I'm a freshman in mechanical engineering, looking to go into propulsion, and even though I have heard some of the beginning stuff and some of the later stuff went over my head, this was great! I am absolutely sure I will be coming back to this in the future as a visual aid. Keep up the good work man!
Oh and idk if im in the place to suggest things, or if this is your area of expertise, but I like reading about high speed, high altitude flight and have no clue how compression lift works, so if you know where I can go to find a more straighforward explaination, I'd welcome it with open arms.
@@jonahgimmi1528 Hi thanks for the kind words. Regarding compression lift, the key thing to understand is oblique shocks. I would recommend Chapter 4 of Modern Compressible Flow by Anderson. You can skip over most of the maths and just read the text + study the images to get a basic understanding. Also this website has a reasonable illustration: www.aerodynamics4students.com/introduction/images-of-lift.php. The important thing to note is that the oblique shocks generate regions of high pressure, but oblique shocks only occur if we are turning/deflecting a supersonic flow. So if we have an aircraft which deflects the flow underneath (high pressure) and doesn't deflect the flow on the top (low pressure), we get a large pressure difference (lift). You can imagine a wedge with the top surface parallel to the direction of flight. I'm planning to make a video in the next few months about waveriders, which are a special kind of hypersonic vehicle, and that video will definitely cover the basics of compression lift.
@@jonahgimmi1528 Hi, I've just finished a video on the basics of compression lift, I hope this can be of some help to you! ruclips.net/video/pFb8GIEkMVw/видео.html
Congratulations Master for your beautiful presentation and interesting manter! I learnt so much! Thank's a lot and my best regards! Jacareí-Sao Paulo- Brazil.
Thanks for watching. I'll do my best to make a video on the MoC in the near future, but I can't promise anything specific! It's a complicated topic to try and summarise.
@@dhrumil979 I would suggest Modern Compressible Flow by Anderson. In Chapter 7 he introduces the Method of Characteristics for unsteady 1D flows like shock waves and expansion fans in tubes. I would suggest reading this chapter very well first. The important point to understand is that with unsteady flows we are 'following' points of constant properties as they travel through 1D space and time. On the x-t diagram we're also studying how the speed of sound (and thus temperature) changes when shock waves and expansion fans interact with each other. Once you understand that, it is much easier to understand the 2D method of characteristics, because all we have done in the 2D method of characteristics is replaced the time axis with the y direction, and assumed constant flow (so no variations in time, only with location in 2D space). In the later chapters Anderson discusses the MoC for 2D and numerical methods etc.
Hi! I'm currently designing a Nozzle for a CubeSat, I used MOC to obtain the design of the wall. I then passed it to ANSYS and i have an underexpansion effect which was expected if my ambient pressure is 0. So, I want to know how much I can expand the exit area to get the pressures to balance. I want to model this just like your video where I can have two charts and see the change in Mach and Pressure changing the velocity and exit area, what program do you use or what can you recommend for this. Thank you!
Hi, thanks for watching. I hope that I am understanding your first question correctly - it is not possible/practical in reality to match the exit static pressure of the nozzle with the vacuum pressure (~0) in space. It is possible to calculate a maximum theoretical exit velocity from a nozzle where all the thermal energy is converted to kinetic energy and the exit temperature is 0 K (and therefore the exit static pressure is 0), but this would require a very long (and heavy) nozzle with a very big expansion ratio. Additionally a real gas would condense into liquid or solid at a temperature >0 K and so the maximum ideal gas expansion could not be achieved. Additionally + 1, viscous losses with the long nozzle wall would remove some amount of energy from the flow, further preventing us reaching this theoretical idea. In practice, all space nozzles are under-expanded, and the designer has to make a trade-off between factors such as performance (Isp), mass, packaging (length and diameter) etc. etc. Regarding the plots of Mach / pressure / temperature - I made them in matlab using the geometry of a simple converging-diverging nozzle, and the isentropic flow relations (please see the link below). If you know the cross section area at each position along the nozzle, you can use equation 9 (the Mach-area relation) to calculate the Mach number at each position. Once you know the Mach number at each position along the nozzle, you can use equations 6 and 7 to calculate the temperature and pressure at each location. www.grc.nasa.gov/www/k-12/airplane/isentrop.html Hope that helps!
Great explanation with very explanatory graphics. Many thanks. A question though: what are design parameters for a nozzle to create certain effects? E.G., if I wanted te create maximum shockwaves, regardless wat happens to the airspeed at the exit, what would that mean for the design of the nozzle?
