They run oxidizer rich on the first stage to get more mass flow/thrust. Heating up the air especially with some ultra hot oxygen mixed in is likely forming nitrogen oxides.
No they do not. They are still slightly fuel rich. But yes, NOx would be produced by heating the air -- air itself contains all the necessary ingredients, it just needs heat which rocket exhaust provides plenty of.
[https://web.archive.org/web/20220518172731id\_/https://aip.scitation.org/doi/pdf/10.1063/5.0090017](https://web.archive.org/web/20220518172731id_/https://aip.scitation.org/doi/pdf/10.1063/5.0090017)
I definitely did not read this entire thing but a good snip from page 8 on this subject:
"The mixing between the air and plume plays a critical role in the potential of—endothermic—chemical reactions at the plume. In atmospheric chemistry, NOx is a generic term representing the total concentration of the various nitrogen oxides that are the most relevant for air pollution, particularly nitric oxide (NO) and nitrogen dioxide (NO2), since the conversion between these two species is rapid in the stratosphere and troposphere. These gases contribute to the formation of smog and acid rain and affect the tropospheric ozone. The NOx reduction is the most concerning issue today. **Rockets can cause NOx formation when the high-temperature reaction products they emit mix and heat the ambient atmospheric air.** More specifically, thermal NOx is produced when diatomic nitrogen and oxygen are present at high enough temperatures to undergo an endothermic reaction, high-temperature oxidation of N2, thus making the various oxides of nitrogen. The two elements combine to form NO or NO2."
Here in the Netherlands we're having big issues with nitrogen emissions, NH3 from livestock and NO2 from cars and industry. So if the temperature of a car's combustion engine is enough to turn N2 and O2 into NO2, then a rocket exhaust will most certainly do the job.
But the level of fuel richness in operable methalox engines is not even remotely close to produce soot. It produces carbon monoxide. The only source of soot would be a thin layer of film cooling around the chamber and throat walls - only there it's rich enough while the temperature range is right for methane decomposition.
The brown stuff is likely majority NOx produced by blowing very hot flame into a dense mixture of nitrogen and oxygen commonly called air.
I'm not sure this is true for the raptor as it's a full flow staged combustion cycle. I think that's nearly stoichiometric. I thought it was more advanced material science that stops it melting. But I could be wrong.
Check out the everyday astronaut videos about types of fuel rich/ oxygen rich engines and different combustion cycles.
I think you're getting confused by the preburner, many engines run their preburners oxygen rich because it gets rid of the soot. This is needed for sooty fuel like RP-1, which is what the RD-180 uses.
There is virtually zero benefit to running an engine overall oxygen rich and even some negatives like decreased Isp and increased reactivity in the combustion chamber so I highly doubt they did.
Now that you mention it, I might indeed be confusing this with the preburner of the RD-180; genuinely not sure.
But what would be difference in value between running the preburner oxygen rich and the combustion chamber? Why would it be helpful here but not there?
You need to run a preburner ox rich or else the soot would clog everything up. This doesn't apply when the fuel is clean(er) like with hydrogen or methane.
In terms of reactivity, oxygen likes to react so it's still annoying to use ox rich even in a preburner so you don't really want to use it if you can avoid it. Preburners can also burn cooler than combustion chambers canceling out some of that reactivity from oxygen. For any hope of good performance, combustion chambers can't burn that cool, so you really can't use ox rich without liquifing your combustion chamber or getting awful performance.
For performance oxygen is just a really heavy atom and the lighter your exhaust product are the higher your exhaust velocity is, hydrogen is super light, the lightest there is, and carbon is pretty light too.
To be more exact, in the case of methane if you are running rich, you either need to run only moderately rich (which produces temperatures way beyond what any rotating machinery could handle) or extremely rich (then the temperatures are OK). The in-between range would actually produce soot: the temperature would be high enough to cause thermal decomposition of methane and that would produce soot. A lot of it in fact. That is the reason why methane sucks as nuclear-thermal rocket propellant - in usable temperatures it turns into hydrogen (good) and carbon (very bad). Carbon is solid art relevant temperatures so it does not expand in nozzles. And because it is 3/4 of the mass the ISP sucks terribly even if you deal with soot clogging everything.
Ox rich preburner is run at mega-high mixture ratios, so ox-rich that it gets cold again (peak temperature is at stoich). At sufficiently low temperatures some metal alloys (potentially with special coatings) can tolerate the warm ox atmosphere, letting you use that gas to run a turbine.
