This see thru video was much more interesting to me just because of how much more visible everything was. Sure, the slo-mo helped a lot in being a better video, but the device itself did as well.
The build of the jet engine was well done, but just the visibility isn't nearly as visible into what's actually going on inside. Sure, you can see the metal casing of the combustion chamber get red hot, but it just didn't have the wow factor I was hoping for. This is not a knock on the person making the video. Sometimes you get snake eyes after all of that effort. Sometimes you roll seven, and get even better than hoped for. All in the day of the life of a content creator.
Yeah? And? So? Keep in mind now, that for the most recent CPUs are no longer 8-bit (20 years?), yet 8 bit microcontrollers are all the fad. Not really sure what your point is? We're not allowed to be fascinated with old tech that still works and has merit in demonstrating how/why it worked?
Speaking of which, I would ban these. It's ridiculous that fuel injection technology is universally used in cars for over 30 years, yet exhaust of garden equipment still stinks horribly. (oh, and also ban oil-burning two-strokes)
This is exactly why I go on HN. I got stuck at sea so many times with no wind to get back and a broken outboard where as a reflex I was taking apart the carburator and putting it back together as a quick fix without really understanding what was going on, this is massively helpful thanks!
This is a very nice demonstration of turbojet. While it has not been claimed as such in the video, this is how most of the modern airplanes DO NOT work.
Modern airplanes use turbofans which has an extremely high bypass ratio. In a turbofan, the turbine mostly rotates the various stages and creates very little thrust.
What’s the deal with high bypass turbofans vs turboprops?
Turbofans look a lot like the jet engine in this video but, as you point out, only the core is the turbine and most of what’s inside the cowling is a fan for pulling air past the engine, which pulls the plane forwards.
Turboprops do this but, to a lay person, in a different way. Four propeller blades instead of hundreds, and no cowling. Is it just cheaper to build a turboprop, and turbofans are the ultimate in terms of performance?
I think turboprops are less efficient, because at high airspeeds, drag is created by vortices generated at the tips of the blades, which are exposed to the free air.
By contrast, the cowling around a turbofan prevents such vortices from being created. Ducted fans are, in general, a more efficient design because of this.
Turboprops indeed fight wing tip vortices. But they also have a problem of not being able to shape air properly like Turbofans can do with stator vanes and such. As you stated, ducts are awesome.
Turbo props and turbo fans, both produce thrust, but very differently.
In a turbofan, the turbine spends most of its effort turning the compressor blades. In a modern turbofan (TF) engine like GE 9x, you are looking at a 80-20 bypass ratio. This means, 80% of the air drawn from the front just flows through. Now, as the "FAN" (duh!) rotates, a huge volume of air is drawn in and compressed. As the volume moves deeper into the engine, the air pressure increases because the engine tapers down, as well as the compressor works in "compressing" the voume of air. As the air pressure increases, so does the speed and temperature. Now, as this compressed air exits from the back of the engine, it provides a forward thrust. The 20% air that DID get into the core, carrys on the combustion in the engine that provides the energy required to rotate the compressor blades. Now this is definitely simplified as there are many additional pieces, like stator vanes, equipment to shape and correct the velocity of the air, fuel injectors, igniters, afterburners (not in civilian airplanes of course) and many more contraptions.
While not completely comparable, a turbo prop (TP) behaves a lot like its piston counterpart. The turbine is simply replacing a traditional 4 stroke IC engine in this case and just tasked with rotating the blades. Each of the blade is a mini (well not so mini) wing that has the shape of an airfoil. As these rotate, they generate lift but since the rotational axis is parallel to the fuselage, it is actually thrust.
> Is it just cheaper to build a turboprop, and turbofans are the ultimate in terms of performance?
It is definitely cheaper to build a TP. And TPs can take a lot of abuse and needs miniscule maintanance compared to TFs. But if you are talking of raw performance, turbojets (TJ) actually run (fly?) around in circles around TFs. TJs are just incredibly wasteful, inefficient and needs a lot of maintenance.
It sounds like you know your way around a turbofan, but you wrote some confusing stuff, mainly this:
> Now, as this compressed air exits from the back of the engine, it provides a forward thrust.