Thanks for watching and the kind words. That's an interesting question and I had to take some time to think about it. On a fundamental level, the most important variable that we can control is the pressure ratio between the total pressure in the chamber and the pressure of the environment at the exit. Please see the image I linked at the end of this message. This following paragraph assumes that we have a constant nozzle geometry, like shown in the linked image. At point 7 we have perfectly expanded flow at the exit and no shocks. As we move upwards to point 6, the environment pressure is a little bit higher than the final nozzle pressure, and so some 'weak' oblique shocks are needed to match the exit pressure to the environment pressure. As we increase the environment pressure more, the weak oblique shocks eventually turn into a 'strong' normal shock at point 5. This shock occurring at point 5 is the 'strongest' shock that a converging-diverging nozzle (of a given geometry) can create. If we increase the environment pressure more, the normal shock has to move into the nozzle and it gets weaker because the subsonic expansion now contributes to some of the pressure recovery. If we want to create a stronger normal shock at the nozzle exit, we need to increase the mach number of the flow at the nozzle exit, and so we need a bigger nozzle exit. However, with the same chamber-to-environment pressure ratio as before, this nozzle will now be over-expanded and the normal shock will actually occur inside the nozzle (situation 4). And so we will need to increase the chamber pressure to move the shock back to the nozzle exit and return to situation 5. So to summarise, if our goal is to make the strongest shock possible, we want a huge expansion ratio nozzle and a (quite high) pressure ratio between the chamber pressure and environment pressure that allows us to achieve situation 5. aerospaceengineeringblog.com/wp-content/uploads/2016/06/NozzleConDiv.jpg This was a longer message than I expected! I hope it is helpful.
Awesome explanation, especially the "supersonic particles interact with particles of the same speed" and normal/diagonal information sending. That makes everything much clearer. I have a question: In rocket engines we accelerate mass to high speed in order to get a high impulse from them. But where exactly does the impulse exchange between the accelerated particles and the nozzle occur. An as a bonus: How does this work in aerospike engines. Normal Nozzles are inefficient when they are not expanding enough, leaving a pressure difference between inside and outside that also "adds" to the thrust (F=deltap*Area) but less than if it were perfectly expanded (assuming bell with no weight blabla). And the outgoing gas stream is not perfectly straight further hurting efficiency or is that already accounted for? Anyways with aerospikes supposedly going around this problem, where exactly is the thrust exchanged on an aerospike, especially when it is in thinner atmosphere where the outgoing gas should pretty much look as "bad" as the traditional nozzle. (As in, expanding almost normal to the rocket axis as we know it from 1st stages in higher atmospheres) Does the thrust exchange work beyond the nozzle? Probably not because the flow is supersonic, so does it exchange "diagonally" from the last part of the nozzle?
Part 1 answer: Regarding the momentum exchange between the rocket and the propellant - it's kind of a chicken and egg problem. The typical text book approach is to do a boundary analysis - e.g. calculate the mass x velocity going into the chamber and out the nozzle and find the difference. This is an analytically simple approach which doesn't answer your question. Another way of looking at the problem which is more 'physical' - the interaction that provides the force which accelerates the rocket is the net gas pressure acting on the combustion chamber walls and nozzle surface. See the picture at this link: . With this logic, we just have to generate high pressures over a large surface area - the velocity of the gas is just a side-effect. In order to generate really high pressures inside the combustion chamber, there must be a restriction that stops too much gas flowing out - the throat in the case of a rocket engine. If the pressure difference between the inside of the combustion chamber and the outside environment is large enough (all rocket engines), the gas flow will be at Mach 1 in the throat (kind of a side-effect). At the throat, the pressure is normally still much higher than the atmosphere pressure, and so we add a nozzle to 'use up' the remaining excess pressure to create more thrust (side-effect is accelerated flow). The thrust coefficient for typical nozzles can be up to about 1.9, so we can almost double our thrust with a well designed nozzle. When the nozzle is overexpanded, the pressure inside the nozzle is less than outside, and so the pressure balance across the nozzle surface creates a 'drag force' pulling the rocket backwards. The opposite happens if the nozzle is under-expanded.
@@TheGravityAssistant Thx for the different ways of looking at the problem. The textbook approach always felt a bit iffy once i started looking closer. GG on your part :) Sadly YT does not show me Part 2 of your answer (weird bug).
