I know that the efficiency will vary by the thrust setting. I assume engines are designed to be most efficient at cruise thrust. In terms of the percentage of the energy in the fuel that is output as usable thrust, what is the efficiency of the current generation of turbofan engines?

planecane wrote:

I know that the efficiency will vary by the thrust setting. I assume engines are designed to be most efficient at cruise thrust. In terms of the percentage of the energy in the fuel that is output as usable thrust, what is the efficiency of the current generation of turbofan engines?

most efficient for the use case.
cruise TSFC is quite a bit higher than full pressure TSFC.
~~.3 something vs ~~.5..something ( lb/lbf/h )

IMU higher thrust capability (TWINs, for One Engine Out ) also comes with a tsfc penalty.
( a boon to the 737: lesser OEO requirements allow less thrust capability giving a slight advantage
in cruise ( with all other design metrics like BPR being the same.)

planecane wrote:

I know that the efficiency will vary by the thrust setting. I assume engines are designed to be most efficient at cruise thrust. In terms of the percentage of the energy in the fuel that is output as usable thrust, what is the efficiency of the current generation of turbofan engines?

The most efficient engines are on the A350, which convert about 40% of the fuel energy.

https://www.nap.edu/read/23490/chapter/6#37

More discussion here:

viewtopic.php?f=5&t=1404589

Ultrafan was said to be 25% more efficient than the first RR Trent turbofans. Wikipedia has the Trent 700 cruise SFC as 0.565 lb/lbf/h, so Ultrafan would've been about 0.42 lb/lbf/h.

CowAnon wrote:

planecane wrote:

I know that the efficiency will vary by the thrust setting. I assume engines are designed to be most efficient at cruise thrust. In terms of the percentage of the energy in the fuel that is output as usable thrust, what is the efficiency of the current generation of turbofan engines?

The most efficient engines are on the A350, which convert about 40% of the fuel energy.

https://www.nap.edu/read/23490/chapter/6#37

I have a question about this graph.
50%*70%=35, not 20% like shown there

Armadillo1 wrote:

CowAnon wrote:

planecane wrote:

I know that the efficiency will vary by the thrust setting. I assume engines are designed to be most efficient at cruise thrust. In terms of the percentage of the energy in the fuel that is output as usable thrust, what is the efficiency of the current generation of turbofan engines?

The most efficient engines are on the A350, which convert about 40% of the fuel energy.

https://www.nap.edu/read/23490/chapter/6#37

I have a question about this graph.
50%*70%=35, not 20% like shown there

Here's the graph. If you're referring to the B747-400, that is around halfway between the 30% and 40% contours, so that's about correct. The 20% label is for a different contour.



Here's a different graph, which shows propulsive efficiency and thermal*transfer efficiency plotted against SFC (source: Seeking Alpha/Aspire Aviation, Rolls-Royce At A Transitional Altitude, 8/31/2015):

In layman terms, where did the efficiency improvements come from in modern turbo fans? Let’s say JT9D vs RR Trent XWB.

T54A wrote:

In layman terms, where did the efficiency improvements come from in modern turbo fans? Let’s say JT9D vs RR Trent XWB.

I guess I'll give this a try.

JT9D -> Boeing 747-100 => 30% overall efficiency (~64% propulsive, ~48% motor thermodynamic)
RR Trent XWB -> Airbus A350 => 40% overall efficiency (~71% propulsive, ~57% motor thermodynamic)

The propulsive improvement is due to the increase in bypass ratio from 4.8:1 to 9.6:1. Bypass ratio is the circular area covered by the turning fan divided by the gas exhaust's cross-sectional area. To convert the hot compressed air resulting from fuel combustion into propulsion, it's more efficient to turn a turbine shaft connected to a fan that pushes a lot of ambient air backward. That beats directly having the heated/compressed air escape backward, which propels the aircraft forward from the air's expansion from cooling to atmospheric temperature and from uncompressing to atmospheric pressure. The bypass ratio increase was possible because of materials improvements, which allowed fan blades to become longer and thinner. Improved materials also allowed for the blade edges to be swept instead of just being straight/flat, which raised the efficiency of the fan.

I understand the engine internals/thermodynamics less, but I can say that the structural materials improved to allow higher temperatures and pressures without melting/damaging the components. Higher temps and pressures mean the engine's turbines can extract more energy from the heated and compressed air. The pressure ratio (the pressure of air exiting the high-pressure compressor divided by the pressure of the outside atmosphere) went from 26.7:1 on the JT9D to 50:1 on the Trent XWB. Also, the JT9D has a two-shaft architecture, while the Trent XWB is three-shaft. For two-shaft, the fan and the low-or-intermediate pressure compressor are driven by the same turbine on one long shaft, so they have to run at the same rotational speed. For three-shaft, those two components are driven by different turbines on separate shafts, so those components can turn at different, more optimal speeds (slower for the fan, faster for the LP/IP compressor). Having separate shafts avoids other design constraints; on the two-shaft architecture, if you optimize the fan speed, you might need to add more LP/IP compression stages, but if you optimize the LP/IP compressor speed, you might have to reduce the diameter of the fan. The three-shaft architecture also makes the engine shorter and lighter.

Comments/corrections welcomed.

CowAnon wrote:

T54A wrote:

In layman terms, where did the efficiency improvements come from in modern turbo fans? Let’s say JT9D vs RR Trent XWB.

I guess I'll give this a try.