This is something you generally don't want to happen in a reaction engine. If you eject above-atmospheric-pressure out of your engine, the expansion takes place downstream, in the wake, and that energy is mostly wasted (some pressure thrust is still created, but it's minuscule). You always want all the expansion to happen inside your engine, in a properly designed propelling nozzle, which should always result in a flow of low-pressure high-speed air leaving the engine.
If the nozzle fails to expand the air (choked flow, nozzle too short, wrong shape), you're losing thrust (and efficiency).
So, to confirm for the comparison, in a nutshell (Some of the upstream posts in this sub-thread imply the bypass section of a TF provides its thrust directly, which I believe your post is showing is incorrect):
- Treat TJ as base model
- TF uses the bypass section to provide additional power to the compressor
- TP generates its thrust from a prop on the shaft, vice the core's output (is this called exhaust?). TJ and TF both use the core's output for thrust.
Energy is proportional with the speed squared, momentum with the speed. You have a fixed budget of energy, initially stored as chemical energy in jet fuel. The engines have a more or less fixed thermodynamic efficiency, so you transform the chemical energy in a fixed amount of kinetic energy of the air being pushed out of the engine. You want that air to have as much momentum as possible (Newton’s action-reaction principle: the momentum the plane gives to the air is the momentum the air gives to the plane).
Now, for a given energy, you can push some mass of air with speed v or 4 times as much with speed v/2. The momentum of the air in the latter case is double the first case: 4 m v/2 = 2mv. You always want to move more air at lower speed.
You achieve that with higher and higher bypass ratios. Of course, with a higher bypass ratio the engine becomes larger and heavier (and draggier), so there are limits. But if you could increase the bypass ratio without increasing the engine weight and drag, you would always do it. As a bonus, with lower air speed come lower vibrations and noise.
Also a limiting factor on turboprops (due to their larger diameter) run at a much lower rpm… they’re geared down quite a bit m.
This is a complicated area, but in general basically: Supersonic flow is bad. At best you get a ton of noise (like the Russian Tu-95). Efficiency also goes to hell.
This is also why the SR-71 had those crazy moving brake cones… they manipulated the shockwave so the air actually entering the engine was subsonic.
Funnily it seems to be the exact opposite of the rocket equation, where you want to eject the limited mass quantity you have as fast as possible. The fuel being more like a mass reserve than a chemical energy reserve.
The fuel in an airplane is also exhausted but I'm not sure how much of momentum is gained due to this rocket fuel effect, in the traditional working regime.
For lower stages, you can inject water into the nozzle and ISP (exhaust velocity) goes down, but thrust goes up, more than compensating for the weight of the water tanks.
But you wouldn't want to do that for an upper stage, as all the rocket below that would need to carry the water tank to a higher velocity.
Saturn V second stage had variable mixture ratio, using a leaner mixture with lower ISP (but better density and thrust) at lower altitude and richer mixture with higher ISP at higher altitudes. https://yarchive.net/space/launchers/saturn_variable_mixture...
So, ISP is clearly not everything. At low altitudes, you're looking at basically maximum thrust and thrust to weight. At some level you're also looking at volume and thus density, to minimize cost.
Cool, glad to learn something :), I'm more of a AI neurosurgeon than a rocket scientist.
My upper point was more about asking about the regime in which we can consider an airplane a constant mass system, and could totally neglect the mass of the consumed fuel compared to the momentum of the air moved.
One of the main difference between a jet engine and a rocket engine is that the rocket engine carry its own oxygen, but I'm also not quite sure how much the momentum of the surrounding air is relevant, in helping push the plane forward.
There is probably some kind of transition where either when air density get lower, or when speed get higher, than your jet engine behave more like a rocket motor than a blowing-fan (like the turbo-prop vs turbo-fan in the other comments).
The Breguet range equation is quite similar to the rocket equation. The airplane has the dimensionless lift to drag ratio to help. And indeed mass ratios are not as drastic for airplanes as rockets, but still matter over long distances.
I wonder how this equation would apply for electrical planes.
And whether it could be useful to carry some water as fuel-mass (like in your saturn motor example) that would be ejected via the use the electrical motor to convert more efficiently the energy that is stored in the batteries into the momentum needed to make the plane go forward.
