Supercruising turbofan engines

chuck4

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For a given airframe, what is the difference between an engine that can supercruise and one that can't? Is it purely a difference of dry thrust? Or would the engines exhibit other differences as well.




Does f-35 lack super cruising capability simply because the airframe is too draggy, and chosen engine can not produce the necessary dry thrust;, or does the engine in fact possess adaquate potential dry thrust for super cruise, but for some other reason is not tuned to support super cruising?
 
To supercruise, an engine must have the lowest bypass ratio possible to keep the exhaust jet's velocity high; a turbojet is preferable to a turbofan. After all, a jet can fly only as fast as the velocity of its exhaust jet and the bypass air slows the exhaust jet's velocity down.

As for the F-35, the F135 is a high bypass ratio engine optimized for low-altitude low-speed operation, to generate the maximum dry thrust at the lading time for the F-35B.
 
SlowMan said:
To supercruise, an engine must have the lowest bypass ratio possible to keep the exhaust jet's velocity high; a turbojet is preferable to a turbofan. After all, a jet can fly only as fast as the velocity of its exhaust jet and the bypass air slows the exhaust jet's velocity down.

As for the F-35, the F135 is a high bypass ratio engine optimized for low-altitude low-speed operation, to generate the maximum dry thrust at the lading time for the F-35B.

Since F-35A and F-35C have no STVOL requirement, why do they not go for a lower bypass variant of the engine?

Would an F-119 engine installed in F-35 be able to provide supercrusing capability?


BTW, "a jet can fly only as fast as the velocity of its exhaust jet" is not technically correct for physics considerations. If a jet is able to eject more than half of its own initial mass as exhaust out the back, then it should be able to be designed to go faster than the speed of its own exhaust. In other words, a jet with fuel fraction > 0.5 can go faster than its own exhaust if that is what you are trying to design it to do.


But that is a theoretical digression.
 
SlowMan said:
As for the F-35, the F135 is a high bypass ratio engine optimized for low-altitude low-speed operation, to generate the maximum dry thrust at the lading time for the F-35B.
Engine Characteristics
Conventional Take Off and Landing CTOL / CV Engine Design

Bypass Ratio: 0.57


Short Take Off and Vertical Landing STOVL Propulsion System Design

Bypass Ratio
Conventional: 0.56
Powered Lift: 0.51
 
chuck4 said:
Since F-35A and F-35C have no STVOL requirement, why do they not go for a lower bypass variant of the engine?
A lower bypass engine would require a new larger core, which defeats the purpose of sharing the same engine across three variants.

Would an F-119 engine installed in F-35 be able to provide supercrusing capability?
An F-35 is as draggy as an F-22, so no.

BTW, "a jet can fly only as fast as the velocity of its exhaust jet" is not technically correct for physics considerations.
You must have slept through in your high school physics class.

then it should be able to be designed to go faster than the speed of its own exhaust.
And violate Newton's Third Law of Motion.
 
SlowMan said:
A lower bypass engine would require a new larger core, which defeats the purpose of sharing the same engine across three variants.?


How so? F-135 already share a core with f-119.


SlowMan said:
An F-35 is as draggy as an F-22, ?


f-22 can super cruise based on similar T/w ratio as f-35.

SlowMan said:
And violate Newton's Third Law of Motion.
Newton's third law says nothing about the speed achieved by two masses thrusting against each other, only the forces acting on each is equal in magnitude and opposite in direction.


Combine newton's second and third law yield the result that the plane could end up traveling faster relative to the initial reference frame than the propellant it ejected if the planet's mass is less than the mass of the resultant of the propellant's combustion.


This result is otherwise known as conservation of momentum.
 
what about pressure inside the engine? Do afterburners increase the pressure aft of the engine core? If so, one could conclude that power is made with added pressure.


Now since there is no afterburner in supercruise, all that added pressure must come within the engine core itself. bypass portion of thrust starts to rapidly drop in transsonic region, so for supercruise one can rely only on non-bypassed thrust, meaning the one coming from the engine core. Older engines couldn't really withstand high pressure so we didn't really have supercruise before. Nowadays, with newer engines supercruise is more prolific.


alternative would be an oversized, overpowered engine for a given aiframe, even if built with older tech, like the pegasus engines on the concorde. But i guess that is not very efficient to use on relatively small planes like the fighters today.