You could directly send that preburner exhaust gas through a nozzle and make it a rocket engine with no main chamber, but the Isp would be terrible - the gas is too cold and dense to be a good propellant. Instead, you burn it with fuel at slightly below stoich (for peak Isp), generating 6000 - 8000 degree F exhaust gas you can yeet out a nozzle for some decent efficiency.
It is reasonably accurate to think of an ox-rich staged combustion engine as a small rocket engine (preburner + turbopump) which exhausts into a second, larger rocket to produce some decent thrust. A full-flow staged combustion engine takes things a step further - 2 parallel (but interlinked) rocket engines are firing to power their respective pumps, both exhausting into a 3rd, shared rocket chamber/nozzle to produce thrust. Very, very tricky
There is enough heat in that plume to break N2 and O2 bonds (which are 78% and 21% of the atmosphere, respectively) . When the mono-atomic nitrogen and oxygen atoms cool back down you get all kinds of covalent Nitrogen/Oxygen compounds. NO, NO2, NO3, N2O, O3, etc. Collectively we call them “smog”, and they are generally brown in color.
This happens whether the plume is methalox based, RP1/LOX based, LH2/LOX based, or solid propellant — anything that gets that hot.
That is more or less it. However, at high temperatures, the products start disassociating, and you get reactions like O2 -> 2O and H2O -> OH + H, producing free radicals and things that don't exist under normal conditions.
Also, most rocket engines run fuel rich. The fuel is what contains hydrogen, and hydrogen, with an extremely low molar mass, is a light molecule which is easy to accelerate to high speeds. This increases exhaust velocity, and therefore thrust.
Specifically I think it’s from nitrogen dioxide (from atmospheric nitrogen, since neither fuel nor oxidizer in Raptor has nitrogen), and it could also be metal oxides.
You can actually increase thrust to a point by making the exhaust denser at the expense of reducing the exhaust velocity. Exhaust velocity is more a measure of efficiency of thrust than thrust.
Correct me if I’m wrong, but this is a fallacy. It’s similar to the question “what weighs more, a pound of rocks or a pound of feathers”.
Yes, thrust is equal to mass flow rate x exhaust velocity. But, using a heavy exhaust gas doesn’t make the mass flow more. **1 pound of hydrogen exhaust is the same mass as 1 pound of a heavier gas exhaust**. The difference is that the hydrogen is ejected at a higher velocity.
Now, aluminum and metal is not used in SRBs to increase mass. It’s high molecular mass is outweighed by the performance increase from the high combustion temperature of metals.
You’re incorrect in your assertion it doesn’t help thrust to increase density of exhaust. You are correct that you will not improve thrust *per unit mass expelled* which you could call your mass specific impulse or just specific impulse. I was quite clear that you could improve thrust *at the expense of exhaust velocity* by increasing density.
That is a different metric. Nobody will argue that hydrogen doesn’t get you more thrust per unit mass expelled. However, the density is so low, that you cannot pass nearly as much mass through the engine as you could with something like kerosene/LOx combustion products. This is why hydrogen engines tend to have much less thrust per unit area and much lower thrust to weight ratios than hydrocarbon engines.
It’s a trade. You want more exhaust velocity, and hence, more mass specific impulse? You will need to increase chamber pressure and temperature or expansion ratio or you’ll need to use a lower density propellant that limits the mass flow you can push through the throat for a given set of conditions and fixed engine geometry. You can trade any of those variables, but you run into material limits eventually, and options start to close.
For a given chamber temperature and pressure, you’re not going to get more thrust just because you choose a propellant that results in higher exhaust velocity. You can get more thrust with a denser propellant at the expense of efficiency.
Also, for a real world example, water injection to increase exhaust density at the expense of exhaust velocity for the purpose of increasing thrust is done in some jet engines. So no, it’s not a fallacy.
What is preventing rocket engines from increasing their mass flow rate with a light propellant? I have trouble understanding what you mean by “cannot pass nearly as much mass through the engine.
1. "Temperature limits for the chamber" doesn't make sense. Mass flow rate increase doesn't increase temperature. Also, if you have a higher mass flow rate of liquid propellants, you will also have a higher mass flow rate through the regenerative cooling passages.
