SDF Aerospace and Aerodynamics Corner

MiG-29 said:

from one source
These necessary variations in mass airflow can be achieved only by altering the size of the intake throat; by making it, in the engineering term, a variable geometry intake rather than a fixed one.

Click to expand...

From the same :



Again, you are which is a fallacy. You deliberately remove materials from your citation to skew the author's viewpoint to match opinion, but the author clear attributed the change in mass flow to spill doors. The author even incorporated diagrams showing how excess air flow is from the bypass:





You made the same fallacy when you cite the following paragraph earlier. Again, you only quoted the part about change in throat area and mass flow rate, but deliberately removed the part about bypass doors being opened at supersonic speed to remove the excess flow:


If you are right, you wouldn't need to resort to misleading like this.

MiG-29 said:

from other source
TechLand's TBCC was designed to address the problem of the changing mass-flow requirements for a supersonic inlet operating across different flight conditions. At the design cruise speed, inlet throat area and mass flow can be matched to maximize engine efficiency, but "off-design" at transonic and low supersonic speeds the inlet needs a much larger throat area to meet the engine's demand for air. This requires an inlet that can accomodate a wide variation in throat area, otherwise engine efficiency suffers .

Click to expand...

To maximize efficiency, normal shock has to be positioned at the throat. Since the size of the normal shock is fixed for a given flow condition, the throat has to be adjusting to the same size as the normal shock, and this is what matching the inlet throat area to mass flow means.

However, you automatically assume change in throat area means change in mass flow, and this is . Repeating the same fallacy ad infinitium isn't going to magically make your false claims into reality. We know throat area does not influence mass flow because of the equation for mass flow ratio, which is independent of throat area:

MiG-29 said:

There is a maximum airflow limit that occurs when the Mach number is equal to one.
contrary to engineer`s fallacy

Click to expand...

The area of normal shock determines the mass flow, which is what it is meant by maximum air flow occurs at Mach 1. You claim I make fallacy, and yet the source you cite from says the exact same thing as I stated. This means you do not know fallacy is, and you should refrain from using a term that you do not understand.

Fallacy refers to improper reasoning used in an argument. When you try to put words in my mouth to misrepresent my position, that's which is a fallacy. When you take materials out of a citation to skew the authors' viewpoint, that's called , and is a fallacy.

My presentation of facts that do not conform to your bias opinions is however, not a fallacy. It only means your view is incorrect.

MiG-29 said:

Spillage is only increased by the engine but it already exists at lower flow ratios than 1
at the inlet design mach number or M=Mdesign near the max speed of the aircraft and where flow ratio is 1

Click to expand...

Spillage is not a result of variation of throat area. It is resulted because the air demand from the engine is reduced, causing a pressure built up which pushes the normal shock out of the inlet. It is explained along with critical and supercritical conditions in one of the sources that you have used:


When the variable-geometry inlet fixes the size of the throat, you assume that the mass flow cannot be increased or decreasedm but this assumption is wrong. When the throat size is too small, the normal shock there shifts downstream into the inlet and grows in area to accommodate the mass flow. This is why in the following graph, you see the line is slanted and not a straight vertical between the critical and supercritical point.


When the engine is throttled down, the air demand is reduced thus decreasing mass flow. The back pressure no longer able to keep the normal shock in the inlet, and the shock shifts upstream until it leaves the inlet's mouth. This creates a gap between the shock and the inlet where pressure from the inlet can escape, consequently reducing mass flow to the engine. This is shown by the near-horizontal line in the above graph, and it happens without any variation in throat geometry.

MiG-29 said:

below the　intake design mach number or M<Mdesign anf flow ratios lower than 1

Click to expand...

Spillage without variation in throat area:


From the above equation, we can see that the amount of spill is determined by two variables only. The first is physical capture area Ac, the other is free-stream cross-sectional area A0. Spill is independent of throat area At, thus throat size does not control mass flow, contrary to your claim.