The kerosene plane is lighter towards the end of the flight, while the electric one's mass doesn't change.
For rockets, there's Rocket Lab's Electron that uses electric pumps to pump the regular propellants to the combustion chamber. This was much easier to develop than a turbine for the turbopump.
It drops some spent batteries along its flight.
Could an electric airplane drop batteries at certain points? Like after climbing to target altitude. A lot of operational headaches there, so not likely. There have been some military planes that have used takeoff trolleys or takeoff rockets so there is some precedence there.
Rocket engines are extremely inefficient, they just have to be used because of the very wide speed range they support. So it wouldn't really make sense to eject water with an electric pump. Batteries have such low energy density and the pumps are heavy too. It's the worst combination to carry your own reaction mass and propel it with battery stored energy!
A rocket engine has tremendous trust to weight and power density, it can be gigawatts in a car sized package. This is enabled by the very energy dense liquid fuel and liquid oxidizer. Even a jet engine pales in comparison, as it has to use gaseous air as oxidizer which is not dense, and which is 80% "fluff" nitrogen. Hence the jet engine compressors have to be large, and the compression ratios are meager 10 to 20 bars. A rocket engine can run at 200 bars main chamber pressure. The compressors (pumps) can operate at hundreds of megawatts (this is actually not far from big jet engines). The thrust to weight ratio can be 100 to 1, while for jets it's 10 to 1.
Edit : (We just have to put an electric jetski inside a fire-fighting plane, and have the Flyride pull the plane forward until it reach lift-off speed, (no expensive fuel needed :) )
With rockets, the mass of whatever you eject is constant, so you can’t apply the lower speed-increased momentum effect.
In outer space, the rocket equation says that the delta-v is only a function of propellant mass and exhaust velocity. It is always better to have higher exhaust velocity, no exceptions.
But until you get to orbit, you experience the “gravity drag”. Gravity acts in the opposite direction of your motion, and reduces your speed, so you end up needing overall more delta-v.
In order to reduce the effect of the gravity drag, you want to accelerate as quickly as possible. That’s why astronauts have to be able to withstand high Gs for a few minutes. Ballistic missiles, which are not limited by the human limits can (and do) acceletate much faster. A Trident II gets to about 50 Gs.
To get to these insane accelerations, you need huge thrust. You achieve that at the cost of lower exhaust velocity (like in the case of the Space Shuttle). Lower exhaust velocity means lower delta-v, but higher thrust means lower delta-v losses due to the gravity drag, so in many cases this tradeoff makes sense.
And it's not only the gravity but density advantage. delta vee = v_ex times ln (mass ratio). It's the mass ratio part.
The problem for first stages is that some high ISP propellants are very non-dense. Liquid hydrogen is 70 kg per cubic meter. Kerosene is about ten times as dense.
So the tanks need to be a lot bigger for the same payload. The same holds for pumps and turbines too. So the engines need to be bigger for the same thrust. Hydrogen is also annoying to deal with as it's much colder than liquid oxygen, the molecule is also much smaller and leaks much more easily.
Technically the fueled all-up hydrogen rocket might be lighter, so the engines should be smaller, but the density disadvantage eats back all that advantage. The dry mass, which is the actual cost driver, is about the same.
For upper stages, it makes sense because the cost driver is not the dry mass but the wet mass. Also they need less thrust so engine mass is a smaller proportion of stage mass.
If you want to maximize the energy efficiency of a rocket, you want variable exhaust velocity, and make the exhaust velocity = - the velocity of the rocket. That leaves the exhaust motionless in the initial reference frame, so it will have no kinetic energy (in that reference frame). Therefore, all the kinetic energy being generated by the rocket is being retained by the rocket.
(This ignores potential energy from changes in altitude.)
This is one reason why it's better to use a lower Isp propellant on the first stage of a multistage launcher (also, those propellants are much denser.)
The other comments have some good explanations, I just want to to confuse you further by pointing out the "open rotor engines"
https://en.m.wikipedia.org/wiki/Propfan
exist and are different from both
They're both exploiting the same principal. It's more efficient to generate thrust by accelerating a large amount of air a little than a little amount of a air a lot.
