I'm not a rocket scientist but from what I recall, designing a reliable high thrust large engine like the F1 from the Saturn V first stage is incredibly hard to get right. The Soviets could never do it which is why their launch vehicles all have lots of smaller engines.
There's an advantage to having many small engines. If you lose one, like what happened on that launch a few months ago, you can still keep trucking if your software can compensate.
If you go with a small number of rockets, it's unlikely you can sustain one (let alone) multiple engine failures. With several rockets arranged in a distributed geometrical fashion, you can significantly raise the probability of proper vehicle function even with several failures (as mentioned before, with the Falcon 9).
Redundancy is a funny issue, though. Replacing one engine with multiple engines means you can survive one failure, but you then have a multiple of the things that can fail.
It is also economy of scale when manufacturing. You need smaller tooling to make smaller engines, and because nobody makes rocket engines in quantities that would load the production line 24x7, almost only additional expense is raw materials which is negligibly small.
Also, gas dynamics in a smaller engine is simpler and easier to simulate.
And yes, it gives a lot of redundancy, if the only engine of the stage fails you are toast, if 1 out of 9 fails it's almost okay, as the most recent F-9 flight shown.
Having multiple engines means having multiple small containers that each take up metal, manufacturing, etc. All of which is dead weight as far as lift goes. By contrast a large engine is by proportion more fuel, which means you can get more stuff to orbit.
Increasing the number of engines can also decrease the reliability, for example if one of the engines explodes or catches on fire. The Falcon 9 has an armored tub around each engine to hopefully reduce the risk of a single engine problem becoming catastrophic.
Apart from the redundancy issue, size matters. The area/volume ratio of the combustion chamber necessarily goes down with size, which changes the flow dynamics. While a lower A/V ratio would seem to make the cooling problem easier to solve, I also remember people talking about the difficulty of getting stable combustion in the F-1 engines because they were so large.
From my amateur reading, I can say that a large number of engines leads to more failure modes due to interactions and complications, especially regarding vibration.
It's not really mentioned in the article but as a starting point they received all the work done for Fastrac [1], a NASA program to build small, cheap, expendable rockets - something like the Falcon 1. That got cancelled as NASA projects are wont to do but it made sense to use that design as a baseline. A clean sheet engine design is very complicated indeed - it's a very high dimensional, highly coupled optimisation problem, so a known good starting point is a very valuable thing to have.
Indeed the turbopump, which is one of the hardest bits of a rocket engine, is made by Barber Nichols, who made the turbopump for Fastrac and who make the turbopumps for the Merlin engines [2]. I've heard rocket engineers describe rocket engines as turbopumps with some extra plumbing. Perhaps a slight exaggeration but not far off, especially since their design is so tightly coupled and dependent upon overall engine parameters. The degree of coupling depends on the topology of the plumbing, or the 'cycle' of the engine, which I'll try and explain:
There are three kinds of rocket engine cycle (well, there are maybe more but these are the three that have been flown historically). The Expander Cycle, the Staged Combustion Cycle, and the Gas Generator cycle. I'll mention the last two.
Merlin, as the article mentions, is an example of a Gas Generator cycle [3]. In this cycle, you take off a little bit of fuel and oxidiser to burn outside the main combustion chamber, to generate some hot energetic gases that you can exhaust over a turbine. This spins the turbine up, which is connected to a shaft with a compressor on the other end. The compressor increases the pressure of the propellents so that they can be injected into the main combustion chamber. This assembly (turbine, shaft, compressor) is called the turbopump. It's necessary because the engines require very high flow rates to get the thrust they need, and that has to be at a high pressure - higher than the pressure of the combusting gases inside the combustion chamber, else you wouldn't be able to inject it!