Is all this that i've written bollocks or is there some truth to it?
 
Actually supercruise is not new. English Electric Lightning from the 1950s could do it. The Lightning didn't really have "oversized" engines. It's T/W ratio would be marginal next to F-35's.
 
One thing I noticed while handling the model of F-35 and F-22 is F-22 seems to try harder to conform to Whitcomb's area rule. F-35 had undercarriage bulges next to the otherwise tubular mid-fuselage precisely where the fuselage needs to slim down to accommodate the sectional area of the wing.
 
chuck4 said:
Actually supercruise is not new. English Electric Lightning from the 1950s could do it. The Lightning didn't really have "oversized" engines. It's T/W ratio would be marginal next to F-35's.

What the Lightning did was nothing special. Lots of fighters can do Mach 1.01 clean and dry. Not so many can do Mach 1.7+ dry with payload.
 
The P1A research aircraft that led to the Lightning exceeded Mach 1 comfortably on unreheated Sapphire engines. Max speed was at least Mach 1.22 - I'll have to check sources.
 
A fact to be borne in mind re. the English-Electric Lightning's supercruise capability (and also that of the Fairey FD.2, while we're at it) is that the Sapphire and Avon were both turbojets, not turbofans.
 
F-22 supercruises with turbofan engines. I believe Tu-144 could supercruise at Mach 2 with turbofan engines.

It would be interesting to find a plot of the actual engine thrust vs air speed for different aircraft / engine combinations, and a plot of drag vs air speed for the same aircrafts.
 
Concorde can supercruise between Mach 1.4-1.7, but needs reheat to get to M1.4.
 
I believe Concord can get all the way to M1.7 on dry thrust alone, but elects to use reheat to get through the transonic region faster. Transonic region is inefficient for the Concord, so it actually save fuel to use after burners to power through the transonic region the least amount of time.
 
chuck4 said:
Actually supercruise is not new. English Electric Lightning from the 1950s could do it. The Lightning didn't really have "oversized" engines. It's T/W ratio would be marginal next to F-35's.

What the Lightning did was nothing special. Lots of fighters can do Mach 1.01 clean and dry. Not so many can do Mach 1.7+ dry with payload.
lightning f6 do M1.4 dry.
 
chuck4 said:
Actually supercruise is not new. English Electric Lightning from the 1950s could do it. The Lightning didn't really have "oversized" engines. It's T/W ratio would be marginal next to F-35's.

What the Lightning did was nothing special. Lots of fighters can do Mach 1.01 clean and dry. Not so many can do Mach 1.7+ dry with payload.
lightning f6 do M1.4 dry.

Source?
 
I’ve seen M1.01 quoted for the Lightning P1 prototype - and that was probably the cleanest Lightning.
 
chuck4 said:
Actually supercruise is not new. English Electric Lightning from the 1950s could do it. The Lightning didn't really have "oversized" engines. It's T/W ratio would be marginal next to F-35's.

What the Lightning did was nothing special. Lots of fighters can do Mach 1.01 clean and dry. Not so many can do Mach 1.7+ dry with payload.
lightning f6 do M1.4 dry.

Source?
5d1ca59c96feb95027.jpg

Although very limited condition.
(the image above may need time to load. plz wait)
 
chuck4 said:
Actually supercruise is not new. English Electric Lightning from the 1950s could do it. The Lightning didn't really have "oversized" engines. It's T/W ratio would be marginal next to F-35's.

What the Lightning did was nothing special. Lots of fighters can do Mach 1.01 clean and dry. Not so many can do Mach 1.7+ dry with payload.
lightning f6 do M1.4 dry.

Source?

And M1.2 dry with 2 RedTop missiles.

5d1d6ac00fe0781992.jpg

5d1d6ac018ebe75903.jpg
 
Which document are these pages from?
Where did you find the document?
 
SlowMan said:
To supercruise, an engine must have the lowest bypass ratio possible to keep the exhaust jet's velocity high; a turbojet is preferable to a turbofan. After all, a jet can fly only as fast as the velocity of its exhaust jet and the bypass air slows the exhaust jet's velocity down.

As for the F-35, the F135 is a high bypass ratio engine optimized for low-altitude low-speed operation, to generate the maximum dry thrust at the lading time for the F-35B.

