://images.nasa.gov/details-ksc-69pc-442.html

Re: The Tech Behind SpaceX’s New Engine

#129

Post by usrusr »

twic wrote: post #12
kumarvvr wrote: post #6

This is an excellent article. Just curious, how much more efficient is the full-flow engine? And, what exactly is the deal with the seals the article is talking about? Anyone have more info?

I think this is about safety. You need two pumps, one for fuel, one for oxygen. The compressor side of the pump will be full of whatever it is you're pumping. But if you use a single type of preburner, then the turbines of both pumps are full of either hot fuel gas, or hot oxygen gas. That means you have one pump where you have hot fuel or oxygen right next to cold oxygen or fuel. If the seals leak, the pump blows up…

In short: you accept a higher failure rate (twice as many pumps) to get a lower explosive failure rate (much less explosive pumps). A very good tradeoff when building a rocket with a high number of engines and some redundancy which does not help against explosions. It gets even better when the pumps turn out to be more reliable individually due to relative simplicity.

Re: The Tech Behind SpaceX’s New Engine

#130

Post by Gravityloss »

The article is a bit incomplete. The space shuttle main engine already had two main turbopumps. IIRC this is because the optimal pump speed is different for pumping hydrogen and oxygen. They have very different densities. You want to avoid gearing as much as possible.

And the main problem problem in the staged combustion cycle - in feeding the gas generator exhaust to the main chamber is not that it's fuel rich or oxidizer rich - it's that it's usually much lower pressure than in the main combustion chamber, because it had to go through the turbine, which causes the pressure to lower.

That's why engines like NK-33 have a separate boost pump for the gas generator. RD-180 has more pump stages for the fuel entering the gas generator (all oxidizer passes through the gas generator).

Full flow staged combustion is another way to solve this - put all the propellants through the gas generators and turbines.

I realize it's hard to write popular technical articles about medium complexity subjects and sometimes you have to take some shortcuts.

https://en.wikipedia.org/wiki/Space_Shuttle_main_engine#/med...

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