Back to the bleed-off to drive the turbine. You usually don't want a perfect stoichiometric mix of fuel and oxidiser for this, or even close, because it generates extraordinary hot gases that no turbine would last long in (The turbines are spinning at many tens of thousands of RPM usually so would be subject to much higher forces than the actively cooled walls of the main combustion chamber). For this reason you usually have a large imbalance of one propellent to the other to keep the temperature down. Usually you run with excess fuel, or 'fuel-rich', as the opposite - oxidiser rich - means you have hot oxidising gases which are harder on the metallurgy. I do know of some russian exceptions to this, though, where fuel rich would have left sooty deposits in the plumbing (The materials science employed in the turbines was apparently so witchcraft that when the US got intelligence of oxidiser-rich turbine precombustors, they thought is was deliberate counterintelligence from the russians to get them to waste billions researching the impossible). The gas generator cycle, as the article mentions, dumps this turbine exhaust overboard separately. The problem with this is that there's a load of uncombusted fuel in this exhaust, which you're just wasting, and this hits your rocket performance - the Specific Impulse ( I_{sp} ), as you're not getting as much bang out of a given mass of fuel as you could.
The answer to this is the Staged Combustion Cycle [4], where you also inject the exhaust of the turbine into the combustion chamber to finish off combustion. The performance of these engines is higher but the thermodynamic balance to design a working system is a greater challenge, and some of the engineering is a bit harder too. Staged Combustion engines are mostly russian, although the Space Shuttle Main Engines are a US-design example of Staged combustion.
SpaceX have been gradually and incrementally improving the Merlin's away from their simpler beginnings, and it's been pleasin...
Thank you for the analysis. I opened the comments expecting at least one with something more in-depth than the article itself, and wasn't disappointed.
Actually, SpaceX manufactures the M1D turbo pumps themselves. If you get a tour of the factory, you can look into the clean room where they are assembling them.
What would you call a turbopump that uses a secondary fuel not used by the rest of the engine? The H2O2/steam powered turbopumps in A4/V2 comes to mind. Would that be another sort of Gas Generator cycle?
From "Equinox - The Engines that came in from the Cold" about the secret soviet moon program, i linked to the staged engine they built, and that they got working.
http://youtu.be/BLg1QUq5GQM?t=14m21s
> For this reason you usually have a large imbalance of one propellent to the other to keep the temperature down.
So how large is that imbalance usually, or in the case of the Merlin 1C?
> although the Space Shuttle Main Engines are a US-design example of Staged combustion.
Is this, why their exhaust is so clean?
How much more efficient is this engine compared to the Merlin?
PS:
I'm excited, this engine will be used again in the new super heavy lift [1] that NASA is developing, though obviously, I can't tell if its a good idea.
>SpaceX's initial plan will be to build a lox/methane rocket for a future upper stage codenamed Raptor. The design of this engine would be a departure from the "open cycle" gas generator system and lox/kerosene propellants that the current Merlin 1 engine series uses. Instead, the new rocket engine would burn lox/methane in a much more efficient "staged combustion" cycle that many Russian rocket engines use.
21 comments
[ 5.1 ms ] story [ 83.1 ms ] threadThere's an advantage to having many small engines. If you lose one, like what happened on that launch a few months ago, you can still keep trucking if your software can compensate.
If you go with a small number of rockets, it's unlikely you can sustain one (let alone) multiple engine failures. With several rockets arranged in a distributed geometrical fashion, you can significantly raise the probability of proper vehicle function even with several failures (as mentioned before, with the Falcon 9).
Also, gas dynamics in a smaller engine is simpler and easier to simulate.
And yes, it gives a lot of redundancy, if the only engine of the stage fails you are toast, if 1 out of 9 fails it's almost okay, as the most recent F-9 flight shown.
Having multiple engines means having multiple small containers that each take up metal, manufacturing, etc. All of which is dead weight as far as lift goes. By contrast a large engine is by proportion more fuel, which means you can get more stuff to orbit.