Since F-35A and F-35C have no STVOL requirement, why do they not go for a lower bypass variant of the engine?

Would an F-119 engine installed in F-35 be able to provide supercrusing capability?


BTW, "a jet can fly only as fast as the velocity of its exhaust jet" is not technically correct for physics considerations. If a jet is able to eject more than half of its own initial mass as exhaust out the back, then it should be able to be designed to go faster than the speed of its own exhaust. In other words, a jet with fuel fraction > 0.5 can go faster than its own exhaust if that is what you are trying to design it to do.


But that is a theoretical digression.
What I want to understand is whether a turbojet would be superior to a low bypass turbofan for supercruise. Assuming both engines have the same exhaust velocity, would the fan be more efficient because of the bypass air? Or would this produce more drag? Let's suppose both engines have the same overall mass flow rate, will the turbofan inherently have a lower exhaust velocity? Do turbojets always have a higher exhaust velocity than turbofans in the same thrust class in the real world?
 
My (limited) understanding is that turbojets were (are ?) superior to turbofans at supersonic speed because hot gases are better than cold air when flying fast.

Flying fast = needs huge volume of gas going out of the nozzle.
Turbojet: all gases are hot as hell
Turbofan: a handful of hot gases are sunk amid a big mass of cold, unburned air.

For Concorde and SST generally, it is either "straight turbojet" or "turbofan with reheat" - none of the two are presently acceptable - guess why Aerion sunk ?

AFAIK the Tu-144 Kuznetsov engine was created by making an Il-62 NK-8 turbofan core, supersonic with an afterburner. The end result was a terrible engine.

Concorde went for straight Olympus turbojet from the beginning and it worked very well. Except for takeoff noise, fuel consumption, and pollution, of course. A turbofan would solve both issues (very critical post-1973 oil shock) but also greatly suck past Mach 1. So it was a conundrum.

Boeing "2707-300" SST engine was a massively scaled up B-70 engine (J93 to GE4): another turbojet with a colossal mass of hot gases going through the exhaust. That what it took to hit MAch 2.7 in cruise - on paper at least.

Turbofans were created only to save fuel by mixing air - cold, unburned air is free, kerosene is not. That's the basic reasoning: "dilution" is the exact word.
a) for airliners, to extend range, cut into pollution, and save oil after 1973
Note that VC-10 and CV-880 speed records will never be matched by present day Airbus and Boeing: airliner speed fell from mach 0.90+ to Mach 0.85- when turbofans kicked away turbojets, in the early 60's.
b) for the military: range again, notably for the B-1A / B-1B either at supersonic speed or low height. Also F-111 with the crappy TF30. RR followed with their Medway / Spey family, and we are three generations into this now.
From there it moved to the F-15, F-16, and then all the others since then.
 
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The current state of the art, the F-119 as used in the F-22, offers supercruise up to Mach 1.8 and is a turbofan.
 
Sure, the bypass ratio is low, but the point is they went to the trouble of designing it with a bypass instead of building a (simpler) turbojet.
 
@GUNDAM123dx You need to read those Lightning charts at ISA temperatures, rather than cherry-picking the most favorable temperatures indicated, which would require arctic weather conditions (!).

The Lightning manual shows an ISA supercruise performance of just over Mach 1.1 clean at 25,000-35,000 ft. Or just under Mach 1.1 with 2x missiles.

For reference, ISA equals 15 degrees celcius at 0ft or -44 degrees at 30,000ft. In order to get the -76 degree temperature needed for Mach 1.4, the ground temperature would need be colder than -15 degrees celcius.
 
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This is one of the reasons the new power plants are variable bypass and why the GE submission for the ATF program was variable bypass; turbojet supersonically, turbofan subsonically. That offers the best optimization. Ultimately what it comes down to is thrust versus drag. You want minimum drag and maximum efficient thrust at design cruise point. The F-22 was optimized for low supersonic drag since it was specifically tasked to super cruise. The F-35 isn't optimized for lower supersonic drag, because that isn't it's primary mission. It has to carry large weapons, relative to the F-22, internally and fit on a ship and keep weight down for STOVL operations. All of those design points aren't great for making a low wave drag airframe. I saw someone refer to the F-35 as a Battle Penguin, and I think that describes it's shape perfectly. From the standpoint of the fineness ratio, it's a bit portly. That's basically the ratio of the length to how fat it is. Generally, when flying fast, you want a high fineness ratio for lower wave drag. Think Blackbird and Concorde. The lower the fineness ratio, the higher the drag, the more thrust will be required. You're always playing with the trades based on the mission(s) requirements.
 