@@uninteressant2196 The problem with the Part 2 of my answer was that I deleted it😂, because I realised I wasn't sure about something and wanted to think some more. I'm still not certain, because I have not worked with aerospikes, but I will try my best! This picture is good for understanding the performance difference between a real bell nozzle, a real aerospike, and an ideal bell nozzle: en.wikipedia.org/wiki/Aerospike_engine#/media/File:Nozzle_performance_comparison.svg An aerospike is never as efficient at a given operating point as an ideal (adjustable) bell nozzle, but it is more efficient over a range of operating points than a real bell nozzle. Again, with an aerospike, pressure and momentum can only be exchanged in the combustion chamber walls and the nozzle/spike walls. Now this next part is just my speculation, not necessarily true! My guess for why aerospikes are never as efficient as ideal bell nozzles is because the flow exiting the combustion chambers must have some angle, and so the portion of the thrust that comes from the pressure acting on the injector minus the throat area creates a force that is not exactly along the axis of the rocket. For a bell nozzle, this force is 'always' along the rocket axis. See this video: ruclips.net/video/-SGIiO1APig/видео.html Finally, about the effect of exhaust angle. We only care about the angle of the flow "exactly" as it is leaving the bottom of the nozzle / spike. Once it exits the bottom of the nozzle, it doesn't matter how many degrees the flow changes direction, the pressure acting on the walls inside the nozzle will be the same and so the thrust will be the same. The problem with flows that turn really far is that they can turn so far that they start hitting the rocket and causing lots of heating. I hope that all helps.
The biggest question I have is just “how do you reduce the pressure at nozzle exit?” Does the air coming out of the nozzle play any part in this? I can certainly understand a pressure controlled room or environment, but I don’t think that’s really what’s meant or else it’d be kind of useless.
What is important is the ratio between the pressure in the combustion chamber (high pressure) and the ambient (low) pressure at the exit. If you want to achieve supersonic flow at the exit, you can increase the high pressure or decrease the ambient pressure (or both). You're right that if we want a rocket nozzle to work at sea level, it doesn't make sense to try and lower the environment pressure - so first stage rocket engines usually have very high combustion chamber pressures to achieve a large pressure ratio. Conversely, 2nd and 3rd stage engines which operate in near-zero pressure environments can have much lower chamber pressures and still achieve supersonic flow / a high pressure ratio. Anything divided by almost zero is very big.
@@TheGravityAssistant I see, that makes intuitive sense in retrospect. If the starting total pressure is higher, to the point that the reduced static pressure is close to atmospheric conditions while holding onto a lot of dynamic pressure, it would just keep moving through similar conditions as though not much had changed. If the static pressure is similar but the temperature is much higher, does that mean the compressed gas is actually released in the exhaust at densities lower than STP? I’ll have to try to wrap my head around that… I’ve been trying to look into a means of numerically estimating the motion of airflow based on pressure differentials and figuring out how far an open air supersonic flow will continue traveling at supersonic speeds before reaching equilibrium with normal air, and I’m not quite sure what-in open air away from the turbine-stops a flow from forming normal shock outside of just powering through it with high pressure…? And how the duration of that air burst impacts this. Do you have any suggestions on where I should look to that? I’m not really sure how to mathematically define the behavior of a compressible thermodynamic gas to move from a region of high pressure to low pressure in the first place. >.
@@TheLordGojira To answer your first question, the ideal gas law PV = nRT can be rearranged to P = rho * R * T, where rho is density and R = R_universal / M_molecule. From that we can see that yes, gasses at higher temperature for a given pressure will have lower density. It also makes sense when you think about a candle or a camp-fire, their combustion is happening at 1 atmosphere of pressure, and the heated gasses float away because of their reduced density (and the effect of buoyancy). I'm not sure I totally understand your second question. I'm going answer the question 'how does a supersonic jet stream break down and dissipate', and hope that helps :). When the supersonic stream exits a nozzle, it is at the same static pressure as the atmosphere, but it is has a much high velocity than the ambient air, so when it moves past there will be a large shear force. This shear force acts to slow down the jet, (and accelerate the ambient air a little bit). Viscous shear and turbulence dissipates the supersonic jet's high dynamic pressure. Have a look at figure 12 and 13 of this paper: iopscience.iop.org/article/10.1088/1742-6596/1240/1/012019/pdf And look at the bottom of the exhaust stream from this photo: space.stackexchange.com/questions/29758/temperature-and-pressure-of-rocket-exhaust Similarly, the temperature will eventually equalise once the hot exhaust gases have radiated or convected their excess heat away to the rest of the atmosphere.