Figure of merit is the bypass ratio. Props are 50-100, fans are 3-10.
So big advantage of a turboprop is a much higher bypass ratio than fans. Downside is props can't spin as fast and so need a high speed gearbox to drive them. And props loose efficiency at high speeds.
I think for large fast aircraft turbofans are a win. For medium sized slower aircraft turboprops are a win.
Basically this plot is why. Efficiency versus speed characteristics due to aerodynamics of propellors versus cowled rotors. Turboprop for short low flights, high bypass turbofan for long and high flights.
The main difference between a turboprop and a turbofan is the speed of sound. A propeller needs to avoid breaking the speed of sound, otherwise you get obscene noise, a reduction in efficiency, and potentially damage to the propeller (although there are some rare propellers that do regularly break the sound barrier). This means that propeller-driven planes must have a speed limit that is quite a lot lower than the speed of sound, to avoid the propeller tips breaking that speed barrier.
A turbofan however is typically installed on planes that want to fly around 80% of the speed of sound. That's really close to the limit, and there isn't enough leeway there for a propeller to get away with it. A turbofan manages by using lots of blades to more gently push the air back, and then the cowling helps by preventing blade-tip vortices, and by constricting the airflow slightly, which increases the air's backwards speed.
So a propeller and a turbofan are suitable for different plane speed regimes. When you're travelling slowly, it is most efficient to push a very large amount of air back just a little bit faster than the plane is moving, and a propeller is great at doing that. As you approach the speed of sound, that becomes infeasible, and a turbofan becomes more efficient. If you're trying to break the speed barrier, then you need a lower bypass ratio, or you may even need to switch to a turbojet.
Virtually all turbofan engine blades exceed the speed of sound at high N1% even when moving slow (like during takeoff). This is part of what generates the distinct "buzzsaw" sound at or near max thrust when seated forward of the engines. The pitch/tone of the buzzsaw sound depends on the number of fan blades and rpm of the main fan (and likely stator design).
In this video of an Iran Air 747SP the buzzsaw effect is particularly isolated/exaggerated due to the camera/mic placement, the use of 1970's design (louder) JT9D engines, and the shorter length of the 747SP. https://youtu.be/3lhHwKK-6ms
On a 737 the pitch is higher due to the smaller diameter fan (higher max RPM), while a 777 has the lowest buzzsaw pitch due to having the largest fan diameter.
The most recent jet engines (like on the 747-8) have highly suppressed this sound with contoured fan blades.
> In a turbofan, the turbine mostly rotates the various stages and creates very little thrust.
That undersells the turbine. Turbine powers not only the various compressor stages, but also the turbofan itself. All power comes from the turbine and it is is arguably the most challenging part of a jet engine. Turbine blades are tiny and fit in the palm of your hand with fingers closed [1]. Temperatures through the roof, it needs compressor bleed air to keep the turbine blades cool or the blades would melt. I was part of the team that did mechanical design of turbine blades and one of interesting challenges I remember was the trade off between the pressure-side bleed vs. trailing-edge dump techniques of cooling turbine blades. Aero engineers prefer razor thin trailing edge for better efficiency (pressure-side bleed as you can see in the picture). These are tiny holes/slots on the edge of the blade that bleed compressor air from various stages which is at 800F and trying to cool 2200F blade edge. So mechanical engineers (us) would prefer a trailing-edge-dump technique of cooling where we would make the edge thicker and drill the holes at the apex of the edge for better cooling performance. Pressure-side bleed usually won because higher fuel efficiency requirements and pushed us to do other things to cool the blade, even if that requires pulling air from a cooler compressor stage, which impacts efficiency too.
Another fun fact, the whole thing leaks hot gas like a sieve until engine reaches temperature and the tip of the turbine blades seal the hot gas path through elongation along the radial direction. Reminds me of the famous SR-71 fuel tank seals.
A turbine, by definition, provides zero thrust, because it works the other way (it extracts energy from the flow).
What you meant to say was that the air flowing through the engine core provides very little thrust, because most of its energy is being extracted by the turbine almost immediately after combustion.