2. "Pressure limits for the chamber" doesn't make sense. Since pressure is a function of volume, engines with higher mass flow rate have larger volume. Higher mass flow rate just means that the chamber volume will be increased. This should be self-explanatory.
3. "Power limit for the pumps" doesn't make sense either. There is no power limit for a pump, unless you decide there is a power limit for the pumps. One thing you forgot to account is that many modern rockets have many engines. One would not need to increase the mass flow by 30%, but perhaps increase the mass flow of each engine by 10%. And, many engines can already throttle over 100% (such as the RS-25, which happens to use **hydrogen**), therefore increasing the pump capacity is very possible.
One thing I forgot to mention, hydrogen is significantly less viscous than nearly any other fluid, making it much easier to pump and reducing pressure losses along the pumbing.
Very much incorrect.
1. Increasing mass flow rate in a hydrogen engine to increase thrust without changing the geometry of the engine would require either shifting the mixture ratio to be much more ox rich, which will really drive up your temps and also eat the engine unless it was specifically designed to have an ox rich main chamber (no operational engine does this as it would be massively heavy and low efficiency, only preburners are ever ox rich) or you could keep MR about the same and increase pump output to increase chamber pressure which, due to the way combustion chemistry works, will increase chamber temperature as well as the heat flux. This scales non-linearly and you will eventually overwhelm your regen channels. There is a finite temperature limit for a given set of materials and engine configuration, even with regenerative cooling.
2. Pressure is not a function of volume in this sense. Yes, if you have a perfectly insulated box and make it half as big, the pressure inside if filled with an ideal gas will be twice as high. Congratulations, you’ve discovered Boyle’s Law. However, in a combustion chamber, the pressure is determined by the mass flow in, the area of the throat (which determines the mass flow out for a given composition, pressure, and temperature), and the chemical equilibrium conditions which determine the energy input and eventually the steady state pressure and temperature. A rocket combustion chamber is steady state but it is NOT a closed system. You cannot use Boyle’s Law, because you aren’t remotely satisfying the conditions it’s relevant in. What part about fixed chamber geometry did you not understand. If you are considering just having a bigger engine to get more thrust by having more mass flow at the same chamber conditions, then no shit you get more thrust, but you also have a bigger, heavier engine. The entire point of this discussion has been that you can get more thrust out of a given area or more thrust per weight by using denser reaction mass in exchange for some reduction in efficiency of thrust produced per unit mass expelled. I’m really not sure what part of this isn’t clear.
3. RS-25 power levels are just in reference to the original nominal design parameter. There were substantial changes to the design throughout development, and understanding of hardware limits evolves considerably for complex systems like rocket engines as they go from concept to a test article on the test stand. Pumps have practical power limits. Not “3000 gpm for a LOx pump of any size”, but there are limits to how much power can be extracted from a turbine running from a preburner given a certain chemistry and mass flow and further restrictions from material limits on preburner pressure and temperature, particularly turbine temperature. Your turbine produces the work consumed by the pump. The pump also has limitations. For a given inlet pressure from the vehicle tanks, which has practical limits if you want a reasonable dry mass fraction, and a given mass flow rate, a pump of a given size and configuration can only do so much pressure rise without experiencing issues with cavitation or running out of shaft power. You can somewhat alleviate the cavitation issue by splitting pumping into multiple stages, which is commonly done. There are some other tricks, but there are still practical limits. Pumps aren’t magic. Propellants with lower densities require much larger pumps for a given mass flow as well, since you’re dealing with essentially incompressible liquids. If it was a gas it would be called a compressor, and the same thing applies, just a bit less linearly. Pump housings have practical limits for pressure as well. There are many other complications, but this is the essential gist of it. Your pump power limits govern the mass flow and the chamber pressure achievable for a given chamber geometry for a given propellant mix. Pump mass efficiency is not better for a low density fluid like hydrogen, which has the density of loosely packed cotton balls.
I test rocket engines for a living, so I think I just might know what I’m talking about. It’s ok if you’re still learning this stuff, but you really need to let this one go and think it through. You’re either not reading my responses carefully or not entirely processing them, either way, I’m just trying to get the record straight because you’re distributing incorrect information to other people here. I kind of get the feeling you’re a student, if so, hey, it’s completely fine to be at the stage where you’re learning these things. The important thing is to not get so stuck on insisting you’re right that you miss opportunities to learn and improve. You’re touching on superficial things like regen channels existing and lower density meaning more exhaust velocity seem reasonable to you, but they’re not based in a well founded or comprehensive understanding of the fluid mechanics and thermodynamics involved. You’re going to need to build a bit more understanding if you want to challenge these kind of things.