Indeed the turbopump, which is one of the hardest bits of a rocket engine, is made by Barber Nichols, who made the turbopump for Fastrac and who make the turbopumps for the Merlin engines [2]. I've heard rocket engineers describe rocket engines as turbopumps with some extra plumbing. Perhaps a slight exaggeration but not far off, especially since their design is so tightly coupled and dependent upon overall engine parameters. The degree of coupling depends on the topology of the plumbing, or the 'cycle' of the engine, which I'll try and explain:
There are three kinds of rocket engine cycle (well, there are maybe more but these are the three that have been flown historically). The Expander Cycle, the Staged Combustion Cycle, and the Gas Generator cycle. I'll mention the last two.
Merlin, as the article mentions, is an example of a Gas Generator cycle [3]. In this cycle, you take off a little bit of fuel and oxidiser to burn outside the main combustion chamber, to generate some hot energetic gases that you can exhaust over a turbine. This spins the turbine up, which is connected to a shaft with a compressor on the other end. The compressor increases the pressure of the propellents so that they can be injected into the main combustion chamber. This assembly (turbine, shaft, compressor) is called the turbopump. It's necessary because the engines require very high flow rates to get the thrust they need, and that has to be at a high pressure - higher than the pressure of the combusting gases inside the combustion chamber, else you wouldn't be able to inject it!
Back to the bleed-off to drive the turbine. You usually don't want a perfect stoichiometric mix of fuel and oxidiser for this, or even close, because it generates extraordinary hot gases that no turbine would last long in (The turbines are spinning at many tens of thousands of RPM usually so would be subject to much higher forces than the actively cooled walls of the main combustion chamber). For this reason you usually have a large imbalance of one propellent to the other to keep the temperature down. Usually you run with excess fuel, or 'fuel-rich', as the opposite - oxidiser rich - means you have hot oxidising gases which are harder on the metallurgy. I do know of some russian exceptions to this, though, where fuel rich would have left sooty deposits in the plumbing (The materials science employed in the turbines was apparently so witchcraft that when the US got intelligence of oxidiser-rich turbine precombustors, they thought is was deliberate counterintelligence from the russians to get them to waste billions researching the impossible). The gas generator cycle, as the article mentions, dumps this turbine exhaust overboard separately. The problem with this is that there's a load of uncombusted fuel in this exhaust, which you're just wasting, and this hits your rocket performance - the Specific Impulse ( I_{sp} ), as you're not getting as much bang out of a given mass of fuel as you could.
The answer to this is the Staged Combustion Cycle [4], where you also inject the exhaust of the turbine into the combustion chamber to finish off combustion. The performance of these engines is higher but the thermodynamic balance to design a working system is a greater challenge, and some of the engineering is a bit harder too. Staged Combustion engines are mostly russian, although the Space Shuttle Main Engines are a US-design example of Staged combustion.
SpaceX have been gradually and incrementally improving the Merlin's away from their simpler beginnings, and it's been pleasin...
> For this reason you usually have a large imbalance of one propellent to the other to keep the temperature down.
So how large is that imbalance usually, or in the case of the Merlin 1C?
> although the Space Shuttle Main Engines are a US-design example of Staged combustion.
Is this, why their exhaust is so clean? How much more efficient is this engine compared to the Merlin?
PS: I'm excited, this engine will be used again in the new super heavy lift [1] that NASA is developing, though obviously, I can't tell if its a good idea.
[1] http://en.wikipedia.org/wiki/Space_Launch_System#Core_stage
The Space Shuttle Main Engines used LOX and LH2, which creates water when burned.
>How much more efficient is this engine compared to the Merlin?
The specific impulse is the metric to look at, and ballooney already talked about it.
Not much of a bet – Musk talked about this two months ago.
http://www.flightglobal.com/blogs/hyperbola/2012/11/musk-goe...
>SpaceX's initial plan will be to build a lox/methane rocket for a future upper stage codenamed Raptor. The design of this engine would be a departure from the "open cycle" gas generator system and lox/kerosene propellants that the current Merlin 1 engine series uses. Instead, the new rocket engine would burn lox/methane in a much more efficient "staged combustion" cycle that many Russian rocket engines use.
http://www.livescribe.com/cgi-bin/WebObjects/LDApp.woa/wa/ML...