Technically, the nozzle provides thrust in the hot gas path. Point taken though. In jet engines, you're correct, most of the power is robbed by the turbine. :-)
Note that here, the largest source of thrust is from the components that jets and rockets have in common (combustion chamber and the divergent section after the turbine) - though here, in a non-afterburning jet engine, the nozzle must be convergent rather than divergent, and that imposes a cost.
I think what you mean is very little thrust exits the turbine. In other words, the air expansion from burning the fuel is used almost entirely by the turbine which, in turn, powers both the compressor and a huge fan which provides the actual thrust used for flight.
Every time I see explanations about jet engines, I have the same basic question that never seems to get addressed. I think I know what the answer, but anyhow: why do the exhaust gases flow in the direction that we want? What prevents them from exiting from the intake?
Not outside the intake (outside the engine). Basically, to answer the question fully, you also need to say why the pressure is high there, and why it is higher than on the other side of the engine.
Because the compressor is on the intake side of the combustion. The compressor increases the pressure more on the intake side than the resistance from the turbine does on the output side.
It's been so long since I took thermo and I have no engineering experience with gas turbines but.
You add heat at constant pressure to the air flowing through the engine. Which means the volume has to increase. Only way balance things out is the velocity of the air has to increase as well.
If you look at the TS diagram state 1 is the intake. 2 is after the compressor. 3 is after the burners. 4 is the exit from the turbine. Temperature is decent proxy for total energy. You can see the vertical lines of constant entropy. That means the energy is being extracted as work. Note real compressors and turbines the lines are angled slightly which reflects inefficiency.
You'll note the difference between 3->4 is larger than 1->2. The turbine expanding the hot gas produces more power than the compressor does compressing the cold air.
It happens - compressor surge (https://en.wikipedia.org/wiki/Compressor_stall). Concorde infamously suffered from this during development. The modern answer is that there's an electric starter motor which starts the compressor before the fuel is injected, and better computers controlling the fuel flow.
Generally, it is because it is harder for the air to exit from the intake than it is to exit through the exhaust.
It can be because the inlet into the combustion chamber is larger than the outlet, because the blade pitch is finer on the compressor side than it is on the turbine side, because of the specific geometry of various components, ... Often all of that.
The intake is optimized for low speed high pressure (lots of elements, small space, fine pitch, ...) , the exhaust is optimized for high speed low pressure (less elements, more space, coarser pitch), and naturally, the air will flow from the high pressure zone to the low pressure zone.
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[ 3.1 ms ] story [ 141 ms ] threadThe build of the jet engine was well done, but just the visibility isn't nearly as visible into what's actually going on inside. Sure, you can see the metal casing of the combustion chamber get red hot, but it just didn't have the wow factor I was hoping for. This is not a knock on the person making the video. Sometimes you get snake eyes after all of that effort. Sometimes you roll seven, and get even better than hoped for. All in the day of the life of a content creator.
Not sure why you're so angry‽
Glad I could help!
Modern airplanes use turbofans which has an extremely high bypass ratio. In a turbofan, the turbine mostly rotates the various stages and creates very little thrust.
Turbofans look a lot like the jet engine in this video but, as you point out, only the core is the turbine and most of what’s inside the cowling is a fan for pulling air past the engine, which pulls the plane forwards.
Turboprops do this but, to a lay person, in a different way. Four propeller blades instead of hundreds, and no cowling. Is it just cheaper to build a turboprop, and turbofans are the ultimate in terms of performance?
By contrast, the cowling around a turbofan prevents such vortices from being created. Ducted fans are, in general, a more efficient design because of this.
Turbo props and turbo fans, both produce thrust, but very differently.
In a turbofan, the turbine spends most of its effort turning the compressor blades. In a modern turbofan (TF) engine like GE 9x, you are looking at a 80-20 bypass ratio. This means, 80% of the air drawn from the front just flows through. Now, as the "FAN" (duh!) rotates, a huge volume of air is drawn in and compressed. As the volume moves deeper into the engine, the air pressure increases because the engine tapers down, as well as the compressor works in "compressing" the voume of air. As the air pressure increases, so does the speed and temperature. Now, as this compressed air exits from the back of the engine, it provides a forward thrust. The 20% air that DID get into the core, carrys on the combustion in the engine that provides the energy required to rotate the compressor blades. Now this is definitely simplified as there are many additional pieces, like stator vanes, equipment to shape and correct the velocity of the air, fuel injectors, igniters, afterburners (not in civilian airplanes of course) and many more contraptions.