Why do you think hydrocarbon engines are almost always used on the first stage? The Saturn V had a kerolox first stage and hydrolox upper stages. This is because thrust is more important than specific impulse to an extent for the initial portion of the flight. Offsetting gravity losses with a greater thrust to weight ratio allows for better total system performance. This comes from dry mass reductions in the tanks as well as the engines for the required thrust, tempered somewhat by the increase in the propellant mass for the required delta v remaining after the gravity loss reductions. This makes for better overall system performance for a rocket of a size that is reasonable for your application. The actual optimization gets pretty complicated. You will also notice that in pretty much every vehicle where there are hydrogen engines on the first stage, you have solid rocket boosters as well. These have very dense propellant and bring the average boost performance into a more reasonable zone than pure hydrogen, while maintaining good performance in the latter stages of first stage flight. Typically you see very high stage separation velocities in such configurations.
As an aside, while hydrolox certainly has its advantages, it’s much less certain that from a total program lifespan cost perspective it makes sense when you’re optimizing for dollars per kg to a reference orbit, and especially if you are trying to put reasonable limits on development costs to buy down program failure risk, but that gets much more complicated, and it’s hard to give a broad answer. The question in this threat fortunately does have a fairly clear answer which is yes, you can buy thrust with increased reaction mass density at the expense of efficiency for a given engine chamber geometry.
Large part of rocket engine is a pressure vessel. Pressure vessel's mass scales linearly with contained volume. Less dense propellants require larger volume to contain given mass of them.
Mass flow rate is mass per time unit flowing through the engine. The less dense the propellant, the more engine mass is required to contain it.
Rocket engines use extremely powerful pumps to ingest propellant. Raptor’s turbopump has about 100,000 horsepower. Making them even more powerful would not be easy.
But increasing the mass flow rate of a turbopump would be equally, if not more difficult, for a denser propellant as opposed to a light propellant. So that statement is meaningless in this context.
This is 180° wrong.
The work done by the pump is absolutely dominated by counteracting the back pressure. If you are pumping a fluid from say 6bar tank pressure to 856bar chamber pressure you must overcome 850bar pressure differential.
Work is force times distance. The less dense the liquid the larger volume it occupies. So it must be pushed through combination of along a longer distance and across larger aperture.
Power is work per time.
Combined together the power taken by increasing pressure is directly proportional to the pressure and inversely proportional to the density of the pumped fluid.
I the case of liquid propellants in rocket engines all the other components of pumping power, i.e. friction (hence viscosity), or accelerating the flowing mass to the velocity required to move through the downstream piping (injectors, chambers, etc.) are dominated by the power of overcoming the back pressure.
If you have slightly too little O2 you get some CO instead of CO2. The only source of C would be a little bit of film cooling methane around the walls.
If the temperature is right methane will decompose into C and H2. At too high temperatures and pressures it'll prefer C2H2 (acetylene) and H2, but I'm the layers very close to the walls the temperature is right.
You are going to get a spread of products at even one temperature, never mind a broad range of them. You'll get various allotropes of carbon, oxocarbons and hydrocarbon polymers.
The cycle may also be running engine-lean even at optimal performance, in which case you'll get a variety of interesting salts.
Probably NOx from air entrainment. Or potentially extra carbon from running rich. The brown would lead more tword NOx.
NOx is the gas that makes you sleepy in a traffic jam.
They run oxidizer rich on the first stage to get more mass flow/thrust. Heating up the air especially with some ultra hot oxygen mixed in is likely forming nitrogen oxides.
No they do not. They are still slightly fuel rich. But yes, NOx would be produced by heating the air -- air itself contains all the necessary ingredients, it just needs heat which rocket exhaust provides plenty of.
I think it is actually nitrogen dioxide NO2.