JT9D -> Boeing 747-100 => 30% overall efficiency (~64% propulsive, ~48% motor thermodynamic)
RR Trent XWB -> Airbus A350 => 40% overall efficiency (~71% propulsive, ~57% motor thermodynamic)

The propulsive improvement is due to the increase in bypass ratio from 4.8:1 to 9.6:1. Bypass ratio is the circular area covered by the turning fan divided by the gas exhaust's cross-sectional area. To convert the hot compressed air resulting from fuel combustion into propulsion, it's more efficient to turn a turbine shaft connected to a fan that pushes a lot of ambient air backward. That beats directly having the heated/compressed air escape backward, which propels the aircraft forward from the air's expansion from cooling to atmospheric temperature and from uncompressing to atmospheric pressure. The bypass ratio increase was possible because of materials improvements, which allowed fan blades to become longer and thinner. Improved materials also allowed for the blade edges to be swept instead of just being straight/flat, which raised the efficiency of the fan.

I understand the engine internals/thermodynamics less, but I can say that the structural materials improved to allow higher temperatures and pressures without melting/damaging the components. Higher temps and pressures mean the engine's turbines can extract more energy from the heated and compressed air. The pressure ratio (the pressure of air exiting the high-pressure compressor divided by the pressure of the outside atmosphere) went from 26.7:1 on the JT9D to 50:1 on the Trent XWB. Also, the JT9D has a two-shaft architecture, while the Trent XWB is three-shaft. For two-shaft, the fan and the low-or-intermediate pressure compressor are driven by the same turbine on one long shaft, so they have to run at the same rotational speed. For three-shaft, those two components are driven by different turbines on separate shafts, so those components can turn at different, more optimal speeds (slower for the fan, faster for the LP/IP compressor). Having separate shafts avoids other design constraints; on the two-shaft architecture, if you optimize the fan speed, you might need to add more LP/IP compression stages, but if you optimize the LP/IP compressor speed, you might have to reduce the diameter of the fan. The three-shaft architecture also makes the engine shorter and lighter.

Comments/corrections welcomed.

Couple of good points in here, ans a few minor clarifications I'd add:
1. The bypass ratio point is mostly on the money, the one thing I'd change there is that it's the mass flow of the bypass versus the core that you need to measure, so the ratios of the areas only works if they have the same jet efflux speed and pressure, which is not generally the case, but it's a close approximation and good visual nevertheless.
2. The compression ratio has to do with the delta-T, so the higher the pressure, the greater the efficiency.
3. One problem with increased engine efficiency is plateauing system efficiency. Higher bypass and pressure ratios involve heavier engines, which need to be carried, so there's an engineering tradeoff at the system level where it makes sense to sacrifice engine efficency for system efficiency.

Thermodynamic efficiencies are around 55%? Seriously?
The best combined cycle gas turbines reach an efficiency of IIRC 63%. Now there may be a few percent losses from running the generator, but I find 55% thermal efficiency hard to believe.
Or are CCGTs run with less pressure/ temperature to minimize maintenance?

I'm also surprised that propulsive efficiency is above 70%. I'm not surprised for a plane in the beginning of the climb, but at high speeds? Maybe I imagine that wrong.
Should I think of it as a stationary fan/ propeller with the air entering with the speed of the plane?
If so, how many km/h does the fan accelerate the air?

LH707330 wrote:

CowAnon wrote:

T54A wrote:

In layman terms, where did the efficiency improvements come from in modern turbo fans? Let’s say JT9D vs RR Trent XWB.

I guess I'll give this a try.

JT9D -> Boeing 747-100 => 30% overall efficiency (~64% propulsive, ~48% motor thermodynamic)
RR Trent XWB -> Airbus A350 => 40% overall efficiency (~71% propulsive, ~57% motor thermodynamic)

The propulsive improvement is due to the increase in bypass ratio from 4.8:1 to 9.6:1. Bypass ratio is the circular area covered by the turning fan divided by the gas exhaust's cross-sectional area. To convert the hot compressed air resulting from fuel combustion into propulsion, it's more efficient to turn a turbine shaft connected to a fan that pushes a lot of ambient air backward. That beats directly having the heated/compressed air escape backward, which propels the aircraft forward from the air's expansion from cooling to atmospheric temperature and from uncompressing to atmospheric pressure. The bypass ratio increase was possible because of materials improvements, which allowed fan blades to become longer and thinner. Improved materials also allowed for the blade edges to be swept instead of just being straight/flat, which raised the efficiency of the fan.

I understand the engine internals/thermodynamics less, but I can say that the structural materials improved to allow higher temperatures and pressures without melting/damaging the components. Higher temps and pressures mean the engine's turbines can extract more energy from the heated and compressed air. The pressure ratio (the pressure of air exiting the high-pressure compressor divided by the pressure of the outside atmosphere) went from 26.7:1 on the JT9D to 50:1 on the Trent XWB. Also, the JT9D has a two-shaft architecture, while the Trent XWB is three-shaft. For two-shaft, the fan and the low-or-intermediate pressure compressor are driven by the same turbine on one long shaft, so they have to run at the same rotational speed. For three-shaft, those two components are driven by different turbines on separate shafts, so those components can turn at different, more optimal speeds (slower for the fan, faster for the LP/IP compressor). Having separate shafts avoids other design constraints; on the two-shaft architecture, if you optimize the fan speed, you might need to add more LP/IP compression stages, but if you optimize the LP/IP compressor speed, you might have to reduce the diameter of the fan. The three-shaft architecture also makes the engine shorter and lighter.

Comments/corrections welcomed.