I have a question, u mentioned that pressure at the exit nozzle must match the atmospheric pressure. So how do you reduce the atmospheric pressure. Also the part abt shock was a bit confusing Looking forward to your reply
Hi, atmospheric pressure at the exit can be reduced by increasing altitude (e.g. a rocket or plane flying higher). Alternatively by conducting a rocket engine test in a vacuum chamber, or with a pressure reducing device such as an inducer. You could instead imagine that we increase the pressure of the fluids going into the engine, as it is the ratio between fluid total pressure and atmospheric pressure that matter. What was confusing about shocks, could you elaborate a bit more please?
12:58 since u said shock is an inefficient compression process, and since shock travels in opposite direction to the flow in the nozzle, how does it's increase help match the nozzle exit pressure to the atmospheric pressure p.s- what is supersonic expansion
@@TheGravityAssistant I'm guessing that Rat King is referring to the graph at 7:27 looking like a tanh(x) curve, and the graph at 10:38 looking like a cosh(x) curve, and wondering why those two functions describe the situation..
BEST FUCKING VIDEO EVER, SAVED MY PROPULSION MID TERMS
Hello Lyle, kudos for this interesting and very well done video. Keep it up!!
I'm a freshman in mechanical engineering, looking to go into propulsion, and even though I have heard some of the beginning stuff and some of the later stuff went over my head, this was great! I am absolutely sure I will be coming back to this in the future as a visual aid. Keep up the good work man!
Oh and idk if im in the place to suggest things, or if this is your area of expertise, but I like reading about high speed, high altitude flight and have no clue how compression lift works, so if you know where I can go to find a more straighforward explaination, I'd welcome it with open arms.
@@jonahgimmi1528 Hi thanks for the kind words.
Regarding compression lift, the key thing to understand is oblique shocks. I would recommend Chapter 4 of Modern Compressible Flow by Anderson. You can skip over most of the maths and just read the text + study the images to get a basic understanding. Also this website has a reasonable illustration: www.aerodynamics4students.com/introduction/images-of-lift.php.
The important thing to note is that the oblique shocks generate regions of high pressure, but oblique shocks only occur if we are turning/deflecting a supersonic flow. So if we have an aircraft which deflects the flow underneath (high pressure) and doesn't deflect the flow on the top (low pressure), we get a large pressure difference (lift). You can imagine a wedge with the top surface parallel to the direction of flight.
I'm planning to make a video in the next few months about waveriders, which are a special kind of hypersonic vehicle, and that video will definitely cover the basics of compression lift.
@@jonahgimmi1528 Hi, I've just finished a video on the basics of compression lift, I hope this can be of some help to you! ruclips.net/video/pFb8GIEkMVw/видео.html
Congratulations Master for your beautiful presentation and interesting manter! I learnt so much! Thank's a lot and my best regards! Jacareí-Sao Paulo- Brazil.
Amazing!!😮😮😮
Oh man thank you for making me understand about CD nozzles in compressible flow case and ..Most importantly, Pe and PA variations.
Glad the video was helpful! Thanks for watching
It would be really helpful if you can discuss the method of characteristics part.
Thanks for watching. I'll do my best to make a video on the MoC in the near future, but I can't promise anything specific! It's a complicated topic to try and summarise.
@@TheGravityAssistant Ya, it is complicated. Do you have any good resources for now?
@@dhrumil979 I would suggest Modern Compressible Flow by Anderson. In Chapter 7 he introduces the Method of Characteristics for unsteady 1D flows like shock waves and expansion fans in tubes. I would suggest reading this chapter very well first.
The important point to understand is that with unsteady flows we are 'following' points of constant properties as they travel through 1D space and time. On the x-t diagram we're also studying how the speed of sound (and thus temperature) changes when shock waves and expansion fans interact with each other.
Once you understand that, it is much easier to understand the 2D method of characteristics, because all we have done in the 2D method of characteristics is replaced the time axis with the y direction, and assumed constant flow (so no variations in time, only with location in 2D space).
In the later chapters Anderson discusses the MoC for 2D and numerical methods etc.
