Forum: F-35 Engine

Regardless, the current figures are impressive nonetheless................

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Elite 1K Joined: Aug 19, 2004 - 09:19 AM Posts: 1092

dwightlooi wrote:

Raptor_One wrote:

I'm sorry, but that is just wrong. A variety of factors determine an engine's thrust output... and that's not taking into account the aircraft it's installed in. Overall pressure ratio and mass flow rate are by no means the entire equation. That would, for example, ignore temperature ratios. If you've ever done engine cycle analysis, you would understand that things are not this simple... not by a long shot.

It is primarily so. Temperature ratios affect the pressure ratio, but in the end the temperature of the exhaust doesn't matter thrust wise -- even though achieving the same thrust with less airflow and higher pressure ratios means having a higher temperature ratio. The only things that matter are how much air is going in the engine and at what pressure that ar is coming out the back! The only things that make thrust are volume and pressure. For a given nozzle area, they determine exhaust mass and exhaust velocity and those are the ONLY things that matter as to how much thrust you make.

How efficient fuel wise, how hot the exhaust plume is, etc are other matters that concern an engine. But thrust is thrust.

Actually, an afterburning turbojet or afterburning turbofan (such as is used in fighter aircraft) generate thrust through high exit velocity. If you mean to tell me that temperature has nothing to do with generating this high exit velocity.... well... okay. Being able to generate higher and higher temperatures inside the engine core and aft of the core are constant goals of engine designers. The hotter you can make it inside, say, an AB turbofan, the more thrust you can generate. Perhaps you should study the Brayton cycle.

Raptor_One wrote:

dwightlooi wrote:

Raptor_One wrote:

I'm sorry, but that is just wrong. A variety of factors determine an engine's thrust output... and that's not taking into account the aircraft it's installed in. Overall pressure ratio and mass flow rate are by no means the entire equation. That would, for example, ignore temperature ratios. If you've ever done engine cycle analysis, you would understand that things are not this simple... not by a long shot.

It is primarily so. Temperature ratios affect the pressure ratio, but in the end the temperature of the exhaust doesn't matter thrust wise -- even though achieving the same thrust with less airflow and higher pressure ratios means having a higher temperature ratio. The only things that matter are how much air is going in the engine and at what pressure that ar is coming out the back! The only things that make thrust are volume and pressure. For a given nozzle area, they determine exhaust mass and exhaust velocity and those are the ONLY things that matter as to how much thrust you make.

How efficient fuel wise, how hot the exhaust plume is, etc are other matters that concern an engine. But thrust is thrust.

Actually, an afterburning turbojet or afterburning turbofan (such as is used in fighter aircraft) generate thrust through high exit velocity. If you mean to tell me that temperature has nothing to do with generating this high exit velocity.... well... okay. Being able to generate higher and higher temperatures inside the engine core and aft of the core are constant goals of engine designers. The hotter you can make it inside, say, an AB turbofan, the more thrust you can generate. Perhaps you should study the Brayton cycle.

Making and running at higher temperatures inside has EVERYTHING to do with exit velocity. What high temperature does is allow you to end up with a HIGHER PRESSURE RATIO! Hence, at the end of the day once you already have a pressure ratio statistic what the temperature ratio is has no bearing on the thrust. None. Its effect is in producing the pressure ratio!

If you hold mass air flow constant, then the pressure ratio is in large part dependent on the temperature ratio in the core, the bypass ratio, the efficiency of the fan, compressor and the turbines. If you hold the pressure ratio constant, then the thrust is dependent on the mass air flow through the engine. How you achieve that again depends on the inner workings of the engine.

As I said, given a mass air flow number and the pressure ratio, you can pretty much know how much air is gong through the engiine and at what pressure that air is prior to going through the exhaust nozzle. If you also know the nozzle dimensions then you CAN calculate thrust. Temperature has no place in the calculations. It doesn't matter at all. Thrust ONLY DEPENDS on two things how much air molecues are coming out the tail pipe per second (per unit time) and how fast they are moving as the come out (exhaust velocity)! This is fundamentally a matter of newtonian physics! The former is measured as mass air flow -- it is the same at the intake and at the exhaust since air that goes in comes out. The latter is a function of PRESSURE in the exhaust can and the area of the nozzle. It is as simple as that!

Elite 1K Joined: Aug 19, 2004 - 09:19 AM Posts: 1092

Okay... write out the basic thrust equation for a turbojet engine. The only place that exit pressure comes into play is when the exit pressure is different from the ambient pressure (in which case you have a thrust loss). And no... temperature is not explicitly in that equation either. In fact, why don't you just write out the full cycle equation for an afterburning turbojet engine and then we can argue more.

For the record, I do not agree with the following statement you made:

"Basically two things determine thrust -- mass air flow and presssure ratio."

I disagree with that statement on a number of different levels. Write out the full cycle equation for an afterburning turbojet engine (or a mixed flow afterburning turbofan) and I will say exactly why I disagree with your statement above by pointing out various "features" of the cycle equation. Have a go at it.

I'm fairly certain that when they're referring to "pressure ratio" they're referring to how much the air gets crunched by the compressor.

Raptor_One wrote:

asiatrails wrote:

Raptor, you are corect. My intention was not to start a debate on cycle decks; all I intended to indicate was that the current system is within existing operational parameters and it should be possible to grow the engine.

Thrust growth with an existing envelope always happens; if you are at the current edge of what is possible it costs a lot more and takes a lot longer to grow.

I wasn't referring to your basic quoting of figures. There's nothing wrong with what you posted (unless it's actually factually inaccurate Smile

).

I quoted publically released numbers for the -129 which can be confirmed on the GE website, I am not aware that the numbers for the -132 have been publically released so I can not discuss them.