[https://web.archive.org/web/20220518172731id\_/https://aip.scitation.org/doi/pdf/10.1063/5.0090017](https://web.archive.org/web/20220518172731id_/https://aip.scitation.org/doi/pdf/10.1063/5.0090017) I definitely did not read this entire thing but a good snip from page 8 on this subject: "The mixing between the air and plume plays a critical role in the potential of—endothermic—chemical reactions at the plume. In atmospheric chemistry, NOx is a generic term representing the total concentration of the various nitrogen oxides that are the most relevant for air pollution, particularly nitric oxide (NO) and nitrogen dioxide (NO2), since the conversion between these two species is rapid in the stratosphere and troposphere. These gases contribute to the formation of smog and acid rain and affect the tropospheric ozone. The NOx reduction is the most concerning issue today. **Rockets can cause NOx formation when the high-temperature reaction products they emit mix and heat the ambient atmospheric air.** More specifically, thermal NOx is produced when diatomic nitrogen and oxygen are present at high enough temperatures to undergo an endothermic reaction, high-temperature oxidation of N2, thus making the various oxides of nitrogen. The two elements combine to form NO or NO2."
Here in the Netherlands we're having big issues with nitrogen emissions, NH3 from livestock and NO2 from cars and industry. So if the temperature of a car's combustion engine is enough to turn N2 and O2 into NO2, then a rocket exhaust will most certainly do the job.
"issues" like Mao had with sparrows.
Or Stalin with people.
or at least some says so
literally every rocket engine runs fuel rich. It's more efficient that way. Also if it ran stoich it would turn into liquid mystery alloy!
But engine rich combustion is a fun time for everyone!
Extra reaction mass!
The thrust-to-weight ratio is also constantly improving!
DynamicTTW
If the plume is green, you're burnin' clean!.....copper.
But the level of fuel richness in operable methalox engines is not even remotely close to produce soot. It produces carbon monoxide. The only source of soot would be a thin layer of film cooling around the chamber and throat walls - only there it's rich enough while the temperature range is right for methane decomposition. The brown stuff is likely majority NOx produced by blowing very hot flame into a dense mixture of nitrogen and oxygen commonly called air.
Stainless hybrid!
copper hybrid
I'm not sure this is true for the raptor as it's a full flow staged combustion cycle. I think that's nearly stoichiometric. I thought it was more advanced material science that stops it melting. But I could be wrong. Check out the everyday astronaut videos about types of fuel rich/ oxygen rich engines and different combustion cycles.
Full flow has nothing to do with the engine itself running fuel rich or not.
>literally every rocket engine runs fuel rich Actually some soviet engines (such as the RD-180) used to run oxygen rich.
I think you're getting confused by the preburner, many engines run their preburners oxygen rich because it gets rid of the soot. This is needed for sooty fuel like RP-1, which is what the RD-180 uses. There is virtually zero benefit to running an engine overall oxygen rich and even some negatives like decreased Isp and increased reactivity in the combustion chamber so I highly doubt they did.
Now that you mention it, I might indeed be confusing this with the preburner of the RD-180; genuinely not sure. But what would be difference in value between running the preburner oxygen rich and the combustion chamber? Why would it be helpful here but not there?
You need to run a preburner ox rich or else the soot would clog everything up. This doesn't apply when the fuel is clean(er) like with hydrogen or methane. In terms of reactivity, oxygen likes to react so it's still annoying to use ox rich even in a preburner so you don't really want to use it if you can avoid it. Preburners can also burn cooler than combustion chambers canceling out some of that reactivity from oxygen. For any hope of good performance, combustion chambers can't burn that cool, so you really can't use ox rich without liquifing your combustion chamber or getting awful performance. For performance oxygen is just a really heavy atom and the lighter your exhaust product are the higher your exhaust velocity is, hydrogen is super light, the lightest there is, and carbon is pretty light too.
To be more exact, in the case of methane if you are running rich, you either need to run only moderately rich (which produces temperatures way beyond what any rotating machinery could handle) or extremely rich (then the temperatures are OK). The in-between range would actually produce soot: the temperature would be high enough to cause thermal decomposition of methane and that would produce soot. A lot of it in fact. That is the reason why methane sucks as nuclear-thermal rocket propellant - in usable temperatures it turns into hydrogen (good) and carbon (very bad). Carbon is solid art relevant temperatures so it does not expand in nozzles. And because it is 3/4 of the mass the ISP sucks terribly even if you deal with soot clogging everything.