Here is a video demonstrating spilling. There is no variation of throat geometry here:
[video=youtube;49sOQ89v-O8]http://www.youtube.com/watch?v=49sOQ89v-O8[/video]

Engineer said:

The area of normal shock determines the mass flow, which is what it is meant by maximum air flow occurs at Mach 1. You claim I make fallacy, and yet the source you cite from says the exact same thing as I stated. This means you do not know fallacy is, and you should refrain from using a term that you do not understand.
My presentation of facts that do not conform to your bias opinions is however, not a fallacy. It only means your view is incorrect.
When the variable-geometry inlet fixes the size of the throat, you assume that the mass flow cannot be increased or decreasedm but this assumption is wrong. .

.

Click to expand...

since you write too much fallacies mosty of them are not even worthed to answer.


These necessary variations in mass airflow can be achieved only by altering the size of the intake throat; by making it, in the engineering term, a variable geometry intake rather than a fixed one.



At take-off and in climb the ramps are in the fully open position, the spill door is closed and its inlet flap is open.


At the design cruise speed, inlet throat area and mass flow can be matched to maximize engine efficiency,


this is the way you think and your basic fallacy
The conservation of mass (continuity) tells us that the mass flow rate mdot through a tube is a constant and equal to the product of the density r, velocity V, and flow area A:

Eq #1:


mdot = r * V * A
Considering the mass flow rate equation, it appears that for a given area and a fixed density, we could increase the mass flow rate indefinitely by simply increasing the velocity.

but saddly air is spilled and according to you the flow after supersonic speeds should go faster than mach 1 to keep bernoulli`s principle, but you know what? it does not it won`t go faster therefore it spills it overflows, therefore you need the intake to get supersonic at the throat and intake cowl lip in supersonic external compression to get a mass flow ratio of 1

because There is a maximum airflow limit that occurs when the Mach number is equal to one. The limiting of the mass flow rate is called choking of the flow. If we substitute M = 1 into Eq #10 we can determine the value of the choked mass flow rate:
If the throat area (A2) is too small, corresponding to point c in Figure 2.25, a detached shock will stand ahead of the inlet, as shown in Figure 2.25a. By increasing the throat area, the shock can be moved to the inlet lip. When the operating point a of Figure 2.25 is reached, the shock will reach the inlet lip and ...
[

MiG-29 said:

since you write too much fallacies mosty of them are not even worthed to answer.

Click to expand...

You are unable to answer not because I write fallacies, but because what have I presented are facts, and your lies and fallacies are not working on me.

Also, it is clear you do not know what fallacy means. Covering your ears and accuse me of making fallacies doesn't make it so. Also, your inability to point out exactly which of my statement is a fallacy and what types of fallacy it is speaks volume. What I am doing is presenting facts and arguing with logic. Just because you find the truth uncomfortable and go into denial, that doesn't make my statements fallacies.

So, let me lecture you on what fallacy means. Fallacies refer to improper reasonings used in argument, such your use of where you intentionally remove materials from citations, and where you invent false statements then claim I made them. These are fallacies because they divert attention rather than address the point, and you are using every fallacy in the book.

MiG-29 said:

These necessary variations in mass airflow can be achieved only by altering the size of the intake throat; by making it, in the engineering term, a variable geometry intake rather than a fixed one.

Click to expand...

This is a fallacy called quoting out of context, because you deliberately removed materials from the paragraph to skew the author's viewpoint. The author's complete statement is as follow:


The change in mass flow is attributed to the bypass system and spill doors, not the variation of throat area. When extra air is needed, spill doors are opened to let air in. When mass flow needs to be reduced, spill doors are open outward to dump the excess air flow. The inlets on F-14 work in a similar way. Once again, from your own source:


This also says that excess air is removed by the bypass system. This air is removed at the throat, but the throat area has nothing to do with the change.

MiG-29 said:

At take-off and in climb the ramps are in the fully open position, the spill door is closed and its inlet flap is open.


At the design cruise speed, inlet throat area and mass flow can be matched to maximize engine efficiency,

Click to expand...

The mass flow is directly related to area of the normal shock, as explained in the following:



This area does not yield to change in throat area, because the shock simply shifts to maintain the area. For example, when the throat size becomes too small, the normal shock shifts until it fits an area further downstream, a condition known as supercritical condition. This allows the shock to take up bigger area, because the inlet duct there is wider, thus mass flow is conserved.