While not completely comparable, a turbo prop (TP) behaves a lot like its piston counterpart. The turbine is simply replacing a traditional 4 stroke IC engine in this case and just tasked with rotating the blades. Each of the blade is a mini (well not so mini) wing that has the shape of an airfoil. As these rotate, they generate lift but since the rotational axis is parallel to the fuselage, it is actually thrust.
> Is it just cheaper to build a turboprop, and turbofans are the ultimate in terms of performance?
It is definitely cheaper to build a TP. And TPs can take a lot of abuse and needs miniscule maintanance compared to TFs. But if you are talking of raw performance, turbojets (TJ) actually run (fly?) around in circles around TFs. TJs are just incredibly wasteful, inefficient and needs a lot of maintenance.
> Now, as this compressed air exits from the back of the engine, it provides a forward thrust.
This is something you generally don't want to happen in a reaction engine. If you eject above-atmospheric-pressure out of your engine, the expansion takes place downstream, in the wake, and that energy is mostly wasted (some pressure thrust is still created, but it's minuscule). You always want all the expansion to happen inside your engine, in a properly designed propelling nozzle, which should always result in a flow of low-pressure high-speed air leaving the engine. If the nozzle fails to expand the air (choked flow, nozzle too short, wrong shape), you're losing thrust (and efficiency).
Now, for a given energy, you can push some mass of air with speed v or 4 times as much with speed v/2. The momentum of the air in the latter case is double the first case: 4 m v/2 = 2mv. You always want to move more air at lower speed.
You achieve that with higher and higher bypass ratios. Of course, with a higher bypass ratio the engine becomes larger and heavier (and draggier), so there are limits. But if you could increase the bypass ratio without increasing the engine weight and drag, you would always do it. As a bonus, with lower air speed come lower vibrations and noise.
glider wing area ~10m2, F16 ~30m2, F22 ~80m2
This is a complicated area, but in general basically: Supersonic flow is bad. At best you get a ton of noise (like the Russian Tu-95). Efficiency also goes to hell.
This is also why the SR-71 had those crazy moving brake cones… they manipulated the shockwave so the air actually entering the engine was subsonic.
The fuel in an airplane is also exhausted but I'm not sure how much of momentum is gained due to this rocket fuel effect, in the traditional working regime.
For lower stages, you can inject water into the nozzle and ISP (exhaust velocity) goes down, but thrust goes up, more than compensating for the weight of the water tanks.
But you wouldn't want to do that for an upper stage, as all the rocket below that would need to carry the water tank to a higher velocity.
Saturn V second stage had variable mixture ratio, using a leaner mixture with lower ISP (but better density and thrust) at lower altitude and richer mixture with higher ISP at higher altitudes. https://yarchive.net/space/launchers/saturn_variable_mixture...
So, ISP is clearly not everything. At low altitudes, you're looking at basically maximum thrust and thrust to weight. At some level you're also looking at volume and thus density, to minimize cost.
My upper point was more about asking about the regime in which we can consider an airplane a constant mass system, and could totally neglect the mass of the consumed fuel compared to the momentum of the air moved.
One of the main difference between a jet engine and a rocket engine is that the rocket engine carry its own oxygen, but I'm also not quite sure how much the momentum of the surrounding air is relevant, in helping push the plane forward.
There is probably some kind of transition where either when air density get lower, or when speed get higher, than your jet engine behave more like a rocket motor than a blowing-fan (like the turbo-prop vs turbo-fan in the other comments).
The other day I saw the TechIngredients video about afterburner https://www.youtube.com/watch?v=EP4Hf63zR6w that I found interesting.
Edit: The wikipedia page about specific impulse https://en.wikipedia.org/wiki/Specific_impulse and in particular its graph https://upload.wikimedia.org/wikipedia/commons/thumb/4/4f/Sp... helped me quantify the importance of air in the propulsion efficiency of a jet engine with respect to a rocket engine.
