This thing I can not comprehend about rocket engines is how the turbopump manages to hold together.
A turbine blade in the SSME about the size of your thumb makes 600 horsepower.
A turbine blade in the SSME about the size of your thumb makes 600 horsepower.
Edit: the SSME high pressure fuel pump turbine produces 63000 hp (46 MW). There's also one for the oxygen, and a pair of low pressure pumps as well. Crazy...
Edit edit: the fuel pump transfers 155 lb/sec of liquid hydrogen. If fully combusted (142 MJ/kg), that would release 10.0 GW of heat per engine.
Whereas "all the power stations in the UK" keep on trucking, day in, day out, year after year.
And gas generators using main propellants are better, certainly, but less than you might think because they have to haul all their oxidizer with them (unlike aircraft) and are also usually run very far from stoichiometric (maybe just 0.3 O:F ratio compared to a stoichiometric 3.4) to keep the temperature down. So unless you have a pretty high temperature turbine, you might not beat peroxide by much!
So the easiest gas generators have worse energy density (keep in mind RocketLab does stage off batteries if necessary…), and the next easiest, while better, aren’t MASSIVELY better without careful efficiency improvements. The real efficiencies come when you use like an expander cycle or a staged combustion cycle or you feed the gas generator exhaust back into the nozzle like Merlin Vacuum or F-1. And those are all much more complicated. A level of complication that is not worth it for small rockets but is for larger.
So it’s really not about electric scaling poorly (electric scales just fine) but about the greater complexity of better engine cycles being worth it at larger scales.
If I remember correctly, with F-1 the gas generator exhaust was sent to the nozzle to cool the nozzle, not to add efficiency to the engine main cycle.
We're still trying to get more efficiency from isochoric combustion, but the expected wins aren't too big. It's good that full-flow combustion becomes more of a norm.
People don't grasp what an unbelievably complex engineering problem that is. It's at least an addition of difficulty at the same magnitude as building a steady-state 63000 hp turbine pump in the first place.
Mechanically implementing it in the inherently steady-state design rocket of most rocket cycles. Having variable controls able to work at those pressures. Testing structural dynamics for a range of harmonic conditions instead of one. And do all of that with materials that need to tolerate temperatures going from cryogenic to white hot, without allowing thermal expansion to affect the mechanical tolerances of parts running at thousands of RPMs. And now you have varying flow rates and negative pressures in the lines coming from the external tanks, so have to design such that cryogenic liquids (that normally would require immense positive pressure to keep liquid) don't spontaneously boil or cavitate or cause a shock-like wave (think water hammer turning off your bath faucet) under changing negative pressures.
It's really difficult even for seasoned engineers to grasp the scale of difficulty involved.
On the one hand, you have https://xkcd.com/793/ responses:
> You're trying to predict the behavior of <complicated system>? Just model it as a <simple object>, and then add some secondary terms to account for <complications I just thought of>. Easy, right? So, why does <your field> need a whole journal, anyway?
Just pressurize the tanks, and meter the flow with some valves. Easy, right?
They move to the other side of the distribution after a little more thought, when they they realize it's simply infeasible to put thousands of horsepower in a pump that size, and declare the whole endeavor completely nonsensical and impossible.
Ran through this on a recent project involving an automated sewing machine. At first, it seems ludicrous that you could tie knots thousands of times per second. Oh wait, it's a single motor and old cam-driven tech from the 1800s, available off the shelf for a couple hundred dollars?
Its rocket science
Don't forget the foundations!
Turbopumps are finicky, but at least heat and cryogenics are separated with the turbine shaft. In combustion chambers, you have both mechanical stress from high pressures (and thin walled constructions) and thermal stress - difference of a few thousands K, so chambers aren't that simple either.
Still - now we have enough knowledge to repeatedly design flyable rockets from scratch, different teams, periods of time, countries. A lot of work still remains - and the plumbing is a good manifestation for that.
The engineering of the seals that can operate at those speeds and pressures is a whole specialised field. If I remember correctly, the SSME design uses a labyrinth seal pressurised with Helium.
In other words, the various gases are kept separate only by more gas.
Similarly, the Saturn V main engines were unlubricated because no lubricant could be found that could tolerate the extreme conditions. They were just designed to wear out slow enough to keep operating over their operating lifetime, which was measured in hundreds of seconds (including static test firing).
What do today's reusable engines use?
"Liquid oxygen is [the high pressure oxidizer turbopump's] only lubricant and a poor one at that."
So it's some lubricant and I imagine also significantly a coolant.
This https://ntrs.nasa.gov/citations/20100023061 goes into a lot more detail. But exotic materials and designs bathed in LOX which provides cooling (and minimal lubrication) seems to be how the space shuttle engines achieved their long bearing lifetimes.
Edit: They used carbon seals. The turbopump uses ball bearings cooled by the fuel. Detailed description here: http://web.archive.org/web/20150509173846/http://agentdc.uah...
Those pressures, rates of fluid flow, and shaft rpms is going to result in the mechanical surfaces being separated by hydrodynamic forces and/or boundary layer flow.
Technically, lubrication is the same physical principles, but viscosity, weight, chemical stability/ durability and other factors become more important for lubricants that are recirculating.
(Oil in eg, a 4-stroke engine operates on these principles, with the rotating shaft causing pressure differences in the oil that "lift" the shaft from contact with the surrounding metal, until pressure (and therefore spacing) is roughly equal on all sides.
