TL; DR Full flow lowers turbine temperatures at the expense of parts complexity. Given turbopumps are the devil’s ass part of rocketry, this has been a sought-after technology. The pay-off isn’t so much efficiency as much as longevity. (This also explains why full flow hasn’t been a priority for anyone until SpaceX.)
Ehhh. I'd argue it also lowers part complexity, because it eliminates interpropellant seals.
> The pay-off isn’t so much efficiency
It improves achievable chamber pressure (assuming similar maximum pressures and temperatures at the turbines), which improves both thrust and Isp.
> (This also explains why full flow hasn’t been a priority for anyone until SpaceX.)
The reason it hasn't been a priority is more that no new development has been a priority since the 70's.
But it adds the whole second gas generator and turbine. No, it doesn't lower part complexity as a whole :) .
> It improves achievable chamber pressure
Yes, I'd even argue that's the penultimate goal of full-flow scheme.
1) For a rocket engine you usually (almost always) want as high Isp (which is approximately the speed of gasses flowing from the rocket engine) as possible.
2) To get that Isp, you need as high pressure in the chamber as possible.
3) To get that high pressure, you want to supply your pumps with as much power (in Watts) as possible from the given overall fuel flow (in kilograms per second) - while not melting turbines blades (so fuel:oxidizer ratio is limited).
4) To get most power to the turbines blades at fixed (maximum) gas temperature and fixed overall fuel flow, you want to use all of that flow (classical staged combustion uses only one component, which is essentially not using the whole fuel flow) and adjust ratios fuel:oxidizer at both gas generators so that total power would be maximum. You want maximum flow both because that increases turbine efficiency and because that actually provides more power to the turbine.
Because the fuel side and oxidizer side typically have a very different flow rates, many engines have separate pumps and turbines for them on separate shafts even if both sides use same fluids in the preburner. For example, see the SSME.
With FFSC, each of the turbines have no sealing requirements. This means, among other things, that bearings become easier and safer, as you can afford larger margins in your design. In many ways, FFSC allows the individual parts of the engine to be simpler.
The main cost of doing it over gas generators (other than oxidation-resistant superalloys...) is that all parts of the engine sort of circularly depend on each other -- when the Merlin engine was in very early development, they drove into the desert and ran the gas generator and fuel pump on it's own with basically no support infrastructure. Testing Raptor in a similar way is just not possible. It would have never gotten built without the NASA Stennis facility that can provide all the intermediate fluids at the pressures required.
I fully agree with the rest.
By "many" you perhaps mean "many American". In Russian engines it's mostly a single shaft.
> The main cost of doing it over gas generators (other than oxidation-resistant superalloys...) is that all parts of the engine sort of circularly depend on each other...
Yes, rocket engine is usually a complex dynamic system, with deep feedback loops. You can still get gas generator and turbines with pumps tested separately, but you need to measure dynamic properties - like resonance frequencies - in order to have better chances of the good work when the system is integrated.
And by "many American," do you perhaps mean "pretty much just the Space Shuttle Main Engine?" Since that one, has America launched any new staged combustion engine off the ground?
Sounds like SpaceX has, with NASA and the State of Mississippi modified the test stand to support methane.
So for the layman, are you saying that they've got a simpler turbine, use more of them, but not twice as many, because the design means you need fewer engines to lift the same payload?
Full-flow scheme has 2 gas generators - with different fuel:oxidizer ratios, and 2 turbines, each driving one pump of the same propellant which constitutes the main flow through the turbine. Staged combustion has 1 gas generator and (usually) 1 turbine with 2 centripetal pumps, so one of the pumps has the flow of the propellant opposite to the main propellant of the turbine. Here is the danger, if seals don't hold.
Your post and the referenced wikipedia article claims the turbines run cooler. Wouldn't that only be true for the fuel-rich side?
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?
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.
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.
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).
>According to SpaceX founder Elon Musk, this design was explicitly intended to achieve full reusability of all rocket stages and, as a result, "a two order of magnitude reduction in the cost of spaceflight" (wikipedia also)
And they hope to make the oxygen and methane on Mars for the return journey, using CO2 and ice found on there plus energy from solar cells.