Couple of good points in here, ans a few minor clarifications I'd add:
1. The bypass ratio point is mostly on the money, the one thing I'd change there is that it's the mass flow of the bypass versus the core that you need to measure, so the ratios of the areas only works if they have the same jet efflux speed and pressure, which is not generally the case, but it's a close approximation and good visual nevertheless.
2. The compression ratio has to do with the delta-T, so the higher the pressure, the greater the efficiency.
3. One problem with increased engine efficiency is plateauing system efficiency. Higher bypass and pressure ratios involve heavier engines, which need to be carried, so there's an engineering tradeoff at the system level where it makes sense to sacrifice engine efficency for system efficiency.

True...witness the more efficient GE9X is actually slightly heavier than the GE90-115B which it replaces...and that at a slightly lower maximum rated thrust too...


Faro

From GE90 an introduction :- https://www.kimerius.com/app/download/5 ... uction.pdf

A GE90 has a mass flow at cruise of 576 kg/s, and generates 69.2 kN thrust, so from F = MA the air must be on average subject to a 120 m/s delta velocity or 433 km/h higher velocity at exit then input.

Obviously the core air will have a higher velocity delta than the by-pass air.

jambrain wrote:

From GE90 an introduction :- https://www.kimerius.com/app/download/5 ... uction.pdf

A GE90 has a mass flow at cruise of 576 kg/s, and generates 69.2 kN thrust, so from F = MA the air must be on average subject to a 120 m/s delta velocity or 433 km/h higher velocity at exit then input.

Obviously the core air will have a higher velocity delta than the by-pass air.

Your F=MA quote reminded me of another basic component to this: the energy (E) required to accelerate a mass scales as the square of the delta-V, so for a given F target you want to maximize M and minimize A to reduce E. This is why turboprops are more efficient than jets at the speeds they fly.

LH707330 wrote:

Couple of good points in here, ans a few minor clarifications I'd add:
1. The bypass ratio point is mostly on the money, the one thing I'd change there is that it's the mass flow of the bypass versus the core that you need to measure, so the ratios of the areas only works if they have the same jet efflux speed and pressure, which is not generally the case, but it's a close approximation and good visual nevertheless.
2. The compression ratio has to do with the delta-T, so the higher the pressure, the greater the efficiency.
3. One problem with increased engine efficiency is plateauing system efficiency. Higher bypass and pressure ratios involve heavier engines, which need to be carried, so there's an engineering tradeoff at the system level where it makes sense to sacrifice engine efficency for system efficiency.

Thanks for the clarifications.

Sokes wrote:

Thermodynamic efficiencies are around 55%? Seriously?
The best combined cycle gas turbines reach an efficiency of IIRC 63%. Now there may be a few percent losses from running the generator, but I find 55% thermal efficiency hard to believe.
Or are CCGTs run with less pressure/ temperature to minimize maintenance?

I just noticed that "motor thermodynamic efficiency" axis in the first graph can't be equivalent to the "thermal efficiency n(th) * transfer efficiency n(tran)" axis on the second graph, since the value for the A350's engine is 57% -- more than the 55% low-NOx limit on the second graph. I guess the transfer efficiency is ignored or applied differently on the first graph.

Never heard about aircraft engines adding an additional turbine to harness the hot exhaust. I've read about recuperator engines that use the exhaust to preheat the compressor outlet (combustor inlet) air, but apparently they're mostly useful for small engines. The Saturn/Lyulka AL-34 engine (an abandoned Soviet engine from the 1990s) was supposed to incorporate that feature. https://www.researchgate.net/figure/Tur ... _269400720

CowAnon wrote:

Sokes wrote:

Thermodynamic efficiencies are around 55%? Seriously?
The best combined cycle gas turbines reach an efficiency of IIRC 63%. Now there may be a few percent losses from running the generator, but I find 55% thermal efficiency hard to believe.
Or are CCGTs run with less pressure/ temperature to minimize maintenance?

....
Never heard about aircraft engines adding an additional turbine to harness the hot exhaust.

I didn't mean to say aircraft engines use exhaust heat to run a steam turbine. I just wondered that the gas turbine alone can reach 55% efficiency.
GE for one of its best engines says around 44% efficiency for simple cycle and 64% for combined cycle.
https://www.ge.com/power/gas/gas-turbines/9ha

It's probably best explained with reduced maintenance demand in stationary application. I still wonder about the difference in efficiency.

Sokes wrote:

CowAnon wrote:

Sokes wrote:

Thermodynamic efficiencies are around 55%? Seriously?
The best combined cycle gas turbines reach an efficiency of IIRC 63%. Now there may be a few percent losses from running the generator, but I find 55% thermal efficiency hard to believe.
Or are CCGTs run with less pressure/ temperature to minimize maintenance?

....
Never heard about aircraft engines adding an additional turbine to harness the hot exhaust.

I didn't mean to say aircraft engines use exhaust heat to run a steam turbine. I just wondered that the gas turbine alone can reach 55% efficiency.
GE for one of its best engines says around 44% efficiency for simple cycle and 64% for combined cycle.
https://www.ge.com/power/gas/gas-turbines/9ha

It's probably best explained with reduced maintenance demand in stationary application. I still wonder about the difference in efficiency.

I wouldn't be surprized if it comes to ability to condense water vapor and reduce exhaust pressure in that way. Water vapor carries a lot of energy... Probably condensation comes with a heavy condenser with water cooling shroud

kalvado wrote:

I wouldn't be surprized if it comes to ability to condense water vapor and reduce exhaust pressure in that way. Water vapor carries a lot of energy... Probably condensation comes with a heavy condenser with water cooling shroud

IIRC steam turbines condensate the steam below atmospheric pressure. But this isn't related to gas turbines reaching 55% efficiency. How should a plane engine expand the exhaust below atmosperic pressure?
On the other side atmospheric pressure in altitude is less than in stationary applications. But does this have any effect? If yes, I doubt it can cause efficiencies of 55% as compared to 44% stationairy.