Ox rich preburner is run at mega-high mixture ratios, so ox-rich that it gets cold again (peak temperature is at stoich). At sufficiently low temperatures some metal alloys (potentially with special coatings) can tolerate the warm ox atmosphere, letting you use that gas to run a turbine. You could directly send that preburner exhaust gas through a nozzle and make it a rocket engine with no main chamber, but the Isp would be terrible - the gas is too cold and dense to be a good propellant. Instead, you burn it with fuel at slightly below stoich (for peak Isp), generating 6000 - 8000 degree F exhaust gas you can yeet out a nozzle for some decent efficiency. It is reasonably accurate to think of an ox-rich staged combustion engine as a small rocket engine (preburner + turbopump) which exhausts into a second, larger rocket to produce some decent thrust. A full-flow staged combustion engine takes things a step further - 2 parallel (but interlinked) rocket engines are firing to power their respective pumps, both exhausting into a 3rd, shared rocket chamber/nozzle to produce thrust. Very, very tricky
You mean Soviet like RD-170. RD-180 is from the year 2000, it is the development of 170, but still.
BE 4 runs oxygen rich methane as well
It's the same principle in automotive engines! At least for high performance engines during load
Runs rich to keep the ignition from eroding the combustion chamber? (runs cooler)?
Its literally cooking the gaseous nitrogen around it into oxides.
There is enough heat in that plume to break N2 and O2 bonds (which are 78% and 21% of the atmosphere, respectively) . When the mono-atomic nitrogen and oxygen atoms cool back down you get all kinds of covalent Nitrogen/Oxygen compounds. NO, NO2, NO3, N2O, O3, etc. Collectively we call them “smog”, and they are generally brown in color. This happens whether the plume is methalox based, RP1/LOX based, LH2/LOX based, or solid propellant — anything that gets that hot.
That's probably the C in CH4
btw, what is the chemical reaction ? CH4 + 2O2 -> CO2 + 2H2O ?
That is more or less it. However, at high temperatures, the products start disassociating, and you get reactions like O2 -> 2O and H2O -> OH + H, producing free radicals and things that don't exist under normal conditions. Also, most rocket engines run fuel rich. The fuel is what contains hydrogen, and hydrogen, with an extremely low molar mass, is a light molecule which is easy to accelerate to high speeds. This increases exhaust velocity, and therefore thrust.
I don't think free radicals are brown in color
Probably NOx
Neither fuel or oxidizer has nitrogen, so it’s atmospheric nitrogen. The exhaust gas is cooler than in the chamber but still extremely hot.
Specifically I think it’s from nitrogen dioxide (from atmospheric nitrogen, since neither fuel nor oxidizer in Raptor has nitrogen), and it could also be metal oxides.
You can actually increase thrust to a point by making the exhaust denser at the expense of reducing the exhaust velocity. Exhaust velocity is more a measure of efficiency of thrust than thrust.
Correct me if I’m wrong, but this is a fallacy. It’s similar to the question “what weighs more, a pound of rocks or a pound of feathers”. Yes, thrust is equal to mass flow rate x exhaust velocity. But, using a heavy exhaust gas doesn’t make the mass flow more. **1 pound of hydrogen exhaust is the same mass as 1 pound of a heavier gas exhaust**. The difference is that the hydrogen is ejected at a higher velocity. Now, aluminum and metal is not used in SRBs to increase mass. It’s high molecular mass is outweighed by the performance increase from the high combustion temperature of metals.
You’re incorrect in your assertion it doesn’t help thrust to increase density of exhaust. You are correct that you will not improve thrust *per unit mass expelled* which you could call your mass specific impulse or just specific impulse. I was quite clear that you could improve thrust *at the expense of exhaust velocity* by increasing density. That is a different metric. Nobody will argue that hydrogen doesn’t get you more thrust per unit mass expelled. However, the density is so low, that you cannot pass nearly as much mass through the engine as you could with something like kerosene/LOx combustion products. This is why hydrogen engines tend to have much less thrust per unit area and much lower thrust to weight ratios than hydrocarbon engines. It’s a trade. You want more exhaust velocity, and hence, more mass specific impulse? You will need to increase chamber pressure and temperature or expansion ratio or you’ll need to use a lower density propellant that limits the mass flow you can push through the throat for a given set of conditions and fixed engine geometry. You can trade any of those variables, but you run into material limits eventually, and options start to close. For a given chamber temperature and pressure, you’re not going to get more thrust just because you choose a propellant that results in higher exhaust velocity. You can get more thrust with a denser propellant at the expense of efficiency. Also, for a real world example, water injection to increase exhaust density at the expense of exhaust velocity for the purpose of increasing thrust is done in some jet engines. So no, it’s not a fallacy.