By matching the inlet throat area to mass flow, the throat is made to be the same size as normal shock so as to position the shock at the throat for maximum pressure recovery. This is one purpose of variable-geometry, but since the variable throat size cannot alter the area of normal shock, it meams variation in throat area cannot influence mass flow.

Finally, for the ramps, they are retracted at subsonic speed and deployed at supersonic speed. Their purpose is to produce shock, and this cannot be done at subsonic speed, which is why they are retracted. Their deployment at supersonic speed is obvious. However, just because the change in throat area and mass flow occur at the same time, that doesn't mean one influences another. Your claim otherwise is known as .

MiG-29 said:

this is the way you think and your basic fallacy
The conservation of mass (continuity) tells us that the mass flow rate mdot through a tube is a constant and equal to the product of the density r, velocity V, and flow area A:

Eq #1:


mdot = r * V * A
Considering the mass flow rate equation, it appears that for a given area and a fixed density, we could increase the mass flow rate indefinitely by simply increasing the velocity.

but saddly air is spilled and according to you the flow after supersonic speeds should go faster than mach 1 to keep bernoulli`s principle, but you know what? it does not it won`t go faster therefore it spills it overflows, therefore you need the intake to get supersonic at the throat and intake cowl lip in supersonic external compression to get a mass flow ratio of 1

because There is a maximum airflow limit that occurs when the Mach number is equal to one. The limiting of the mass flow rate is called choking of the flow. If we substitute M = 1 into Eq #10 we can determine the value of the choked mass flow rate:

Click to expand...

You are wrong in assuming that flow in an inlet is uniformly Mach 1.

Flow does not go faster than Mach 1 at normal shock, but that doesn't mean everywhere else the flow is at Mach 1. In fact, for normal shock to exists in an inlet, there must be supersonic flow (flow that is faster than Mach 1) upstream and subsonic flow (flow that is slower than Mach 1) downstream. Putting it in another way, the normal shock can only exist at one location, and anywhere else the flow is not Mach 1. Bernoulli's principle still applies. Your attempt at poking holes at this simple fact betrays your desperation at preaching your opinions, and shows your lack of understanding in the subject.

Now, because the flow upstream of the normal shock is supersonic, the signal that there is a restriction at the throat cannot make it out of the inlet. This is because that signal propagates at the speed-of-sound, which is slower than the supersonic flow. In short, the air inside the inlet cannot warn the air outside the inlet to move away, so the air outside would go into the inlet regardless. Then, no matter how your vary the throat size, it does not affect mass flow.

Spillage occurs behind a shock wave, usually the normal shock is referred. Spillage occurs when the normal shock is pushed out of the inlet's mouth due to pressure build-up, leaving a gap between the shock and the opening of the inlet where air can escape. This condition is a result subcritical condition. This has nothing to do with variation in throat area, as spillage occurs on both fixed inlet and DSI.

Here is a video demonstrating spilling. There is no variation of throat geometry here:
[video=youtube;49sOQ89v-O8]http://www.youtube.com/watch?v=49sOQ89v-O8[/video]

MiG-29 said:

If the throat area (A2) is too small, corresponding to point c in Figure 2.25, a detached shock will stand ahead of the inlet, as shown in Figure 2.25a. By increasing the throat area, the shock can be moved to the inlet lip. When the operating point a of Figure 2.25 is reached, the shock will reach the inlet lip and ...

Please, Log in or Register to view URLs content!

Click to expand...

This citation of your's show the process of starting an inlet, which does not prove your claim that variation in throat area is in control of mass flow. Furthermore, from that same , we see that flow ahead of the normal shock is supersonic and flow behind the normal shock is subsonic:


Note that in diagram (a) the flow at the throat is less than Mach 1 and in (c) the flow at the throat is greater than Mach 1. This verifies what I have said, and disprove your claim that flow cannot move faster than Mach 1 at the throat. The simple matter is that when the throat is too narrow, the normal shock simply shifts downstream to seek a bigger area to accommodate the flow. Thus, the throat area does not influence mass flow:


Finally, throat size does not influence mass flow ratio because the equation for this ratio is independent from throat area As:

Engineer said:

You are unable to answer not because I write fallacies, but because what have I presented are facts, and your lies and fallacies are not working on me.