Hi! I'm currently designing a Nozzle for a CubeSat, I used MOC to obtain the design of the wall. I then passed it to ANSYS and i have an underexpansion effect which was expected if my ambient pressure is 0. So, I want to know how much I can expand the exit area to get the pressures to balance. I want to model this just like your video where I can have two charts and see the change in Mach and Pressure changing the velocity and exit area, what program do you use or what can you recommend for this. Thank you!
Hi, thanks for watching. I hope that I am understanding your first question correctly - it is not possible/practical in reality to match the exit static pressure of the nozzle with the vacuum pressure (~0) in space. It is possible to calculate a maximum theoretical exit velocity from a nozzle where all the thermal energy is converted to kinetic energy and the exit temperature is 0 K (and therefore the exit static pressure is 0), but this would require a very long (and heavy) nozzle with a very big expansion ratio. Additionally a real gas would condense into liquid or solid at a temperature >0 K and so the maximum ideal gas expansion could not be achieved. Additionally + 1, viscous losses with the long nozzle wall would remove some amount of energy from the flow, further preventing us reaching this theoretical idea. In practice, all space nozzles are under-expanded, and the designer has to make a trade-off between factors such as performance (Isp), mass, packaging (length and diameter) etc. etc.
Regarding the plots of Mach / pressure / temperature - I made them in matlab using the geometry of a simple converging-diverging nozzle, and the isentropic flow relations (please see the link below). If you know the cross section area at each position along the nozzle, you can use equation 9 (the Mach-area relation) to calculate the Mach number at each position. Once you know the Mach number at each position along the nozzle, you can use equations 6 and 7 to calculate the temperature and pressure at each location.
www.grc.nasa.gov/www/k-12/airplane/isentrop.html
Hope that helps!
Great explanation with very explanatory graphics. Many thanks. A question though: what are design parameters for a nozzle to create certain effects? E.G., if I wanted te create maximum shockwaves, regardless wat happens to the airspeed at the exit, what would that mean for the design of the nozzle?
Thanks for watching and the kind words. That's an interesting question and I had to take some time to think about it.
On a fundamental level, the most important variable that we can control is the pressure ratio between the total pressure in the chamber and the pressure of the environment at the exit. Please see the image I linked at the end of this message.
This following paragraph assumes that we have a constant nozzle geometry, like shown in the linked image. At point 7 we have perfectly expanded flow at the exit and no shocks. As we move upwards to point 6, the environment pressure is a little bit higher than the final nozzle pressure, and so some 'weak' oblique shocks are needed to match the exit pressure to the environment pressure. As we increase the environment pressure more, the weak oblique shocks eventually turn into a 'strong' normal shock at point 5. This shock occurring at point 5 is the 'strongest' shock that a converging-diverging nozzle (of a given geometry) can create. If we increase the environment pressure more, the normal shock has to move into the nozzle and it gets weaker because the subsonic expansion now contributes to some of the pressure recovery.
If we want to create a stronger normal shock at the nozzle exit, we need to increase the mach number of the flow at the nozzle exit, and so we need a bigger nozzle exit. However, with the same chamber-to-environment pressure ratio as before, this nozzle will now be over-expanded and the normal shock will actually occur inside the nozzle (situation 4). And so we will need to increase the chamber pressure to move the shock back to the nozzle exit and return to situation 5.
So to summarise, if our goal is to make the strongest shock possible, we want a huge expansion ratio nozzle and a (quite high) pressure ratio between the chamber pressure and environment pressure that allows us to achieve situation 5.
aerospaceengineeringblog.com/wp-content/uploads/2016/06/NozzleConDiv.jpg
This was a longer message than I expected! I hope it is helpful.
Thank you so much for your elaborated answer. Is there I way I can email you directly?
@@ariebos7872 No worries. I can be reached at lyle.tac@gmail.com, and I will reply when I can
Awesome explanation, especially the "supersonic particles interact with particles of the same speed" and normal/diagonal information sending. That makes everything much clearer.
I have a question: In rocket engines we accelerate mass to high speed in order to get a high impulse from them. But where exactly does the impulse exchange between the accelerated particles and the nozzle occur.
An as a bonus: How does this work in aerospike engines.
Normal Nozzles are inefficient when they are not expanding enough, leaving a pressure difference between inside and outside that also "adds" to the thrust (F=deltap*Area) but less than if it were perfectly expanded (assuming bell with no weight blabla). And the outgoing gas stream is not perfectly straight further hurting efficiency or is that already accounted for?