The F-35 front section (just like the F-22) is designed to create multiple weaker shockwaves instead of a single strong one. That allows the center and rear sections to be larger than if they had to fit inside a strong shockwave cone.
Being fatter means really nothing once you are beyond the shock wave. That has been the main principle fighter jets have used since the supersonic age.
As an example, the X-3 was a poor supersonic performer despite being the most pointy one among the early generation of X-planes.
 
I've read that even with the F-22 it is preferred to use afterburner to get past the sound barrier before cutting it back because it's actually more efficient than getting there with dry thrust alone although that is possible.
 
what about pressure inside the engine? Do afterburners increase the pressure aft of the engine core? If so, one could conclude that power is made with added pressure.


Now since there is no afterburner in supercruise, all that added pressure must come within the engine core itself. bypass portion of thrust starts to rapidly drop in transsonic region, so for supercruise one can rely only on non-bypassed thrust, meaning the one coming from the engine core. Older engines couldn't really withstand high pressure so we didn't really have supercruise before. Nowadays, with newer engines supercruise is more prolific.


alternative would be an oversized, overpowered engine for a given aiframe, even if built with older tech, like the pegasus engines on the concorde. But i guess that is not very efficient to use on relatively small planes like the fighters today.


Is all this that i've written bollocks or is there some truth to it?
Afterburners do not increase the pressure after the turbine to increase the thrust of the engine, they add heat and a little bit of additional mass flow due to the weight of the fuel being injected into the exhaust flow.

The afterburner is essentially a constant pressure device. There can be a diffuser section prior to the flameholder area to increase the static pressure by reducing velocity, but the flame area is basically constant pressure until it reaches the exhaust nozzle, where the pressure is converted to velocity.

The nozzle pressure ratio will determine the exhaust flow Mach number - the higher nozzle pressure ratio, the higher the exhaust Mach number. A NPR of approximately 2 will give Mach 1 thru a convergent nozzle. A NPR greater than 2 needs a divergent nozzle for supersonic expansion with exit Mach numbers greater than 1.

The difference in thrust between Military and AB settings is due to the increased temperature of the exhaust flow. The local Mach velocity goes up by the square root of absolute temperature ratio. So if the afterburner increased the exhaust absolute temperature by 2x, the exhaust Mach number would remain the same, but the exhaust velocity would increase by a factor of 1.4. If the AB increased the absolute temperature by 4x, you would approximately double the exhaust velocity and static thrust of the engine, all else remaining equal.

Because higher bypass engines typically have a cooler average exhaust temperature and more oxygen available to burn, their AB can achieve a larger increase in exhaust temperature than a turbojet / low bypass turbofan. This is reflected in their higher AB thrust increase, but also their worse AB specific fuel consumption.
 
For a given airframe, what is the difference between an engine that can supercruise and one that can't? Is it purely a difference of dry thrust? Or would the engines exhibit other differences as well.
To supercruise, an engine must have the lowest bypass ratio possible to keep the exhaust jet's velocity high; a turbojet is preferable to a turbofan. After all, a jet can fly only as fast as the velocity of its exhaust jet and the bypass air slows the exhaust jet's velocity down.
Correct - to supercruise, an engine must have a high exhaust velocity, at the supercruise engine inlet conditions. And a clean airframe with low supersonic drag. Both are required.

High exhaust velocity means a high nozzle pressure ratio going thru an optimized convergent / divergent exhaust nozzle. Nozzle pressure ratio is engine pressure ratio (exhaust pressure / engine inlet pressure) x ram pressure (engine inlet pressure / ambient pressure). Higher exhaust temperature also increases thrust, but it only goes up with the square root of the absolute temperature increase (same relationship as speed vs Mach number).