A ping-pong ball floating on a column of air self-stabilizes in the same way.)
The primary concern for these turbopumps becomes heat. So heat transfer and temperature of the "lubricating" fluid become more important than its other nominative qualities as a lubricant. Fluids being pumped at cryogenic temperatures can obviously help here.
So instead of a normal "lubricant", most of these turbopump designs just run a portion of the fuel/oxidizer fluid through the critical areas to provide the surface separation and cooling required.
Ball bearings require "lubrication", and the lack of good lubricants is exactly why they use fluids that end up behaving as I described, and why "lubricant" is the wrong way to think of it in this case.
If anything, the existence of a good lubricant would prevent the surface wear. Temp control and fluid dynamics are what they have to work with, and yes it results in surface wear because it's suboptimal.
But so does running your 4-stroke too hot.
To give a far less glamourous example, I am a mech eng who works in rail. We needed some more orecars to complement an existing fleet that only had about 10 years' operational life left in them. Thus, they wanted me to design for a shorter life than the 25-30 year standard that we target, to save cost. However, trying to thin out the structure so it only had 10-15 years' of fatigue life in it meant that it fell well short of the proof load requirements needed to stop it ripping in half in a worst-case shock load. Put differently, the constraint around peak loads effectively baked about 20 years of operational life into the structure, and in turn made it difficult to save money on a shorter-life design.
It would not surprise me if many of the non-ablative parts in a rocket are in fact fairly durable without the stop-start cycles. So whilst a launch may only take 5-8 minutes, I could totally believe that a 15-20 minute launch wouldn't demand heaps more from the parts.
Of course, this absolutely doesn't translate to ablative parts, or items that undergo stop-start cycles. The latter of course is where the devil is for reusable equipment. Depending on the failure modes in question for turbo pumps and throttling controls, those may or may not apply.
A .22 LR rifle bullet might acquire 200 J in 2 ms, which means the firing gun is producing 100 kW mechanical, plus probably another 300 kW thermal. So another way of thinking of this is that an engine dissipating 10 GW is equivalent to something like 25000 handguns firing at once, without ever stopping.
That's not the power output of the rocket (which someone quoted below). That's the power required to move the fuel and oxidizer into the rocket engines. Pumping the same amount of liquid can move a 100,000 ton floating city across the ocean at 40mph.
As I understand it, a huge factor in the power requirement of the pumps is raising the pressure to the required level which is very high (combustion chamber pressure might be hundreds of atmospheres).
Designing for thermal cycles and serviceability[0] is at least as difficult a problem as running a hypothetical rocket engine an equal amount of time in one longer, hotter burn.
(Such a design isn't needed and wouldn't be practical, but then again multiple aspects of SSMEs being reusable turned out not very practical either, depending on what version of design criterea you evaluate and how the expected vs actual usage changed over the lifetime of the program.)
[0]In both the engineering sense, as durability of the various loading cycles (ie lifetime turbine rotations or number of thermal cycles before eol or failure), and as being constructed as able to undergo maintenance and refurbishment between launches.
Not as much as RL-10s. SSME you can disassemble - because you should do that, as thermal stresses on turbine blades are too dangerous, so you have to periodically replace the parts which are nearing the fault.
edit: if you watch gas generator tests on youtube the gas coming out is dark because it's very fuel rich which keeps it cool (in a relative sense)
Russians had oxydizer-rich gas generators in large engines since e.g. early 1960-s (see Proton 1st stage engines). Oxygen-rich gas generators are at least since mid-1980-s (RD-170). So not exactly never done before.
A jet engine can't have such power density because it uses gaseous air which is about 1000 times less dense than liquid oxygen and only contains 20% oxygen.
Everything follows from that.
The pumps are not challenging temperature wise since they pump cold liquids.
The turbine is challenging, but the temperature can be limited by varying the ratio of propellants in the preburner (very lean or very rich means lower temperature). If you use lean, then it's a very oxidizing environment. If you go very rich, there's soot (if you use fuels with carbon).
And the chamber is not so challenging because there is so much cool liquid available for cooling.
You can boil water with a candle and a paper cup.
A high performance jet engine is a harder problem than a medium performance rocket engine.
Considering all the manufacturing, precision manufacturing involved the V2 effort, it's astounding they could do it at all.
A gear tooth the size of your thumb pulls a semi truck up a mountain.
Agree that these are rather different stress modes, just wanted to point to another place in the engine with pretty extreme stresses.
Another interesting thing to think about, it's the turbopump keeping the combustion from running back up the injectors. So, the turbopump has to outperform the combustion chamber in terms of pressure and flow rate.
Not quite literally, in e.g. Soviet engine designs some gas generators are oxygen-rich, they produce oxidized exhaust of high temperature, and the engine still doesn't burn (usually).
> So, the turbopump has to outperform the combustion chamber in terms of pressure and flow rate.
Yes, the injector pressure drop can't be too small.
The most similar technology is a boiler feed pump (used to feed the boilers in a steam turbine system, e.g. in a ship or power station) - these work in a similar way, high pressure steam is bled off from the system to drive a small turbine, which drives a high pressure pump to feed the boiler, which feeds the 'main' steam turbine. Similar pressure (>100 bar), but the turbine on the rocket engine needs to handle much higher temperatures (hot gas rather than steam).
The fuel rail in a modern diesel engine is operating at 25-30,000 psi all the time.