Anyways with aerospikes supposedly going around this problem, where exactly is the thrust exchanged on an aerospike, especially when it is in thinner atmosphere where the outgoing gas should pretty much look as "bad" as the traditional nozzle. (As in, expanding almost normal to the rocket axis as we know it from 1st stages in higher atmospheres)
Does the thrust exchange work beyond the nozzle? Probably not because the flow is supersonic, so does it exchange "diagonally" from the last part of the nozzle?
Hi, I'm very sorry for the slow reply, somehow I missed this comment. I will give a detailed answer to your questions tonight. Thanks for watching!
Part 1 answer:
Regarding the momentum exchange between the rocket and the propellant - it's kind of a chicken and egg problem. The typical text book approach is to do a boundary analysis - e.g. calculate the mass x velocity going into the chamber and out the nozzle and find the difference. This is an analytically simple approach which doesn't answer your question.
Another way of looking at the problem which is more 'physical' - the interaction that provides the force which accelerates the rocket is the net gas pressure acting on the combustion chamber walls and nozzle surface. See the picture at this link: . With this logic, we just have to generate high pressures over a large surface area - the velocity of the gas is just a side-effect.
In order to generate really high pressures inside the combustion chamber, there must be a restriction that stops too much gas flowing out - the throat in the case of a rocket engine. If the pressure difference between the inside of the combustion chamber and the outside environment is large enough (all rocket engines), the gas flow will be at Mach 1 in the throat (kind of a side-effect). At the throat, the pressure is normally still much higher than the atmosphere pressure, and so we add a nozzle to 'use up' the remaining excess pressure to create more thrust (side-effect is accelerated flow). The thrust coefficient for typical nozzles can be up to about 1.9, so we can almost double our thrust with a well designed nozzle.
When the nozzle is overexpanded, the pressure inside the nozzle is less than outside, and so the pressure balance across the nozzle surface creates a 'drag force' pulling the rocket backwards. The opposite happens if the nozzle is under-expanded.
@@TheGravityAssistant Thx for the different ways of looking at the problem. The textbook approach always felt a bit iffy once i started looking closer.
GG on your part :)
Sadly YT does not show me Part 2 of your answer (weird bug).
@@uninteressant2196 The problem with the Part 2 of my answer was that I deleted it😂, because I realised I wasn't sure about something and wanted to think some more. I'm still not certain, because I have not worked with aerospikes, but I will try my best!
This picture is good for understanding the performance difference between a real bell nozzle, a real aerospike, and an ideal bell nozzle: en.wikipedia.org/wiki/Aerospike_engine#/media/File:Nozzle_performance_comparison.svg
An aerospike is never as efficient at a given operating point as an ideal (adjustable) bell nozzle, but it is more efficient over a range of operating points than a real bell nozzle. Again, with an aerospike, pressure and momentum can only be exchanged in the combustion chamber walls and the nozzle/spike walls. Now this next part is just my speculation, not necessarily true! My guess for why aerospikes are never as efficient as ideal bell nozzles is because the flow exiting the combustion chambers must have some angle, and so the portion of the thrust that comes from the pressure acting on the injector minus the throat area creates a force that is not exactly along the axis of the rocket. For a bell nozzle, this force is 'always' along the rocket axis. See this video: ruclips.net/video/-SGIiO1APig/видео.html
Finally, about the effect of exhaust angle. We only care about the angle of the flow "exactly" as it is leaving the bottom of the nozzle / spike. Once it exits the bottom of the nozzle, it doesn't matter how many degrees the flow changes direction, the pressure acting on the walls inside the nozzle will be the same and so the thrust will be the same.
The problem with flows that turn really far is that they can turn so far that they start hitting the rocket and causing lots of heating.
I hope that all helps.
@@TheGravityAssistant Thank you for your honest answer regarding the limits of your expertise
Does RUclips also put cookies in my brain because I literally thought of this " ok the math says so but what's the reason for supersonic flow".
The biggest question I have is just “how do you reduce the pressure at nozzle exit?” Does the air coming out of the nozzle play any part in this? I can certainly understand a pressure controlled room or environment, but I don’t think that’s really what’s meant or else it’d be kind of useless.
What is important is the ratio between the pressure in the combustion chamber (high pressure) and the ambient (low) pressure at the exit. If you want to achieve supersonic flow at the exit, you can increase the high pressure or decrease the ambient pressure (or both).