The combined-cycle engines have two thermodynamic advantages:

1. As mentioned, the hot exhaust warms up water to power a steam turbine
2. Many CCGTs also use water in the compressor to reduce compressor temperatures, work required to compress the air mass, lower NOx, and allow higher fuel addition before reaching the turbine temp limits. Because they're on land, the weight and complexity is much less of an issue.

@Sokes
I have a feeling you're not "comparing apples to apples" with your efficiency questions, but I don't have an adequate explanation for them, either.

CowAnon wrote:

T54A wrote:

In layman terms, where did the efficiency improvements come from in modern turbo fans? Let’s say JT9D vs RR Trent XWB.

I guess I'll give this a try.

JT9D -> Boeing 747-100 => 30% overall efficiency (~64% propulsive, ~48% motor thermodynamic)
RR Trent XWB -> Airbus A350 => 40% overall efficiency (~71% propulsive, ~57% motor thermodynamic)

The propulsive improvement is due to the increase in bypass ratio from 4.8:1 to 9.6:1. Bypass ratio is the circular area covered by the turning fan divided by the gas exhaust's cross-sectional area. To convert the hot compressed air resulting from fuel combustion into propulsion, it's more efficient to turn a turbine shaft connected to a fan that pushes a lot of ambient air backward. That beats directly having the heated/compressed air escape backward, which propels the aircraft forward from the air's expansion from cooling to atmospheric temperature and from uncompressing to atmospheric pressure. The bypass ratio increase was possible because of materials improvements, which allowed fan blades to become longer and thinner. Improved materials also allowed for the blade edges to be swept instead of just being straight/flat, which raised the efficiency of the fan.

I understand the engine internals/thermodynamics less, but I can say that the structural materials improved to allow higher temperatures and pressures without melting/damaging the components. Higher temps and pressures mean the engine's turbines can extract more energy from the heated and compressed air. The pressure ratio (the pressure of air exiting the high-pressure compressor divided by the pressure of the outside atmosphere) went from 26.7:1 on the JT9D to 50:1 on the Trent XWB. Also, the JT9D has a two-shaft architecture, while the Trent XWB is three-shaft. For two-shaft, the fan and the low-or-intermediate pressure compressor are driven by the same turbine on one long shaft, so they have to run at the same rotational speed. For three-shaft, those two components are driven by different turbines on separate shafts, so those components can turn at different, more optimal speeds (slower for the fan, faster for the LP/IP compressor). Having separate shafts avoids other design constraints; on the two-shaft architecture, if you optimize the fan speed, you might need to add more LP/IP compression stages, but if you optimize the LP/IP compressor speed, you might have to reduce the diameter of the fan. The three-shaft architecture also makes the engine shorter and lighter.

Comments/corrections welcomed.

Very good summary. From a thermodynamic perspective the higher bypass ratio allows more air to bypass the engine core. The greater the temperature difference between the engine core and the bypassed air the more "work" the engine can do (that is, efficiency increases).

Steam engines, gas turbines, and internal combustion engines all work under the identical thermodynamic principles. That is why the big breakthroughs in jet engine tech are in the material sector. The hotter the engine core the better.

That is also why jet engines perform less efficiently in hot conditions. Think DXB in summer when it is 45 c. The greatest volume of colder bypass air versus the hottest engine core the materials can tolerant is the ideal.

ElroyJetson wrote:

That is also why jet engines perform less efficiently in hot conditions. Think DXB in summer when it is 45 c. The greatest volume of colder bypass air versus the hottest engine core the materials can tolerant is the ideal.

Isn't that primarily due to reduced air density? Thurst is M.dot, carnot efficiency derating should be in the 1.5 .. 2% domain.

RR years ago shew a table of thermal efficiency vs optimal bypass ration. ( link lost, sorry )

WIederling wrote:

ElroyJetson wrote:

That is also why jet engines perform less efficiently in hot conditions. Think DXB in summer when it is 45 c. The greatest volume of colder bypass air versus the hottest engine core the materials can tolerant is the ideal.

Isn't that primarily due to reduced air density? Thurst is M.dot, carnot efficiency derating should be in the 1.5 .. 2% domain.

RR years ago shew a table of thermal efficiency vs optimal bypass ration. ( link lost, sorry )

Both actually. Colder denser air increases jet turbine efficiency. Warmer less dense air decreases efficiency. The thermodynamic principles I mentioned above is basically describing entropy. Warmer always must go to colder, never the reverse.

Steam engines, jet turbines, and internal combustion engines all work on the identical principle. Heat travels from one chamber that is hot to another chamber that is cooler. The greater the temperature variance between the chambers the more "work" the engine can do.

In jet engines colder air passes over and around a hot engine core. Core temps in jet turbines have increased about 1,000 degrees F since the early turbo jet era. Hence a huge increase in efficiency. That is why the next big breakthroughs in engine tech will be in the materials sector. CMC is already being incorporated in engine components, its use will increase. The problem with CMC is it is brittle and difficult to machine. If there was a ceramic material that could be machined like steel and be as pliable as steel that would be an enormous breakthrough.

Engineers are taught themodyanmic principles, but it sort of like their organic chemistry. Classes to be endured and quickly left.

LH707330 wrote:

The combined-cycle engines have two thermodynamic advantages:

1. As mentioned, the hot exhaust warms up water to power a steam turbine

This. There is no rankine bottoming cycle for aviation GT's.

