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?
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
https://www.nextplatform.com/2015/03/27/rockets-shake-and-ra...
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
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[1] are in boxes, but the units everyone uses are wrong. Actually inappropriately simplified.
[1] 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.
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,[4] or 28,415 lb (12,890 kg). Each F-1 engine had more thrust than three Space Shuttle Main Engines combined.[5]
https://en.wikipedia.org/wiki/Rocketdyne_F-1#Design
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:
https://en.wikipedia.org/wiki/Raptor_(rocket_engine_family)#...
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.
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.
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.
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