To get a high EPR at Mil power with a turbofan engine, you need a high pressure ratio Fan section, and a core module big and powerful enough to drive that high pressure Fan. The F100-220 fan pressure ratio is around 3:1, and the supercruise engines are significantly higher than that. At the typical quoted supercruise condition of 40K ft, 1.5Mn, the inlet temperature of the engine is approximately 100F. Most engines reach their rotor speed and turbine temperature limits in the 60-70F range. As the inlet gets hotter, these engines lose airflow, EPR, and thrust. At 100F inlet supercruise, both the Fan/LPT and Core have have the design margin to turn faster and hotter than at 60F inlet, just to keep making the same thrust (airflow and EPR). This points to having both a robust high pressure ratio fan, and probably an even larger core module. Thus, your design points to a low bypass turbofan - i.e. the F119.
more like a leaky turbojet, certainly not in the bypass range of a TF30 even, so not truly a turbofan but rather a very optimized turbojet for supercruise..

The F119 is a low bypass turbofan. Why turbofan instead of turbojet? A turbojet has a very hot outer skin in the main combustor, turbine, and AB areas, requiring heat shielding (weight) and a significant amount of secondary airflow (i.e. drag) to keep the airframe cool. The low bypass design of the F119 enables the structural fan ducts to be the heatshield, making it a "self cooled turbojet". The bypass air is also used extensively to cool the exhaust and nozzle components.

The other part of that thrust equation in the inlet ram recovery. At 40K, 1.5Mn the engine inlet pressure is around 10 psia, approximately 3.5 times the ambient pressure. Multiply that by the supercruise EPR, and you have a nozzle pressure ratio greater than 10:1. The convergent section of the nozzle accelerates the flow to Mach 1 using about 2:1 pressure ratio, leaving 5:1 or greater to be expanded in the divergent nozzle for supersonic flow. If you have a typical fixed ratio convergent / divergent nozzle, it will not have a large enough divergent section to efficiently utilize that available nozzle pressure ratio - i.e. under expanded flow. You could make the divergent section larger, but that adds weight, drag, and overexpansion concerns at lower speeds. The F119 engine, since it has full control of the 2D nozzle divergent flaps for thrust vectoring, can also position the divergent flap angles to optimize the supersonic expansion over a wide range of flight conditions and power settings.
 
My (limited) understanding is that turbojets were (are ?) superior to turbofans at supersonic speed because hot gases are better than cold air when flying fast.

Flying fast = needs huge volume of gas going out of the nozzle.
Turbojet: all gases are hot as hell
Turbofan: a handful of hot gases are sunk amid a big mass of cold, unburned air.

For Concorde and SST generally, it is either "straight turbojet" or "turbofan with reheat" - none of the two are presently acceptable - guess why Aerion sunk ?

AFAIK the Tu-144 Kuznetsov engine was created by making an Il-62 NK-8 turbofan core, supersonic with an afterburner. The end result was a terrible engine.

Concorde went for straight Olympus turbojet from the beginning and it worked very well. Except for takeoff noise, fuel consumption, and pollution, of course. A turbofan would solve both issues (very critical post-1973 oil shock) but also greatly suck past Mach 1. So it was a conundrum.

Boeing "2707-300" SST engine was a massively scaled up B-70 engine (J93 to GE4): another turbojet with a colossal mass of hot gases going through the exhaust. That what it took to hit MAch 2.7 in cruise - on paper at least.

Turbofans were created only to save fuel by mixing air - cold, unburned air is free, kerosene is not. That's the basic reasoning: "dilution" is the exact word.
a) for airliners, to extend range, cut into pollution, and save oil after 1973
Note that VC-10 and CV-880 speed records will never be matched by present day Airbus and Boeing: airliner speed fell from mach 0.90+ to Mach 0.85- when turbofans kicked away turbojets, in the early 60's.
b) for the military: range again, notably for the B-1A / B-1B either at supersonic speed or low height. Also F-111 with the crappy TF30. RR followed with their Medway / Spey family, and we are three generations into this now.
From there it moved to the F-15, F-16, and then all the others since then.
I'm aware that turbofans generally have lower exhaust velocity than turbojets, they make up for this by increasing mass flow rate, moving more air than a turbojet with minimal increases in fuel consumption compared to a turbojet of the same size as a given turbofan's core. The lower exhaust velocity is also more desirable for airliners as it is quieter. The question I'm trying to ask is, what makes a turbojet more efficient at supersonic speeds than a LOW BYPASS turbofan like what we see on modern fighter aircraft. Will a low bypass fan always have a lower exhaust velocity than a turbojet of a similar thrust class?
 