Click to expand...

you write only fallacies and as long as you have a that attitude, you will continue saying fallacies

　　These necessary variations in mass airflow can be achieved only by altering the size of the intake throat; by making it, in the engineering term, a variable geometry intake rather than a fixed one.　　　

this is a fact, you are just saying fallacies, , no reasoning is possible with you

For typical HSCT-type engines the requiered cruise air flow can be as low as 70% of the required air flow at transonic conditions, To provide for the large airflow rates at transonic condition the inlet throat area must be larger than at the cruise condition, this increased area is provided by the variable inlet geometry


this graph shows a fixed intake, they show subcritical, critical, started supercritical state this prove you know nothing and it is pure posture and fallacies what you say since it is freestream mach number and the intake`s design mach number a major fallacy and blunt lie from you since the shock does not equate freestream mach number, they show supercritical state and the shock at different positions in the diffuser and free stream mach number does not equate mass flow ratio.

and what they show is the shock position at different freestream mach number speeds, first it is below the inlet design mach number, later almost at the design mach number and last above the inlet design mach number showing a started shock position and a too aft position of low pressure recovery

MiG-29 said:

you write only fallacies and as long as you have a that attitude, you will continue saying fallacies

Click to expand...

You have no idea what fallacy is. Fallacies refer to improper reasonings, my act of presenting facts and connecting them together is called logical arguments. Just because the facts I have presented don't conform to your opinions, that doesn't make my statements fallacies. It only means your views are incorrect. Deal with it.

In addition, just because you use fallacies, that doesn't mean other people use the same tactics. Your of your own actions and attributes on me only serve to speak more about you.

MiG-29 said:

　　These necessary variations in mass airflow can be achieved only by altering the size of the intake throat; by making it, in the engineering term, a variable geometry intake rather than a fixed one.　　　

this is a fact, you are just saying fallacies, , no reasoning is possible with you

Click to expand...

The one who is not here to reason is you. You have an axe to grind against the J-20, so you try to turn any feature on the aircraft into some form of disadvantage. This discussion on inlet is just an extension of that, because you want to argue DSI has worse performance than inlets on third generation aircraft. You claim variation of throat size can control mass flow as a premise for that argument, which is proved to be wrong by facts, but you are relentless in preaching your opinions. Preaching opinions is not reasoning.

The fact is, that paragraph attributes that change in mass flow to spill doors and bypass. When more flow is need, the door opens inward to allow air in. When less flow is need, the door opens outward to bleed excess air. Your deliberate removal of materials from that citation is a fallacy called , because you are skewing the author's viewpoint. I have took a screenshot of the entire paragraph and highlighted the parts that you have taken out:


The author even incorporated diagrams showing how variation to air flow is made by spill door and bypass:

MiG-29 said:

For typical HSCT-type engines the requiered cruise air flow can be as low as 70% of the required air flow at transonic conditions, To provide for the large airflow rates at transonic condition the inlet throat area must be larger than at the cruise condition, this increased area is provided by the variable inlet geometry

Click to expand...

This is , and repeating this ad infinitium won't make the two events occurring together to imply one is in control of another -- . Using large font size isn't going to change the fact that throat area does not influence mass flow either.

The collapsed position of the ramp is the default position, and it is the position the ramp is in when no shock is created. Once the aircraft goes supersonic, the ramp moves downward to create oblique shocks, and the resulting throat creates the normal shock. This is called creating a throat when it is needed, and putting it away when it is not, which is not controlling mass flow rate as you have claimed.

The difference in mass flow between subsonic and supersonic speed is accounted for bypass, as explained in multiple sources. For example, in regard to F-14's inlets one of your very own sources say the following:


Then it continues:


In another source of your's, it difference in mass flow between the air that goes into the inlet and the air that goes into the engine is accounted for by the bypass system:


Furthermore, yet another source of your's show that excess air must be removed by the bypass system. If throat area has ability to restrict air flow as you claimed, then this wouldn't be needed:


The fact that throat area does not influence mass flow is also confirmed by the equation below, which shows mass flow ratio is dependent on two variables. They are the physical capture area A1 and free-stream area A0i. It has no involvement of throat area As:


Why don't you talk about this equation? Because you like to spin, and you cannot put a spin on equation like you can do with people's words. But your denial of the existence of this equation doesn't mean the equation is not there. Throat size does not affect the mass flow rate, that's one of the things says by the above equation. As simple as that.