For rockets, there's Rocket Lab's Electron that uses electric pumps to pump the regular propellants to the combustion chamber. This was much easier to develop than a turbine for the turbopump. It drops some spent batteries along its flight.
Could an electric airplane drop batteries at certain points? Like after climbing to target altitude. A lot of operational headaches there, so not likely. There have been some military planes that have used takeoff trolleys or takeoff rockets so there is some precedence there.
Rocket engines are extremely inefficient, they just have to be used because of the very wide speed range they support. So it wouldn't really make sense to eject water with an electric pump. Batteries have such low energy density and the pumps are heavy too. It's the worst combination to carry your own reaction mass and propel it with battery stored energy!
A rocket engine has tremendous trust to weight and power density, it can be gigawatts in a car sized package. This is enabled by the very energy dense liquid fuel and liquid oxidizer. Even a jet engine pales in comparison, as it has to use gaseous air as oxidizer which is not dense, and which is 80% "fluff" nitrogen. Hence the jet engine compressors have to be large, and the compression ratios are meager 10 to 20 bars. A rocket engine can run at 200 bars main chamber pressure. The compressors (pumps) can operate at hundreds of megawatts (this is actually not far from big jet engines). The thrust to weight ratio can be 100 to 1, while for jets it's 10 to 1.
Until then we have to stay tethered to the sea with something like Zapata's Flyride / Flyboard https://www.youtube.com/watch?v=PCAvIHGylKo
Edit : (We just have to put an electric jetski inside a fire-fighting plane, and have the Flyride pull the plane forward until it reach lift-off speed, (no expensive fuel needed :) )
In outer space, the rocket equation says that the delta-v is only a function of propellant mass and exhaust velocity. It is always better to have higher exhaust velocity, no exceptions.
But until you get to orbit, you experience the “gravity drag”. Gravity acts in the opposite direction of your motion, and reduces your speed, so you end up needing overall more delta-v.
In order to reduce the effect of the gravity drag, you want to accelerate as quickly as possible. That’s why astronauts have to be able to withstand high Gs for a few minutes. Ballistic missiles, which are not limited by the human limits can (and do) acceletate much faster. A Trident II gets to about 50 Gs.
To get to these insane accelerations, you need huge thrust. You achieve that at the cost of lower exhaust velocity (like in the case of the Space Shuttle). Lower exhaust velocity means lower delta-v, but higher thrust means lower delta-v losses due to the gravity drag, so in many cases this tradeoff makes sense.
The problem for first stages is that some high ISP propellants are very non-dense. Liquid hydrogen is 70 kg per cubic meter. Kerosene is about ten times as dense. So the tanks need to be a lot bigger for the same payload. The same holds for pumps and turbines too. So the engines need to be bigger for the same thrust. Hydrogen is also annoying to deal with as it's much colder than liquid oxygen, the molecule is also much smaller and leaks much more easily.
Technically the fueled all-up hydrogen rocket might be lighter, so the engines should be smaller, but the density disadvantage eats back all that advantage. The dry mass, which is the actual cost driver, is about the same.
For upper stages, it makes sense because the cost driver is not the dry mass but the wet mass. Also they need less thrust so engine mass is a smaller proportion of stage mass.
(This ignores potential energy from changes in altitude.)
This is one reason why it's better to use a lower Isp propellant on the first stage of a multistage launcher (also, those propellants are much denser.)
Figure of merit is the bypass ratio. Props are 50-100, fans are 3-10.
So big advantage of a turboprop is a much higher bypass ratio than fans. Downside is props can't spin as fast and so need a high speed gearbox to drive them. And props loose efficiency at high speeds.
I think for large fast aircraft turbofans are a win. For medium sized slower aircraft turboprops are a win.
https://en.wikipedia.org/wiki/Turbofan#Efficiency
https://en.m.wikipedia.org/wiki/Propulsive_efficiency#/media...
A turbofan however is typically installed on planes that want to fly around 80% of the speed of sound. That's really close to the limit, and there isn't enough leeway there for a propeller to get away with it. A turbofan manages by using lots of blades to more gently push the air back, and then the cowling helps by preventing blade-tip vortices, and by constricting the airflow slightly, which increases the air's backwards speed.