I'm aware that turbofans generally have lower exhaust velocity than turbojets, they make up for this by increasing mass flow rate, moving more air than a turbojet with minimal increases in fuel consumption compared to a turbojet of the same size as a given turbofan's core. The lower exhaust velocity is also more desirable for airliners as it is quieter. The question I'm trying to ask is, what makes a turbojet more efficient at supersonic speeds than a LOW BYPASS turbofan like what we see on modern fighter aircraft. Will a low bypass fan always have a lower exhaust velocity than a turbojet of a similar thrust class?

Bypass air will always reduce the exhaust velocity of the mixed stream. However at low bypass ratios, this reduction will be relatively minor and compensated by a reduction on SFC and a useful cooling effect on the engine outer casing. Also for any given engine, I think the exhaust velocity is related to engine pressure ratio, engine temperature ratio, and nozzle performance, so more modern engines with higher technical parameters might outperform older ones even if the newer engine is a low pressure turbofan versus an older pure turbojet.
 
I'm aware that turbofans generally have lower exhaust velocity than turbojets, they make up for this by increasing mass flow rate, moving more air than a turbojet with minimal increases in fuel consumption compared to a turbojet of the same size as a given turbofan's core. The lower exhaust velocity is also more desirable for airliners as it is quieter. The question I'm trying to ask is, what makes a turbojet more efficient at supersonic speeds than a LOW BYPASS turbofan like what we see on modern fighter aircraft. Will a low bypass fan always have a lower exhaust velocity than a turbojet of a similar thrust class?

Bypass air will always reduce the exhaust velocity of the mixed stream. However at low bypass ratios, this reduction will be relatively minor and compensated by a reduction on SFC and a useful cooling effect on the engine outer casing. Also for any given engine, I think the exhaust velocity is related to engine pressure ratio, engine temperature ratio, and nozzle performance, so more modern engines with higher technical parameters might outperform older ones even if the newer engine is a low pressure turbofan versus an older pure turbojet.
Actually, bypass air will always reduce the exhaust temperature of the mixed stream, not necessarily the pressure. The lower temperature is why SFC is reduced (i.e. less wasted heat going out the exhaust), but it does reduce exhaust velocity by the square root of the absolute temperature decrease ratio to standard day temperature [519R (59F) or 288K (15C)], with all else being equal.

With the same core performance, increasing the fan airflow to increase the bypass ratio will reduce the exhaust pressure and velocity because it takes more work out of the core exhaust flow to turn the bigger fan, and the fan pressure ratio is usually reduced proportionally with a mixed flow exhaust.

But, as PaulMM notes, a higher performing core can bring the fan and engine pressure ratio back to the same level as before, and exhaust velocity increase is roughly proportional to engine pressure ratio (EPR) increase.

The real challenge is to maintain airflow and EPR as the inlet temperature increases with increased Mn. As the inlet temperature increases, the rotor speed and internal temperatures of the engine have to keep increasing to maintain that flow. Once you reach the rotor and temperature limits of the engine, the performance of the engine decreases, either airflow, EPR, or both.

As you fly faster, the increasing intake ram pressure ratio multiplies the available EPR to provide nozzle pressure ratio, which determine the exhaust flow Mn. On the SR-71 at Mach 3, the 600F inlet has the engine turning 100% mechanical speed, but the temperature corrected rotor speed is only around 65% with an EPR of about 1.1 - just above Idle thrust. But the ram recovery is on the order of 30:1, giving a nozzle pressure ratio of 33:1, which results in a very high exhaust Mn. Of course, this is not a supercruising engine, but it gives you an idea how the engine and inlet work together to provide supersonic thrust.
 