I find 55% to be unrealistically high for an aircraft GT without a bottoming cycle, as the bottoming cycle contributes around 35% to overall efficiency of a CC gas turbine. The current efficiency record is around 65%-66% for a CC GT with a 1600 deg C turbine inlet temp. Supposedly solid oxide fuel cells in the bottoming cycle could raise that to almost 80% (again, not applicable to aircraft though unless the GT is powering an electric propulsion system)

Something else to note: The efficiency of the JT9D was improved by 15% with the PW4000, yet bypass ratio was not increased at all.

Heck, the TF-39 had an 8-to-1 BPR and the CF6-80 is only 5 to1

ElroyJetson wrote:

Very good summary.

Thanks!

ElroyJetson wrote:

In jet engines colder air passes over and around a hot engine core. Core temps in jet turbines have increased about 1,000 degrees F since the early turbo jet era. Hence a huge increase in efficiency. That is why the next big breakthroughs in engine tech will be in the materials sector. CMC is already being incorporated in engine components, its use will increase. The problem with CMC is it is brittle and difficult to machine. If there was a ceramic material that could be machined like steel and be as pliable as steel that would be an enormous breakthrough.

How much more of an increase in temperature is possible/probable in the future?

744SPX wrote:

Heck, the TF-39 had an 8-to-1 BPR and the CF6-80 is only 5 to1

Interesting. I'm not familiar with a lot of older engines and aircraft, so I didn't know that a BPR 8 turbofan was available in 1964. It makes the advances in turbofan technology since that time seem less impressive. Why did GE have to reduce the BPR in its CF6 followup?

CowAnon wrote:

ElroyJetson wrote:

Very good summary.

Thanks!

ElroyJetson wrote:

In jet engines colder air passes over and around a hot engine core. Core temps in jet turbines have increased about 1,000 degrees F since the early turbo jet era. Hence a huge increase in efficiency. That is why the next big breakthroughs in engine tech will be in the materials sector. CMC is already being incorporated in engine components, its use will increase. The problem with CMC is it is brittle and difficult to machine. If there was a ceramic material that could be machined like steel and be as pliable as steel that would be an enormous breakthrough.

How much more of an increase in temperature is possible/probable in the future?

744SPX wrote:

Heck, the TF-39 had an 8-to-1 BPR and the CF6-80 is only 5 to1

Interesting. I'm not familiar with a lot of older engines and aircraft, so I didn't know that a BPR 8 turbofan was available in 1964. It makes the advances in turbofan technology since that time seem less impressive. Why did GE have to reduce the BPR in its CF6 followup?

The highest grade metal alloys can withstand temps up to 1300 C. The best ceramics can withstand temps up to 1600 C. That is roughly 3000 F. The turbine intake on jet turbines can reach 3600 F at full power. Melting is prevented by micro pores in the metal or ceramic material. The fact that the best ceramic materials can take temps 300 C hotter is hugely significant. That is over 570 F difference.
Again, CMC is being incorporated in jet turbines, but it is brittle and difficult to machine. As machining gets better expect to see more of these components in jet turbines and subsequent increases in efficiency.

GE for example is spending 1.5 billion in CMC and they see the material as a potential major leap in jet turbine technology. This link is a great article on where the technology may be going.


https://blog.geaviation.com/technology/ ... 20coatings.


Key quote, The CMC benefits are seductive, but industrializing this sophisticated material system has posed a huge challenge to private industry for decades. Difficult to fabricate, CMCs also have brittle properties. The U.S. government has funded CMC research since the early 1970s, and GE scientists have wrestled with the technology ever since then. In the 1980s, GE pursued CMCs for large ground-based gas turbines and filed for its first CMC patent in 1986. Within 25 years, the company successfully ran CMC turbine shrouds in multiple industrial gas turbine applications.

One other thing I want to emphasize. When you talk with engineers they rarely if ever mention thermodynamics. Honestly I do not think most understand it that well, and they tend to get lost in the micro engineering details.

Bottom line: The greater the temperature variance between the hot chamber and cool chamber of an engine the more "work" it can do (i.e. the efficiency increases). This is true of steam engines, internal combustion engines, and jet turbines, It is an immutable Law of Physics.

By the way, you can think of a hurricane as a giant engine. A hurricanes power increases the greater the temperature difference between hot surface water, and cold upper atmosphere air. Per the Law of Thermodynamics hot must always go to cold. Never the reverse. If you understand that at a fundamental level you can understand precisely how every jet turbine works.

CowAnon wrote:

Ultrafan was said to be 25% more efficient than the first RR Trent turbofans.

That’s since around 1985, typically the improvements have been less than 1% in TSFC over per year.

Faro wrote:

the more efficient GE9X is actually slightly heavier than the GE90-115B

It’s around 10% heavier, as well as physically larger which means more drag.

744SPX wrote:

Heck, the TF-39 had an 8-to-1 BPR and the CF6-80 is only 5 to1

Interesting. I'm not familiar with a lot of older engines and aircraft, so I didn't know that a BPR 8 turbofan was available in 1964. It makes the advances in turbofan technology since that time seem less impressive. Why did GE have to reduce the BPR in its CF6 followup?[/quote]

I'm not really sure, although the TF-39 uses a rather unique and complex fan design to achieve that BPR and maybe they just wanted to simplify...

ElroyJetson wrote:

One other thing I want to emphasize. When you talk with engineers they rarely if ever mention thermodynamics. Honestly I do not think most understand it that well, and they tend to get lost in the micro engineering details.

Bottom line: The greater the temperature variance between the hot chamber and cool chamber of an engine the more "work" it can do (i.e. the efficiency increases). This is true of steam engines, internal combustion engines, and jet turbines, It is an immutable Law of Physics.