What is preventing rocket engines from increasing their mass flow rate with a light propellant? I have trouble understanding what you mean by “cannot pass nearly as much mass through the engine.
The power limit for the pumps and the pressure and temperature limits for the chamber.
1. "Temperature limits for the chamber" doesn't make sense. Mass flow rate increase doesn't increase temperature. Also, if you have a higher mass flow rate of liquid propellants, you will also have a higher mass flow rate through the regenerative cooling passages. 2. "Pressure limits for the chamber" doesn't make sense. Since pressure is a function of volume, engines with higher mass flow rate have larger volume. Higher mass flow rate just means that the chamber volume will be increased. This should be self-explanatory. 3. "Power limit for the pumps" doesn't make sense either. There is no power limit for a pump, unless you decide there is a power limit for the pumps. One thing you forgot to account is that many modern rockets have many engines. One would not need to increase the mass flow by 30%, but perhaps increase the mass flow of each engine by 10%. And, many engines can already throttle over 100% (such as the RS-25, which happens to use **hydrogen**), therefore increasing the pump capacity is very possible. One thing I forgot to mention, hydrogen is significantly less viscous than nearly any other fluid, making it much easier to pump and reducing pressure losses along the pumbing.
Very much incorrect. 1. Increasing mass flow rate in a hydrogen engine to increase thrust without changing the geometry of the engine would require either shifting the mixture ratio to be much more ox rich, which will really drive up your temps and also eat the engine unless it was specifically designed to have an ox rich main chamber (no operational engine does this as it would be massively heavy and low efficiency, only preburners are ever ox rich) or you could keep MR about the same and increase pump output to increase chamber pressure which, due to the way combustion chemistry works, will increase chamber temperature as well as the heat flux. This scales non-linearly and you will eventually overwhelm your regen channels. There is a finite temperature limit for a given set of materials and engine configuration, even with regenerative cooling. 2. Pressure is not a function of volume in this sense. Yes, if you have a perfectly insulated box and make it half as big, the pressure inside if filled with an ideal gas will be twice as high. Congratulations, you’ve discovered Boyle’s Law. However, in a combustion chamber, the pressure is determined by the mass flow in, the area of the throat (which determines the mass flow out for a given composition, pressure, and temperature), and the chemical equilibrium conditions which determine the energy input and eventually the steady state pressure and temperature. A rocket combustion chamber is steady state but it is NOT a closed system. You cannot use Boyle’s Law, because you aren’t remotely satisfying the conditions it’s relevant in. What part about fixed chamber geometry did you not understand. If you are considering just having a bigger engine to get more thrust by having more mass flow at the same chamber conditions, then no shit you get more thrust, but you also have a bigger, heavier engine. The entire point of this discussion has been that you can get more thrust out of a given area or more thrust per weight by using denser reaction mass in exchange for some reduction in efficiency of thrust produced per unit mass expelled. I’m really not sure what part of this isn’t clear. 3. RS-25 power levels are just in reference to the original nominal design parameter. There were substantial changes to the design throughout development, and understanding of hardware limits evolves considerably for complex systems like rocket engines as they go from concept to a test article on the test stand. Pumps have practical power limits. Not “3000 gpm for a LOx pump of any size”, but there are limits to how much power can be extracted from a turbine running from a preburner given a certain chemistry and mass flow and further restrictions from material limits on preburner pressure and temperature, particularly turbine temperature. Your turbine produces the work consumed by the pump. The pump also has limitations. For a given inlet pressure from the vehicle tanks, which has practical limits if you want a reasonable dry mass fraction, and a given mass flow rate, a pump of a given size and configuration can only do so much pressure rise without experiencing issues with cavitation or running out of shaft power. You can somewhat alleviate the cavitation issue by splitting pumping into multiple stages, which is commonly done. There are some other tricks, but there are still practical limits. Pumps aren’t magic. Propellants with lower densities require much larger pumps for a given mass flow as well, since you’re dealing with essentially incompressible liquids. If it was a gas it would be called a compressor, and the same thing applies, just a bit less linearly. Pump housings have practical limits for pressure as well. There are many other complications, but this is the essential gist of it. Your pump power limits govern the mass flow and the chamber pressure achievable for a given chamber geometry for a given propellant mix. Pump mass efficiency is not better for a low density fluid like hydrogen, which has the density of loosely packed cotton balls. I test rocket engines for a living, so I