It will be impressive if they do any of those.
Here is a fantastic talk from the NVIDIA conference:
It wouldn't be shocking if they've removed some ITAR-covered parts so no one can sneak up in the middle of the night and steal rocket tech, but it appears to be the actual booster.
2020? I honestly doubt we'll see the SLS launch before 2024 at it's current rate.
In 2017 the SLS launch was about 2 years out, but likely to slip: https://www.nasaspaceflight.com/2017/11/sls-managers-troops-...
In 2019 the SLS launch is a bit under 2 years out, but likely to slip: https://arstechnica.com/science/2019/02/nasa-still-working-t...
This strongly reminds me of the fact that late software projects are promised to be on time until about 6 weeks before launch, and then launch keeps getting delayed. And this is true no matter how late it eventually turns out to be. The reason why 6 weeks is the magic figure is that for a software project, that's the point where you can no longer paper over the inevitability of failure with wishful thinking. Rockets have slower schedule, but it is strongly looking like 2 years is a similar magic figure in that industry, for similar reasons.
In that light, the money quote from the second article is this:
However, the agency and its prime contractor for the core stage, Boeing, are on a tight timeline that has little margin for technical problems that might occur during the structural tests of the tank or the green run tests. Historically, during this integration and test process with other large rocket programs, major problems have often occurred.
I am generally a believer that what happened historically shouldn't be ignored. There is therefore no way that the SLS will launch in 2020. Or 2021. In fact nobody really knows when it will launch. Furthermore the upper stage, aka Orion, is apparently in even worse shape. Right now they are going back to the drawing board to try to find a design that gets costs down.
Admittedly SpaceX itself is promising the BFR in about 2 years, and also had a history of overruns. I don't think that they will launch on time. But they have a better history of getting launch vehicles up.
I will therefore happily take your bet for $100, but I would like to formalize it a bit. I win if Falcon Super Heavy+Starship or whatever it gets renamed to gets successfully launched to orbit first OR if SLS+Orion gets canceled first. Vice versa you win if SLS+Orion gets successfully launched first OR if Falcon Super Heavy+Starship or whatever gets canceled first. Note that "SpaceX goes out of business" counts as canceled, even if someone else (eg Bezos) buys the remnants and then launches something based on the work.
If my version is acceptable, you can contact me by email per my profile.
That said, I'm still on the side of betting with history. In most organizations, the deadline is the first date that nobody can (yet) disprove. When a deadline depends on problems not happening on this project that historically have been common, I think it is safe to bet that history will repeat itself.
This goes doubly for the SLS. Which is more ambitious than past launch systems, and is being built so long after the last new launch system was designed by the companies involved that there is little institutional knowledge left about how to do it. (Furthermore building with competing companies contracting for pieces that need to integrate just sounds like a recipe for expensive overruns to me.)
As opposed to the BFR. Which is being designed by a company with more recent experience of how to build new launch systems than the rest of the planet put together.
The Starship hopper will definitely be working long before either of those two take flight.
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.
In pictures of launches , 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.
And, what exactly is the deal with the seals the article is talking about?
Anyone have more info?
In terms of efficiency, it's worth mentioning that even tiny gains have big impacts because of the rocket equation. Having a 1% boost the entire time the rocket is running can have an enormous impact in the overall velocity, or reduce the overall amount of fuel needed for the same targets.
Compared to a oxygen-rich full flow you can reach a higher pressure while having better safety margins which will allow you to re-use the engine. The higher your chamber pressure the closer the rocket's thrust at sea level is to its maximum thrust in vacuum. Also higher pressure tends to correlate with higher thrust to weight ratios.
Also, having both inputs already be gasses mean they mix better. This will have a tiny benefit in them burning more completely. But more importantly they'll burn better across a wider range of thrusts, which could be important for throttling down the engine for landing.
Also, and importantly for the design of the engine, it's much much easier* to computationally model the mixing of two gases, to ensure complete mixing and no combustion instability (especially when the reaction components are limited to ch4, o2, and partial- and complete-reaction products like c02, c0, h2o, oh-, ...).