But, as PaulMM notes, a higher performing core can bring the fan and engine pressure ratio back to the same level as before, and exhaust velocity increase is roughly proportional to engine pressure ratio (EPR) increase.
"Same level as it was before" as in a turbojet? Are you saying a high performing turbofan core can achieve exhaust velocities of a similarly sized turbojet?
The real challenge is to maintain airflow and EPR as the inlet temperature increases with increased Mn. As the inlet temperature increases, the rotor speed and internal temperatures of the engine have to keep increasing to maintain that flow. Once you reach the rotor and temperature limits of the engine, the performance of the engine decreases, either airflow, EPR, or both.
So as airspeed increases, the temperature (and as a result, pressure) in the inlet increases, which reduces EPR? Or am I reading this wrong? Admittedly my understanding of how pressure and airspeed differences affect jet engine (specifically gas turbine jets) thrust performance is lacking, I understand the basic theory of operation of turbine engines and that thrust is a product of mass flow and velocity, and now I'm understanding pressure as well affects thrust output. So tell me, does an afterburner increase thrust primarily through increasing velocity or pressure in the exhaust outlet?
 
For a given airframe, what is the difference between an engine that can supercruise and one that can't? Is it purely a difference of dry thrust? Or would the engines exhibit other differences as well.
To supercruise, an engine must have the lowest bypass ratio possible to keep the exhaust jet's velocity high; a turbojet is preferable to a turbofan. After all, a jet can fly only as fast as the velocity of its exhaust jet and the bypass air slows the exhaust jet's velocity down.
Correct - to supercruise, an engine must have a high exhaust velocity, at the supercruise engine inlet conditions. And a clean airframe with low supersonic drag. Both are required.

High exhaust velocity means a high nozzle pressure ratio going thru an optimized convergent / divergent exhaust nozzle. Nozzle pressure ratio is engine pressure ratio (exhaust pressure / engine inlet pressure) x ram pressure (engine inlet pressure / ambient pressure). Higher exhaust temperature also increases thrust, but it only goes up with the square root of the absolute temperature increase (same relationship as speed vs Mach number).

To get a high EPR at Mil power with a turbofan engine, you need a high pressure ratio Fan section, and a core module big and powerful enough to drive that high pressure Fan. The F100-220 fan pressure ratio is around 3:1, and the supercruise engines are significantly higher than that. At the typical quoted supercruise condition of 40K ft, 1.5Mn, the inlet temperature of the engine is approximately 100F. Most engines reach their rotor speed and turbine temperature limits in the 60-70F range. As the inlet gets hotter, these engines lose airflow, EPR, and thrust. At 100F inlet supercruise, both the Fan/LPT and Core have have the design margin to turn faster and hotter than at 60F inlet, just to keep making the same thrust (airflow and EPR). This points to having both a robust high pressure ratio fan, and probably an even larger core module. Thus, your design points to a low bypass turbofan - i.e. the F119.
more like a leaky turbojet, certainly not in the bypass range of a TF30 even, so not truly a turbofan but rather a very optimized turbojet for supercruise..

The F119 is a low bypass turbofan. Why turbofan instead of turbojet? A turbojet has a very hot outer skin in the main combustor, turbine, and AB areas, requiring heat shielding (weight) and a significant amount of secondary airflow (i.e. drag) to keep the airframe cool. The low bypass design of the F119 enables the structural fan ducts to be the heatshield, making it a "self cooled turbojet". The bypass air is also used extensively to cool the exhaust and nozzle components.

The other part of that thrust equation in the inlet ram recovery. At 40K, 1.5Mn the engine inlet pressure is around 10 psia, approximately 3.5 times the ambient pressure. Multiply that by the supercruise EPR, and you have a nozzle pressure ratio greater than 10:1. The convergent section of the nozzle accelerates the flow to Mach 1 using about 2:1 pressure ratio, leaving 5:1 or greater to be expanded in the divergent nozzle for supersonic flow. If you have a typical fixed ratio convergent / divergent nozzle, it will not have a large enough divergent section to efficiently utilize that available nozzle pressure ratio - i.e. under expanded flow. You could make the divergent section larger, but that adds weight, drag, and overexpansion concerns at lower speeds. The F119 engine, since it has full control of the 2D nozzle divergent flaps for thrust vectoring, can also position the divergent flap angles to optimize the supersonic expansion over a wide range of flight conditions and power settings.
There's something that's bugged me for years. Based on the Dash-1s (F-16 I believe) the -229s exhaust is about a thousand degrees hotter, and much higher in velocity, than that of the -129. Yet (apparently) the -129 performs better at high velocity. Any idea why this might be the case?
 

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