MiG-29 said:

this graph shows a fixed intake, they show subcritical, critical, started supercritical state this prove you know nothing and it is pure posture and fallacies what you say since it is freestream mach number and the intake`s design mach number a major fallacy and blunt lie from you since the shock does not equate freestream mach number, they show supercritical state and the shock at different positions in the diffuser and free stream mach number does not equate mass flow ratio.

and what they show is the shock position at different freestream mach number speeds, first it is below the inlet design mach number, later almost at the design mach number and last above the inlet design mach number showing a started shock position and a too aft position of low pressure recovery

Click to expand...

You keep accusing me of using fallacies, but you are unable to provide an example. Not a single one. This speaks volume. Your incoherent sentences also show that you are desperately trying to continue to argue despite facts have already shown you to be incorrect. This says you are here to preach your opinion, and argue for the sake of arguing.

This image is from your very own source, and shows your claim that air flow at throat cannot be greater than Mach 1 to be wrong. The fact that you have no idea this could happen shows you have no understanding in the subject in which you are arguing.


Thus the limiting factor on flow is the area of the normal shock. In a variable-geometry inlet, when the throat varies in size the normal shock simply shifts, a fact admitted by yourself. This shifts occur because the normal shock conserves its area. When throat size is too narrow, the normal shock moves further into the inlet to find a bigger area, resulting in supercritical condition. In other words, the normal shock shifts to accommodate flow regardless of throat size.


This is known as supercritical condition, and it gives the engine the correct mass flow despite the throat being too narrow. Thus, variation in throat size does not affect the area of the normal shock, consequently does not affect the mass flow.

Engineer said:

This is known as supercritical condition, and it gives the engine the correct mass flow despite the throat being too narrow. Thus, variation in throat size does not affect the area of the normal shock, consequently does not affect the mass flow.

Click to expand...

you just say fallacies, and blunt lies, the figure you presented showed only free stream mach number, mach design number, critical, subcritical and supercritical conditions with respect the normal shock on an internal compression fixed inlet like this


the rest is a bunch of fallacies, because you can not comprehend what is critical positon with respect mach design number and mass flow ratio, this figure shows you that at the mach design number the critical condition exist


and it is fact
These necessary variations in mass airflow can be achieved only by altering the size of the intake throat; by making it, in the engineering term, a variable geometry intake rather than a fixed one. fallacies will try to say lies about a fact as you do

MiG-29 said:

you just say fallacies, and blunt lies

Click to expand...

Covering your ears and accusing me of saying fallacies doesn't make it so. What I have presented are facts, facts that you are unable to accept, which doesn't make my statement fallacies either. Deal with it.

MiG-29 said:

the figure you presented showed only free stream mach number, mach design number, critical, subcritical and supercritical conditions with respect the normal shock on an internal compression fixed inlet like this

Click to expand...

The figure above shows that air flow at the throat can have a speed lower than Mach 1, as well as a speed higher than Mach 1. This is indicated by the small circle, pointed at by the arrow with the flow velocity written right beside. You claimed that flow velocity cannot be higher than Mach 1 at the throat, and the diagrams showed you to be incorrect. You made the erroneous claim because you are lacking knowledge in the subject, and you want to appear smart by disagreeing with everything that I have said. However, the fact is there for everybody to see, and that is normal shock (flow at Mach 1) can shift location to vary its area, which enables the shock to accommodate the mass flow regardless of throat size.

MiG-29 said:

the rest is a bunch of fallacies, because you can not comprehend what is critical positon with respect mach design number and mass flow ratio, this figure shows you that at the mach design number the critical condition exist

Click to expand...

It is funny how you keep accusing me of making fallacies when all I do is presenting facts. Obviously, by accusing my facts as fallacies, you are also accusing all those authors of those books and papers that we both cited from as making fallacies. Clearly, that cannot be the case, and the only explanation left that you are unable to accept facts when they don't conform to your opinion, and you are the only who made fallacies to divert attention.

The three conditions shown in the diagram how normal shock is positioned with respect to the throat. Subcritical has the normal shock upstream of the throat. Critical has the shock at the throat, and supercritical has the shock downstream of the throat.