* by much easier, i mean less impossible
Raptor: Thrust 1900 KN / 330 specific impulse
Merlin 1D: 480KN / 275 s.i.
i can't find the weight of the Raptor. Merlin is 630Kg
Also some people might not realize how much liquid is being pumped by rocket engines (the Saturn V F-1 could drain a 30,000 gallon swimming pool in 10 seconds):
Each second, a single F-1 burned 5,683 pounds (2,578 kg) of oxidizer and fuel: 3,945 lb (1,789 kg) of liquid oxygen and 1,738 lb (788 kg) of RP-1, generating 1,500,000 lbf (6.7 MN; 680 tf) of thrust. This equated to a flow rate of 671.4 US gal (2,542 l) per second; 413.5 US gal (1,565 l) of LOX and 257.9 US gal (976 l) of RP-1. During their two and a half minutes of operation, the five F-1s propelled the Saturn V vehicle to a height of 42 miles (222,000 ft; 68 km) and a speed of 6,164 mph (9,920 km/h). The combined flow rate of the five F-1s in the Saturn V was 3,357 US gal (12,710 l) per second, or 28,415 lb (12,890 kg). Each F-1 engine had more thrust than three Space Shuttle Main Engines combined.
So the methane-oxygen Raptor at 380 SI is almost as powerful as the hydrogen-oxygen Space Shuttle SLS at 453 SI, if I have the math right:
The tradeoff is reduced price and improved safety with the choice of methane over hydrogen, but with a reduced specific impulse.
Instead of doing that with all it's cost and safety implications, SpaceX intends to put 31 raptors on the bottom of their rocket. You could not fit the same thrust in other major US engines without making the rocket comically large.
Additionally, their control scheme was designed to shut down the engine opposite to any failed engine, meaning that they had very little engine-out margin in the inevitability of an engine failure.
All in all, we can (and have, with the Falcon Heavy) do much better.
N-1 could complete the mission with up to 4 engines turned off from the start. And I guess it it's not from the start, you could turn off even more engines and still have a successful flight.
Even though Saturn-V is heavier, N-1 has more lift-off thrust. Partially to have this kind of redundancy.
SpaceX's lots-of-engines approach has been well proven on the Falcon 9 thus far.
This is just to compare optimism for heavily multi-engined BFR with suggestions that N-1 was doomed just because of the number of engines.
So 330/275 = 20% better.
Yep. The actual physically relevant number is exhaust velocity, which is the mean velocity of the particles in the rocket exhaust.
Isp = Specific Impulse = exhaust velocity/(9.81ms^-2).
9.81ms^-2 there is not any actual acceleration, just an agreed conversion factor.
The unit of Isp has no physical significance, the convention just arose because early rocketry was a collaboration of people who all used different units, and they wanted to be able to compare numbers. Since the unit they all shared was a second, and they all agreed (roughly) on the real value of g, they just divided exhaust velocity by their g to get a number that could be compared. The convention stuck.
I'd look it up but my textbooks are in boxes, but the units everyone uses are wrong. Actually inappropriately simplified.
 I have one and exactly one textbook that has the full units for gc.
due to the exponential nature of the rocket equation, 100 tons of fuel will result in 2 tons to LEO in the former case and just 1 ton in the later.
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 about accidentally getting a little of the hot gas into the liquid propellant, because that would be mixing fuel with fuel-rich gas or oxygen with oxygen-rich gas. But in a normal staged combustion engine, you have to carefully seal off the oxygen from the fuel-rich gas (in a fuel-rich staged combustion engine) or the oxygen-rich gas from the fuel (in an oxygen-rich staged combustion engine.
You’re all over this thread being wrong and confused.
Source: am a rocket engine designer.
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?
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. May be even tertiary - how about possibility to optimize turbine frequencies for both propellant flows without losses on gears?
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.
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.
You say the main benefit is just increased pressure?
This guy has some simple short videos on rocketry. this one is on combustion cycles
Closed cycle is not a new idea: https://en.wikipedia.org/wiki/Staged_combustion_cycle#Histor...
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, lives on inside the design of such a new technology, the rocket.
it wasnt steam engine per.se., it was catalytic peroxide decomposition.
Are there? Here's a diagram of the RD-180, which uses an oxygen-rich preburner:
Where is anything escaping other than through the combustion chamber?