Elite 1K Joined: Aug 19, 2004 - 09:19 AM Posts: 1092

asiatrails wrote:

I quoted publically released numbers for the -129 which can be confirmed on the GE website, I am not aware that the numbers for the -132 have been publically released so I can not discuss them.

No problemo. I'm sure the -132 has impressive specs. The -129 ain't no slouch.

Elite 1K Joined: Aug 19, 2004 - 09:19 AM Posts: 1092

sferrin wrote:

I'm fairly certain that when they're referring to "pressure ratio" they're referring to how much the air gets crunched by the compressor.

That's what does the majority of the compression, but there is an overall pressure ratio for the engine which is what they are probably quoting. Each stage has a pressure ratio. Put them all together and you get the overall pressure ratio.

An engine's pressure ratio (EPR) is defined as the total pressure ratio across the engine. As EPR is the ratio of the nozzle total pressure to the compressor face total pressure it can be easily measured and displayed.

It is important to realize that the total pressure losses in the inlet are not contained in the calculation, installation effects (Reality) changes theoretical performance.

EPR = compressor pressure ratio * burner pressure ratio * turbine pressure ratio * nozzle pressure ratio

asiatrails wrote:

An engine's pressure ratio (EPR) is defined as the total pressure ratio across the engine. As EPR is the ratio of the nozzle total pressure to the compressor face total pressure it can be easily measured and displayed.

It is important to realize that the total pressure losses in the inlet are not contained in the calculation, installation effects (Reality) changes theoretical performance.

EPR = compressor pressure ratio * burner pressure ratio * turbine pressure ratio * nozzle pressure ratio

Okay. I'd always thought it had to do with how far the engine compressed the air. For example 320 in^3 of air would be compressed to 10 in^3 by the time it hit the combustor in a 32-1 engine.

Raptor_One wrote:

Okay... write out the basic thrust equation for a turbojet engine. The only place that exit pressure comes into play is when the exit pressure is different from the ambient pressure (in which case you have a thrust loss). And no... temperature is not explicitly in that equation either. In fact, why don't you just write out the full cycle equation for an afterburning turbojet engine and then we can argue more.

For the record, I do not agree with the following statement you made:

"Basically two things determine thrust -- mass air flow and presssure ratio."

I disagree with that statement on a number of different levels. Write out the full cycle equation for an afterburning turbojet engine (or a mixed flow afterburning turbofan) and I will say exactly why I disagree with your statement above by pointing out various "features" of the cycle equation.

Have a go at it.

Call me stupid but doesn't F=ma even for turbojets? Rolling Eyes

Wouldn't mass and exit velocity be the two key elements in determining thrust? The only place temp would come into play is a hotter engine enabling greater exhaust velocity because of greater pressure for a given mass of air.

Excellent news indeed! Definately a step in the right direction, especially for a single engine a/c. Supercruise ought to be well within plausable range, at least in clean (internal stores only) configuration...

Folks,

The overall goal is to take air in at Vo (flight speed), throw it out at Vo + DV.

Thrust is largely composed of the net change in momentum of the air entering and leaving the engine, with a typically small adjustment for the differences in pressure between the inlet and the exit

The NASA Glenn has some excellent detailed information and a basic engine simulator.

http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html

sferrin wrote:

Raptor_One wrote:

Okay... write out the basic thrust equation for a turbojet engine. The only place that exit pressure comes into play is when the exit pressure is different from the ambient pressure (in which case you have a thrust loss). And no... temperature is not explicitly in that equation either. In fact, why don't you just write out the full cycle equation for an afterburning turbojet engine and then we can argue more.

For the record, I do not agree with the following statement you made:

"Basically two things determine thrust -- mass air flow and presssure ratio."

I disagree with that statement on a number of different levels. Write out the full cycle equation for an afterburning turbojet engine (or a mixed flow afterburning turbofan) and I will say exactly why I disagree with your statement above by pointing out various "features" of the cycle equation.

Have a go at it.

Call me stupid but doesn't F=ma even for turbojets? Rolling Eyes

Wouldn't mass and exit velocity be the two key elements in determining thrust? The only place temp would come into play is a hotter engine enabling greater exhaust velocity because of greater pressure for a given mass of air.

[quote="asiatrails"]Folks,

The overall goal is to take air in at Vo (flight speed), throw it out at Vo + DV.

Thrust is largely composed of the net change in momentum of the air entering and leaving the engine, with a typically small adjustment for the differences in pressure between the inlet and the exit

The NASA Glenn has some excellent detailed information and a basic engine simulator.

http://www.grc.nasa.gov/WWW/K-12/airplane/ngnsim.html

I don't know that I'd bank on it's accuracy as they have an F100 overheating at Mach 1.7 at 36,000ft.

Alright people, there's about 20 variants of thrust equations all of which are right and can be translated to other forms. Often when you see a pressure ratio with an engine it is the compressor pressure ratio. This means that the STAGNATION pressure is that many times higher than ambient. What you want on the back end is a stag pressure about 2x ambient (actually 1.891). This means that if you shove it through a converging diverging nozzle, the flow will choke and you'll get supersonic exit flow. Now, the equation in my main book is all put in terms of temperature ratios, its way WAY to long to right out here.

Right now there are engines with pressure ratios in the 40's and one in the 50's. I hate to say it, but the one with a CPR of 50 is a Rolls.

By the way, easiest thrust equation: T=Mdot*(u9-u0)
Mdot= mass flow
u9= exit velocity
u0=entry velocity
That is where the pressure comes in is the need to choke the flow for the extra speed on exit. The turbine pressure ratio is not going to be the same because you need the extra pressure on the back end.

Forum: F-35 Engine

F-35's engine will have 43,000 lbs of thrust.