By the way, you can think of a hurricane as a giant engine. A hurricanes power increases the greater the temperature difference between hot surface water, and cold upper atmosphere air. Per the Law of Thermodynamics hot must always go to cold. Never the reverse. If you understand that at a fundamental level you can understand precisely how every jet turbine works.

Gas turbine system engineering very heavily relies on the thermodynamics.

The summaries so far have been excellent. I will point out a few details:
1. Surface area is drag. Part of the efficiency reduction is fewer stages to do the same work. A big reduction is the combustor. The current combustor length is half prior. This is one reason there is such an obsession to reduce part count is to reduce surface area in the engines.
2. Mach numbers are much more optimised. On non-GTF engines the scimitar fan blades allow much higher low turbine mach numbers which really improves efficiency. On GTFs, we still are not near the optimal, but baby steps.
3. High spools have very far to go to achieve optimal Mach #. Alas, as the PW1100G showed, we need to develop seals, bearing, and rotor dynamics for accelerated RPM (needed to achieve the Mach #s).
4. Materials (besides CMCs). All new engines are gaining weight. A higher pressure requires a thicker pressure vessel (casing) booth for hoop stress and because increasing temperature reduces material strength

How do we break this cycle? Fuel cells. Eventually, we won't have gas turbine engines, but fuel cells powering fans. But that is just my opinion and decades away.

Lightsaber

zeke wrote:

CowAnon wrote:

Ultrafan was said to be 25% more efficient than the first RR Trent turbofans.

That’s since around 1985, typically the improvements have been less than 1% in TSFC over per year.

Thanks for that tip. I looked around and found that the Trent 600, which was to go on the McDonnell Douglas MD-11 before the engine was canceled, had a TSFC of 0.59 lb/lbf/hr. If that's the earliest Trent, then the Ultrafan would come in at 0.44 lb/lbf/hr. To me that sounds more realistic than 0.42.

lightsaber wrote:

2. Mach numbers are much more optimised. On non-GTF engines the scimitar fan blades allow much higher low turbine mach numbers which really improves efficiency. On GTFs, we still are not near the optimal, but baby steps.
3. High spools have very far to go to achieve optimal Mach #. Alas, as the PW1100G showed, we need to develop seals, bearing, and rotor dynamics for accelerated RPM (needed to achieve the Mach #s).

Interesting to hear that compressor & turbine speeds are still not where they could/should be.

ElroyJetson wrote:

One other thing I want to emphasize. When you talk with engineers they rarely if ever mention thermodynamics. Honestly I do not think most understand it that well, and they tend to get lost in the micro engineering details.

Bottom line: The greater the temperature variance between the hot chamber and cool chamber of an engine the more "work" it can do (i.e. the efficiency increases). This is true of steam engines, internal combustion engines, and jet turbines, It is an immutable Law of Physics.

By the way, you can think of a hurricane as a giant engine. A hurricanes power increases the greater the temperature difference between hot surface water, and cold upper atmosphere air. Per the Law of Thermodynamics hot must always go to cold. Never the reverse. If you understand that at a fundamental level you can understand precisely how every jet turbine works.

Thermodynamics is mentioned early in a design. New engines that rely on variable cycle technology are detailed thermodynamics analysis.

However practical aspects often result in severe compromises to the thermodynamics. For example, decades ago I was part if a large engine design team for an engine not taken up. In the design study, we found the costs skyrocketed for manufacturing the engine above a very specific fan diameter. This was due to the lowest cost vendor having automated equipment up to a certain diameter and there still isn't enough volume to automate as effectively at too much larger. It made for a cost difference of $1 million per engine which just destroyed the economic model of going only a little larger.

Now larger tools have been bought, but the PW1100G and LEAP set the standard. Pratt made vendors buy for the MoM, but not for GE9x or Txwb. e.g., at vendors, the Txwb, even for small parts I would have thought were fine in existing tooling, the vendors needed new tooling. So much that RR bought the automated tooling and the vendors pay rent for non Txwb work on the tools and no excuses, Txwb parts go before all others. The tools are so efficient that 40% of the work on them (my best estimate) is rented time.

Some vendors wised up and expanded factories to where all non-Txwb work is on vendor owned automated tooling.

Jet engines are squeeze-bang-blow. It is all about getting out inefficiencies and weight. Thermodynamics will play a key role in:
1. Variable turbine cooling (cut down during cruise, ~3% efficiency gain in LEAP and GE8x only so far).
2. Variable fan pitch. A 2% to 3% gain for a lot of weight. Bummer all those Hamilton-Standard engineers who developed constant speed props are long dead. There are several good WW2 concepts to model.
4. Variable fan nozzle. Less gain, takes 3+ hours if mission to pay for itself (fuel burn break even in a half hour of cruise or less, but then manufacturing and maintenance have to be paid for in fuel savings during cruise).
5. variable stators in the turbine. Obviously only low turbine due to heat.
6. More turbine stages allows higher pressure ratios. GE's foamed nickel helps reduce weight, allowing lighter blades. This shifts the thermodynamic optimum bypass ratio and pressure ratio by allowing more low turbine stages at less weight. For lighter blades mean a lighter rotor, lighter blade out containment,

It isn't just component efficiency, but weight and nacelle drag are always part of the design optimum.

A few examples might help. Please recall all engines must set many design parameters years before the fuselage.

For the PW1500G, the original mission was a 1 hour optimization. Sorry, no link, take this as rumor. This set the thermodynamics (bypass ratio and pressure ratio). Later Bombardier switched to a 90 minute mission. Pratt acheived this, in my opinion, by increasing high turbine cooling and using margin in combustor and low turbine life (increased maintenance costs). I believed they used up too much combustor margin and this is why they must improve overhaul intervals, but that is my speculation.