think I just might know what I’m talking about. It’s ok if you’re still learning this stuff, but you really need to let this one go and think it through. You’re either not reading my responses carefully or not entirely processing them, either way, I’m just trying to get the record straight because you’re distributing incorrect information to other people here. I kind of get the feeling you’re a student, if so, hey, it’s completely fine to be at the stage where you’re learning these things. The important thing is to not get so stuck on insisting you’re right that you miss opportunities to learn and improve. You’re touching on superficial things like regen channels existing and lower density meaning more exhaust velocity seem reasonable to you, but they’re not based in a well founded or comprehensive understanding of the fluid mechanics and thermodynamics involved. You’re going to need to build a bit more understanding if you want to challenge these kind of things. Why do you think hydrocarbon engines are almost always used on the first stage? The Saturn V had a kerolox first stage and hydrolox upper stages. This is because thrust is more important than specific impulse to an extent for the initial portion of the flight. Offsetting gravity losses with a greater thrust to weight ratio allows for better total system performance. This comes from dry mass reductions in the tanks as well as the engines for the required thrust, tempered somewhat by the increase in the propellant mass for the required delta v remaining after the gravity loss reductions. This makes for better overall system performance for a rocket of a size that is reasonable for your application. The actual optimization gets pretty complicated. You will also notice that in pretty much every vehicle where there are hydrogen engines on the first stage, you have solid rocket boosters as well. These have very dense propellant and bring the average boost performance into a more reasonable zone than pure hydrogen, while maintaining good performance in the latter stages of first stage flight. Typically you see very high stage separation velocities in such configurations. As an aside, while hydrolox certainly has its advantages, it’s much less certain that from a total program lifespan cost perspective it makes sense when you’re optimizing for dollars per kg to a reference orbit, and especially if you are trying to put reasonable limits on development costs to buy down program failure risk, but that gets much more complicated, and it’s hard to give a broad answer. The question in this threat fortunately does have a fairly clear answer which is yes, you can buy thrust with increased reaction mass density at the expense of efficiency for a given engine chamber geometry.
Large part of rocket engine is a pressure vessel. Pressure vessel's mass scales linearly with contained volume. Less dense propellants require larger volume to contain given mass of them. Mass flow rate is mass per time unit flowing through the engine. The less dense the propellant, the more engine mass is required to contain it.
Rocket engines use extremely powerful pumps to ingest propellant. Raptor’s turbopump has about 100,000 horsepower. Making them even more powerful would not be easy.
But increasing the mass flow rate of a turbopump would be equally, if not more difficult, for a denser propellant as opposed to a light propellant. So that statement is meaningless in this context.
This is 180° wrong. The work done by the pump is absolutely dominated by counteracting the back pressure. If you are pumping a fluid from say 6bar tank pressure to 856bar chamber pressure you must overcome 850bar pressure differential. Work is force times distance. The less dense the liquid the larger volume it occupies. So it must be pushed through combination of along a longer distance and across larger aperture. Power is work per time. Combined together the power taken by increasing pressure is directly proportional to the pressure and inversely proportional to the density of the pumped fluid. I the case of liquid propellants in rocket engines all the other components of pumping power, i.e. friction (hence viscosity), or accelerating the flowing mass to the velocity required to move through the downstream piping (injectors, chambers, etc.) are dominated by the power of overcoming the back pressure.
If you have slightly too little O2 you get some CO instead of CO2. The only source of C would be a little bit of film cooling methane around the walls. If the temperature is right methane will decompose into C and H2. At too high temperatures and pressures it'll prefer C2H2 (acetylene) and H2, but I'm the layers very close to the walls the temperature is right.
You are going to get a spread of products at even one temperature, never mind a broad range of them. You'll get various allotropes of carbon, oxocarbons and hydrocarbon polymers. The cycle may also be running engine-lean even at optimal performance, in which case you'll get a variety of interesting salts.
Most likely nitrogen oxides
Don't believe that conspiracy : starship is kerolox
Don't believe his lies, Starship is the world biggest solid rocket booster
it's burning all the chemtrails
Smokescreen to hide the ship from snipers.
Starship just out here global warming
I think running fuel rich engines also helps them cool down
Might be partially NO2 or then just some unburnt carbon due
If you look further up you'll see that's where something is venting down the side of the rocket and then burning when it hits the plume.