You're right that if we want a rocket nozzle to work at sea level, it doesn't make sense to try and lower the environment pressure - so first stage rocket engines usually have very high combustion chamber pressures to achieve a large pressure ratio.
Conversely, 2nd and 3rd stage engines which operate in near-zero pressure environments can have much lower chamber pressures and still achieve supersonic flow / a high pressure ratio. Anything divided by almost zero is very big.
@@TheGravityAssistant I see, that makes intuitive sense in retrospect. If the starting total pressure is higher, to the point that the reduced static pressure is close to atmospheric conditions while holding onto a lot of dynamic pressure, it would just keep moving through similar conditions as though not much had changed.
If the static pressure is similar but the temperature is much higher, does that mean the compressed gas is actually released in the exhaust at densities lower than STP? I’ll have to try to wrap my head around that…
I’ve been trying to look into a means of numerically estimating the motion of airflow based on pressure differentials and figuring out how far an open air supersonic flow will continue traveling at supersonic speeds before reaching equilibrium with normal air, and I’m not quite sure what-in open air away from the turbine-stops a flow from forming normal shock outside of just powering through it with high pressure…? And how the duration of that air burst impacts this. Do you have any suggestions on where I should look to that? I’m not really sure how to mathematically define the behavior of a compressible thermodynamic gas to move from a region of high pressure to low pressure in the first place. >.
@@TheLordGojira To answer your first question, the ideal gas law PV = nRT can be rearranged to P = rho * R * T, where rho is density and R = R_universal / M_molecule. From that we can see that yes, gasses at higher temperature for a given pressure will have lower density. It also makes sense when you think about a candle or a camp-fire, their combustion is happening at 1 atmosphere of pressure, and the heated gasses float away because of their reduced density (and the effect of buoyancy).
I'm not sure I totally understand your second question. I'm going answer the question 'how does a supersonic jet stream break down and dissipate', and hope that helps :).
When the supersonic stream exits a nozzle, it is at the same static pressure as the atmosphere, but it is has a much high velocity than the ambient air, so when it moves past there will be a large shear force. This shear force acts to slow down the jet, (and accelerate the ambient air a little bit). Viscous shear and turbulence dissipates the supersonic jet's high dynamic pressure.
Have a look at figure 12 and 13 of this paper: iopscience.iop.org/article/10.1088/1742-6596/1240/1/012019/pdf
And look at the bottom of the exhaust stream from this photo: space.stackexchange.com/questions/29758/temperature-and-pressure-of-rocket-exhaust
Similarly, the temperature will eventually equalise once the hot exhaust gases have radiated or convected their excess heat away to the rest of the atmosphere.
Why this video has so little views?
I have a question, u mentioned that pressure at the exit nozzle must match the atmospheric pressure. So how do you reduce the atmospheric pressure. Also the part abt shock was a bit confusing
Looking forward to your reply
Hi, atmospheric pressure at the exit can be reduced by increasing altitude (e.g. a rocket or plane flying higher). Alternatively by conducting a rocket engine test in a vacuum chamber, or with a pressure reducing device such as an inducer.
You could instead imagine that we increase the pressure of the fluids going into the engine, as it is the ratio between fluid total pressure and atmospheric pressure that matter.
What was confusing about shocks, could you elaborate a bit more please?
12:58 since u said shock is an inefficient compression process, and since shock travels in opposite direction to the flow in the nozzle, how does it's increase help match the nozzle exit pressure to the atmospheric pressure
p.s- what is supersonic expansion
Apologies for the slow reply. Can you please restate your questions, I don't really understand what you're asking.
@@TheGravityAssistant how does increase in shock help exit pressure to match atmospheric pressure
"What happens next will SHOCK you"
...badum tss
tanh(x)for fluid supersonic.
cosh(x) for object driven in flow..... why?
I'm sorry, I don't understand your question. Could you please clarify what you mean?
@@TheGravityAssistant I'm guessing that Rat King is referring to the graph at 7:27 looking like a tanh(x) curve, and the graph at 10:38 looking like a cosh(x) curve, and wondering why those two functions describe the situation..
OMG THIS VIDEO SAVED ME IN HYPERSONIC AERODYNAMICS!!!!!!! Thanks from Brazil🟢🟡
Glad it was helpful!
Hi I left a comment on your follow up comment looking forward to ur response