So a propeller and a turbofan are suitable for different plane speed regimes. When you're travelling slowly, it is most efficient to push a very large amount of air back just a little bit faster than the plane is moving, and a propeller is great at doing that. As you approach the speed of sound, that becomes infeasible, and a turbofan becomes more efficient. If you're trying to break the speed barrier, then you need a lower bypass ratio, or you may even need to switch to a turbojet.
In this video of an Iran Air 747SP the buzzsaw effect is particularly isolated/exaggerated due to the camera/mic placement, the use of 1970's design (louder) JT9D engines, and the shorter length of the 747SP. https://youtu.be/3lhHwKK-6ms
On a 737 the pitch is higher due to the smaller diameter fan (higher max RPM), while a 777 has the lowest buzzsaw pitch due to having the largest fan diameter.
The most recent jet engines (like on the 747-8) have highly suppressed this sound with contoured fan blades.
That undersells the turbine. Turbine powers not only the various compressor stages, but also the turbofan itself. All power comes from the turbine and it is is arguably the most challenging part of a jet engine. Turbine blades are tiny and fit in the palm of your hand with fingers closed [1]. Temperatures through the roof, it needs compressor bleed air to keep the turbine blades cool or the blades would melt. I was part of the team that did mechanical design of turbine blades and one of interesting challenges I remember was the trade off between the pressure-side bleed vs. trailing-edge dump techniques of cooling turbine blades. Aero engineers prefer razor thin trailing edge for better efficiency (pressure-side bleed as you can see in the picture). These are tiny holes/slots on the edge of the blade that bleed compressor air from various stages which is at 800F and trying to cool 2200F blade edge. So mechanical engineers (us) would prefer a trailing-edge-dump technique of cooling where we would make the edge thicker and drill the holes at the apex of the edge for better cooling performance. Pressure-side bleed usually won because higher fuel efficiency requirements and pushed us to do other things to cool the blade, even if that requires pulling air from a cooler compressor stage, which impacts efficiency too.
Another fun fact, the whole thing leaks hot gas like a sieve until engine reaches temperature and the tip of the turbine blades seal the hot gas path through elongation along the radial direction. Reminds me of the famous SR-71 fuel tank seals.
Jet engines are insane. So much complexity.
[1] https://www.theengineer.co.uk/media/4kgoz1mg/rr-turbine-blad...
What you meant to say was that the air flowing through the engine core provides very little thrust, because most of its energy is being extracted by the turbine almost immediately after combustion.
https://i.stack.imgur.com/47W9J.jpg
Note that here, the largest source of thrust is from the components that jets and rockets have in common (combustion chamber and the divergent section after the turbine) - though here, in a non-afterburning jet engine, the nozzle must be convergent rather than divergent, and that imposes a cost.
[0]: https://www.youtube.com/c/TechIngredients
You add heat at constant pressure to the air flowing through the engine. Which means the volume has to increase. Only way balance things out is the velocity of the air has to increase as well.
Ah, there's the key.
If you look at the TS diagram state 1 is the intake. 2 is after the compressor. 3 is after the burners. 4 is the exit from the turbine. Temperature is decent proxy for total energy. You can see the vertical lines of constant entropy. That means the energy is being extracted as work. Note real compressors and turbines the lines are angled slightly which reflects inefficiency.
You'll note the difference between 3->4 is larger than 1->2. The turbine expanding the hot gas produces more power than the compressor does compressing the cold air.
Another related question: what happens if you don't rotate the engine before you ignite it?
But jet engines can and do backfire spectacularly at times. You can find lots of cool footage from hobby turbine engine enthusiasts on the tubes.
It can be because the inlet into the combustion chamber is larger than the outlet, because the blade pitch is finer on the compressor side than it is on the turbine side, because of the specific geometry of various components, ... Often all of that.
The intake is optimized for low speed high pressure (lots of elements, small space, fine pitch, ...) , the exhaust is optimized for high speed low pressure (less elements, more space, coarser pitch), and naturally, the air will flow from the high pressure zone to the low pressure zone.
Did he mention how much thrust it produced?
https://youtu.be/L24Wf0VlTE0