Normal shock means flow velocity is at Mach 1. In an inlet, flow upstream of this shock is supersonic, while flow downstream the shock is subsonic. In other words, the three conditions also refer to flow velocity at the throat. So:

• Subcritical condition = subsonic flow at throat
• Critical condition = sonic flow at throat
• Supercritical condition = supersonic flow at throat

Thus, the term supercritical condition means that air at the throat can be higher than Mach 1, proving your claim as incorrect.

Here is another graph proving you are wrong. It shows that mass flow varies throughout subcritical, critical, and supercritical conditions independent of throat size. As the normal shock shifts down the inlet, its area increases because that part of the inlet is divergent. As a result, mass flow increases, which explains why the line representing supercritical condition is not vertical:


This shows that mass flow is related to area of the normal shock, not throat area.

The graph also shows that maximum pressure recovery is attained when the normal shock is positioned at the throat. Since the area of the normal shock does not change, the only way to position the normal shock at the throat is by making the throat the same size as the area of the normal shock. This is what it means by matching the inlet throat area to mass flow, because the area of the normal shock represents the mass flow. This carries the connotation that the mass flow controls throat area, not the other way around as you have claimed.

But no matter how the throat size is varied, it doesn't affect the area of normal shock, which means variation in throat area cannot influence mass flow. This is proven by the following equation, showing mass flow ratio is independent to throat area As:

Engineer said:

Covering your ears and accusing me of saying fallacies doesn't make it so. What I have presented are facts, facts that you are unable to accept, which doesn't make my statement fallacies either. Deal with it.


The figure above shows that air flow at the throat can have a speed lower than Mach 1, as well as a speed higher than Mach 1.

Click to expand...

pure fallacies, mach 1 is the speed of sound, the shock can not go beyond mach 1, you are saying fallacies, behind the shock the speed is subsonic that is the reason it spills in subcritical condition if it does not go subsonic then why using shockwaves for slowing down the speed of the air entering the diffuser?.

Freestream is achieveable by the airplane it self the airplane can go mach 2, the shock can not, it always goes mach 1.

at supercritical conditions in an internal compression intake there are oblique and normal shocks.


but what can expect from you? yeah i know just fallacies
you say so much fallacies that you did not even check on the fig 10.15 they show capture area equals the air that enters the intake as such the flow ratio is one at the intake`s design mach number


in the other graph the slope mc0 is the engine corrected mass rate, and it shows the engine demands with respect the mass flow, it shows when the engine demands are higher than the critical operation and as the mass flow rate remains constant, in few words it shows the engine operating regimes of an external compression intake

MiG-29 said:

pure fallacies, mach 1 is the speed of sound, the shock can not go beyond mach 1, you are saying fallacies, behind the shock the speed is subsonic that is the reason it spills in subcritical condition if it does not go subsonic then why using shockwaves for slowing down the speed of the air entering the diffuser?.

Freestream is achieveable by the airplane it self the airplane can go mach 2, the shock can not, it always goes mach 1.

at supercritical conditions in an internal compression intake there are oblique and normal shocks.

Click to expand...

Nope. My statements are pure facts, not fallacies. Just because the facts I have presented do not conform to your opinion, that doesn't make them fallacies, and covering your ears and making empty accusations won't hide the fact that you are the one who employs fallacies.

Your accusations also show that you have no idea what a fallacy is. Fallacy refers to improper reasonings used in an argument, such as you are displaying in your above statement, because you are trying to use irrelevant facts to divert attention away from your incorrect claim that flow at the throat cannot be higher than Mach 1.

The flow velocity at the normal shock is Mach 1. Elsewhere, it is not. In fact, for there to be a normal shock in an inlet, the upstream of the shock flow has to be supersonic and downstream of the shock has to be subsonic. The flow velocity at the throat can be below Mach 1, at Mach 1, or above Mach 1:

• Subcritical condition = subsonic flow at throat
• Critical condition = sonic flow at throat
• Supercritical condition - flow at throat is supersonic

Since the flow velocity ahead of the shock is supersonic, it means that signal of a restriction at the throat won't make it out of the inlet, as this signal propagates at the speed-of-sound. In kiddies' term, the air inside the inlet cannot warn the air outside of the inlet to move away, resulting in the same mass flow regardless of how throat area is varied.