The A320NEO set a 2 hour mission. This allowed more weight and bypass ratio, which hurts climb fuel burn. GE broke part of the compromise by variable turbine cooling, but the mission sets optimization.

Lightsaber

Lightsaber,

Variable fan blades pitch, only 2 to 3% gain? Is that less than what was gained when props were made variable pitch? If so, how come there would be a difference?

I remember being given an ADP for Advanced Ducted Prop brochure by a Pratt folk, at Paris Air Show in the late '80s or early '90s. I wondered what had happened to that project. I guess part of it is in today's GTF, but what happened to the variable pitch that was to make reversing air flow so easy?

lightsaber wrote:

However practical aspects often result in severe compromises to the thermodynamics. For example, decades ago I was part if a large engine design team for an engine not taken up. In the design study, we found the costs skyrocketed for manufacturing the engine above a very specific fan diameter. This was due to the lowest cost vendor having automated equipment up to a certain diameter and there still isn't enough volume to automate as effectively at too much larger. It made for a cost difference of $1 million per engine which just destroyed the economic model of going only a little larger.

Aircraft engine production seems like such a horrible business to be in. Face huge upfront costs to create much of the fuel savings for the airlines and aircraft OEMs, yet get squeezed by the plane OEMs to the point of taking decades to return a profit, but receive little credit when things go smoothly and lots of blame if things go badly!

6. More turbine stages allows higher pressure ratios. GE's foamed nickel helps reduce weight, allowing lighter blades. This shifts the thermodynamic optimum bypass ratio and pressure ratio by allowing more low turbine stages at less weight. For lighter blades mean a lighter rotor, lighter blade out containment,

Quick survey of Wikipedia finds 7 to be the highest number of low-turbine stages in an engine. How much higher can that go to? And does this mean that GE can reach a bypass ratio of 12-15 with its engines without having to resort to a gearbox?

Ultrafan gear system has been run up to 85,000 hp. Clearly a very different design to PW's gear - so no patent lawyers necessary!

"(The gear system) is a planetary-style gearbox with a ring gear on the outside and five planet gears inside, rotating around a central sun gear. The design drives the fan from a centrally mounted planet carrier unlike the star-style gear system used in Pratt & Whitney’s geared turbofan."

https://aviationweek.com/air-transport/ ... wer-record

https://www.airlinerwatch.com/2021/09/r ... -tops.html

JerseyFlyer wrote:

Ultrafan gear system has been run up to 85,000 hp. Clearly a very different design to PW's gear - so no patent lawyers necessary!

"(The gear system) is a planetary-style gearbox with a ring gear on the outside and five planet gears inside, rotating around a central sun gear. The design drives the fan from a centrally mounted planet carrier unlike the star-style gear system used in Pratt & Whitney’s geared turbofan."

https://aviationweek.com/air-transport/ ... wer-record

https://www.airlinerwatch.com/2021/09/r ... -tops.html

That's a nice achievement. Although I've never understood Roll's strategy. If they didn't bother with Ultrafan, wouldn't the company's Trent engines still have a competitive-to-strong position in the widebody engine space? And a previous article (can't find it right now) at least hinted that the Ultrafan might not be able to scale down to the lower thrust levels of narrowbody engines?

Also, since the early 1990s Allison has produced the T406 and AE2100 engines (turboshaft and turboprop derivatives of the military T56 engine), which have gearboxes that last 30,000 hours before replacement. While those are only 6,000-horsepower engines, Allison's gearboxes are closely related to one in the 578-DX propfan demonstrator that flew in 1989, which had a design that scaled to 25,000 hp. Rolls bought out Allison in the mid-1990s, so RR has had available a high-reliability gearbox design that could power many narrowbody aircraft, but still hasn't been able to come up with a proper narrowbody engine competitor in the last 25 years.

CowAnon wrote:

Never heard about aircraft engines adding an additional turbine to harness the hot exhaust. I've read about recuperator engines that use the exhaust to preheat the compressor outlet (combustor inlet) air, but apparently they're mostly useful for small engines. The Saturn/Lyulka AL-34 engine (an abandoned Soviet engine from the 1990s) was supposed to incorporate that feature. https://www.researchgate.net/figure/Tur ... _269400720

CFM Details Open-Fan Plan For Next-gen Engine (Aviation Week, 6/25/2021)

The RISE open fan will include a new compact high-pressure core to boost thermodynamic efficiency, as well as a recuperating system to preheat combustion air with waste heat from the exhaust.

That's for an open-rotor engine, but I don't think there's anything that prevents a recuperator from being added to turbofan engines.

Sokes wrote:

CowAnon wrote:

Sokes wrote:

Thermodynamic efficiencies are around 55%? Seriously?
The best combined cycle gas turbines reach an efficiency of IIRC 63%. Now there may be a few percent losses from running the generator, but I find 55% thermal efficiency hard to believe.
Or are CCGTs run with less pressure/ temperature to minimize maintenance?

....
Never heard about aircraft engines adding an additional turbine to harness the hot exhaust.

I didn't mean to say aircraft engines use exhaust heat to run a steam turbine. I just wondered that the gas turbine alone can reach 55% efficiency.
GE for one of its best engines says around 44% efficiency for simple cycle and 64% for combined cycle.
https://www.ge.com/power/gas/gas-turbines/9ha

It's probably best explained with reduced maintenance demand in stationary application. I still wonder about the difference in efficiency.

Could the difference also be due, in part, to the fuel - Natural Gas vs. Jet-A? I'm not familiar with power turbines, but are they compressing the air for combustion like an engine? Even if they are, I would be curious to what extent.

bkflyguy wrote:

Sokes wrote:

CowAnon wrote:

....
Never heard about aircraft engines adding an additional turbine to harness the hot exhaust.