But full flow staged combustion lets you fully vaporize both propellants before they mix, leading to more optimal burning.
> In the 1950’s, Soviet scientists came up with something of a compromise. Instead of using a fuel-rich mixture in the preburner which produced an exhaust that couldn’t be safely recirculated into the engine, they experimented with running the preburner oxygen-rich.
> Eventually the Soviets mastered the metallurgy required to build the turbine and developed several engines that operated on the oxygen-rich preburner concept. The exhaust was piped into the engine’s combustion chamber and recovered at least some of the propellants which would otherwise have been dumped overboard. The modern day Russian RD-180 engine, currently in use by the American Atlas V, is a continuation of this technology.
From what I understand, the fuel and oxidizer that has been through the preburners to spin the turbines is basically fungible with fuel and oxidizer that is used in the combustion chamber.
If it's incompletely burned in the preburner, it gets completely burned in the combustion chamber.
And even if that were true that combustion happening in the preburner was wasted, then full-flow staged combustion would be equally bad because it still has preburners.
Mixing should be limited even if nothing special is done, due to the temperature and phase of matter and short timeframe. One could of course pay the weight penalty of a piston (need not have a perfect seal) or collapsing bag.
Doing a heat exchanger (to boil and thus pressurize) is another option, but then you're back to needing a place for the exhaust. It would let you do a sort of full-flow engine without turbopumps however, which is great. All those issues with cavitation and lubrication and stress cracking just go away.
This combines the worst of both worlds. You get the weight and complexity of turbopumps. And you get the weight, explosion risk and leakiness of pressure vessels.
Pressurizing has weight advantages. You can use a balloon tank. The tank no longer has to have the rigidity to support itself.
Rockets with typical cycles have used balloon tanks. Even modern ones like the Falcon 9 are partially that way, with just enough rigidity to be erected empty on the pad. The Falcon 9 uses helium to pressurize; that could be changed to preburner exhaust even if you did keep the moderate pressure and the turbopumps.
To drive the flow of fuel and oxidizer into the combustion chamber, you must overcome the pressure inside. To drive this flow, the turbopump output pressure will be of comparable magnitude to the combustion chamber. This pressure is in the realm of a few 10s of megapascals, or a few thousand PSI. This is a hard design requirement.
If you're proposing to "skip the turbopump", then necessarily you'd have to pressurize the entire fuel and oxidizer tanks to these high pressures. This is, to put it very mildly, utterly infeasible.
Pumps are still good because you really don't want the tanks to pressurize at the same pressure as the combustion chamber. If you do, you'll see a rapid disassembly, but your claim on it being unscheduled will be questioned.
IIRC, Block D taps some preburner exhaust for that...
> American engineers went in the opposite direction. They believed that a fuel-rich mixture in the preburner was possible and could be done with existing metal alloys, so long as hydrogen was used as the fuel instead of kerosene. This ultimately lead to the development of the Space Shuttle Main Engine, which to date remains the most efficient liquid fuel rocket engine ever flown.
SSME performance was due to H2 vs kerosene, it was not a full flow engine.
Edit; also no mention of Blue Origins BE-4 which is also a full flow engine.
BE-4 is not full flow, but oxidizer-rich (https://en.wikipedia.org/wiki/Staged_combustion_cycle#Oxidiz...).
Any small improvement in exhaust velocity of the engine makes a huge difference(sort of exponential) in the amount of payload it can take to orbit. In this case the 2 pre-burners also make relighting the engine in a vacuum a lot more reliable.
For more on the math:
Laws of physics forbidding it is understandable. We may likely never see the development of FTL because our understanding as we know it would see it impossible to do due to its strange implication about cause and effect.
And they used other words next to the unquoted impossible in the article itself like "deemed by scientists to be next to impossible". Context matters.
The part I'm confused about is how it talks about all of these 60s rockets that close the cycle but aren't as efficient because they mess up the mixture in the combustion chamber, but shouldn't the injectors be tuned for the excess fuel/oxidizer from the get go? It doesn't seem like you should be exhausting unburnt fuel/oxidizer if you have it tuned correctly.