The area of the normal shock determines the mass flow, but the shock can shift downstream in response to narrowing of the throat resulting in supercritical condition. When this occurs, it means the flow at the throat is higher than Mach 1. Thus, your claim that supersonic flow cannot occur at the throat is incorrect. This also means that when the throat narrows, the flow there just moves faster to compensate, thus mass flow is unaffected.

MiG-29 said:

but what can expect from you? yeah i know just fallacies
you say so much fallacies that you did not even check on the fig 10.15 they show capture area equals the air that enters the intake as such the flow ratio is one at the intake`s design mach number

Click to expand...

Wrong. I say so much facts, not fallacies. Just because your own fallacies are not working on me, that doesn't make my statements fallacies.

The fact remains that flow velocity at the normal shock is Mach 1, but ahead of the shock the flow is supersonic. When the throat area is too narrow, and the normal shock moves deeper into the inlet to seek a larger area. This causes the flow velocity at the throat to be supersonic and is your so called supercritical condition. Variation of throat area simply causes the normal shock to shift but has no affect on the area of the normal shock, thus no effect is made to mass flow.

MiG-29 said:

in the other graph the slope mc0 is the engine corrected mass rate, and it shows the engine demands with respect the mass flow, it shows when the engine demands are higher than the critical operation and as the mass flow rate remains constant, in few words it shows the engine operating regimes of an external compression intake

Click to expand...

In a few words, both graphs show that mass flow can be changed regardless of throat size, proving your claim that throat area being in control of mass flow is incorrect:



Beyond critical condition, the mass flow slightly increases because that part of the curve isn't a straight vertical. This is a result of the normal shock shifting downward into the part of the inlet that is divergent, allowing the normal shock to seek a larger area to accommodate for the mass flow:


The horizontal part of the curve shows what happens as the engine throttle decreases. The back pressure builds up and pushes the normal shock out of the inlet, resulting in a gap between the shock and the inlet's mouth. This gap allows the air to rapidly escape, resulting in a fast decrease in mass flow, causing the curve to be close to horizontal.


This graphs show that you are incorrect because your assumption won't allow the mass flow to change, whereas in the real word, mass flow can be changed regardless of throat size.

Engineer said:

yes My statements are pure fallacies.

Click to expand...

I agree you say only fallacies, you only say fallacy after fallacy, that is right

Note:

The limiting mass flow rate is directly proportional to the stagnation pressure, and inversely proportional to the square root of the stagnation temperature. Thus, if more work is taken out of the flow by the turbine per unit mass, the stagnation pressure drops sharply, and the mass flow and thus the thrust drop sharply as well.



The effects of inlet design and engine operation on total pressure recovery are expressed in an inlet performance plot called a recovery cane. The figure shows inlet recovery for a supersonic external compression inlet plotted versus mass flow ratio at a single free stream Mach number. The mass flow ratio is the engine mass flow rate divided by the maximum mass flow rate that can be captured by the inlet. On the schematic at the bottom of the upper figure, we show a typical capture streamline as the thin horizontal line upstream of the inlet lip. The recovery is always less than 1.0 because of shock losses and boundary layer losses. The knee of the curve is labeled critical and at this maximum mass flow ratio, the terminal normal shock sits at the cowl lip. For the lower mass flow ratios, labeled sub Critical・ the normal shock sits off of the cowl lip and excess mass is spilled around the cowl. At the lowest mass flow ratio the inlet is in buzz, an unsteady condition in which the normal moves at high frequency in the streamwise direction. For the Super Critical・portion of the curve, the normal shock is pulled inside the cowl. Super critical mass flow ratio remains a constant because the mass flow through the cowl lip plane is fixed by supersonic conditions . The recovery decreases along the curve because the normal shock is pulled farther back into the diffuser, with an increase in Mach number upstream of the normal shock, and a resulting increase in total pressure loss. Notice that the mass flow ratio never equals 1.0 because there is always some amount of supersonic spillage from the external oblique shocks. During inlet wind tunnel testing, recovery canes are generated over a range of free stream Mach numbers.