I didn't mean to say aircraft engines use exhaust heat to run a steam turbine. I just wondered that the gas turbine alone can reach 55% efficiency.
GE for one of its best engines says around 44% efficiency for simple cycle and 64% for combined cycle.
https://www.ge.com/power/gas/gas-turbines/9ha

It's probably best explained with reduced maintenance demand in stationary application. I still wonder about the difference in efficiency.

Could the difference also be due, in part, to the fuel - Natural Gas vs. Jet-A? I'm not familiar with power turbines, but are they compressing the air for combustion like an engine? Even if they are, I would be curious to what extent.

Compressing? Yes, even more so than typical engines. For example, the GE9X overall pressure ratio is 60:1.

The average modern gasoline engine compression ratio is 11:1 (highest I’ve seen is 13.5:1) and Diesel engines are about 16 or 16.5:1.

Compression ratio isn’t a good measure of efficiency comparison between to different engine types. It’s rather helpful within the same engine type. In reciprocating engines, expansion ratio is a better indicator of overall efficiency.

lightsaber wrote:

ElroyJetson wrote:

One other thing I want to emphasize. When you talk with engineers they rarely if ever mention thermodynamics. Honestly I do not think most understand it that well, and they tend to get lost in the micro engineering details.

Bottom line: The greater the temperature variance between the hot chamber and cool chamber of an engine the more "work" it can do (i.e. the efficiency increases). This is true of steam engines, internal combustion engines, and jet turbines, It is an immutable Law of Physics.

By the way, you can think of a hurricane as a giant engine. A hurricanes power increases the greater the temperature difference between hot surface water, and cold upper atmosphere air. Per the Law of Thermodynamics hot must always go to cold. Never the reverse. If you understand that at a fundamental level you can understand precisely how every jet turbine works.

Thermodynamics is mentioned early in a design. New engines that rely on variable cycle technology are detailed thermodynamics analysis.

However practical aspects often result in severe compromises to the thermodynamics. For example, decades ago I was part if a large engine design team for an engine not taken up. In the design study, we found the costs skyrocketed for manufacturing the engine above a very specific fan diameter. This was due to the lowest cost vendor having automated equipment up to a certain diameter and there still isn't enough volume to automate as effectively at too much larger. It made for a cost difference of $1 million per engine which just destroyed the economic model of going only a little larger.

Now larger tools have been bought, but the PW1100G and LEAP set the standard. Pratt made vendors buy for the MoM, but not for GE9x or Txwb. e.g., at vendors, the Txwb, even for small parts I would have thought were fine in existing tooling, the vendors needed new tooling. So much that RR bought the automated tooling and the vendors pay rent for non Txwb work on the tools and no excuses, Txwb parts go before all others. The tools are so efficient that 40% of the work on them (my best estimate) is rented time.

Some vendors wised up and expanded factories to where all non-Txwb work is on vendor owned automated tooling.

Jet engines are squeeze-bang-blow. It is all about getting out inefficiencies and weight. Thermodynamics will play a key role in:
1. Variable turbine cooling (cut down during cruise, ~3% efficiency gain in LEAP and GE8x only so far).
2. Variable fan pitch. A 2% to 3% gain for a lot of weight. Bummer all those Hamilton-Standard engineers who developed constant speed props are long dead. There are several good WW2 concepts to model.
4. Variable fan nozzle. Less gain, takes 3+ hours if mission to pay for itself (fuel burn break even in a half hour of cruise or less, but then manufacturing and maintenance have to be paid for in fuel savings during cruise).
5. variable stators in the turbine. Obviously only low turbine due to heat.
6. More turbine stages allows higher pressure ratios. GE's foamed nickel helps reduce weight, allowing lighter blades. This shifts the thermodynamic optimum bypass ratio and pressure ratio by allowing more low turbine stages at less weight. For lighter blades mean a lighter rotor, lighter blade out containment,

It isn't just component efficiency, but weight and nacelle drag are always part of the design optimum.

A few examples might help. Please recall all engines must set many design parameters years before the fuselage.

For the PW1500G, the original mission was a 1 hour optimization. Sorry, no link, take this as rumor. This set the thermodynamics (bypass ratio and pressure ratio). Later Bombardier switched to a 90 minute mission. Pratt acheived this, in my opinion, by increasing high turbine cooling and using margin in combustor and low turbine life (increased maintenance costs). I believed they used up too much combustor margin and this is why they must improve overhaul intervals, but that is my speculation.


The A320NEO set a 2 hour mission. This allowed more weight and bypass ratio, which hurts climb fuel burn. GE broke part of the compromise by variable turbine cooling, but the mission sets optimization.

Lightsaber

Love the post. Wish we had a like button.

But I thought that the big gains were to be made in allowing ever higher temps in the core, which allowed for higher bypass ratios. Is that not a fair 12 word oversimplified summary of progress to date, and our best hope for the future too?

kitplane01 wrote:

lightsaber wrote:

ElroyJetson wrote:

One other thing I want to emphasize. When you talk with engineers they rarely if ever mention thermodynamics. Honestly I do not think most understand it that well, and they tend to get lost in the micro engineering details.

Bottom line: The greater the temperature variance between the hot chamber and cool chamber of an engine the more "work" it can do (i.e. the efficiency increases). This is true of steam engines, internal combustion engines, and jet turbines, It is an immutable Law of Physics.

By the way, you can think of a hurricane as a giant engine. A hurricanes power increases the greater the temperature difference between hot surface water, and cold upper atmosphere air. Per the Law of Thermodynamics hot must always go to cold. Never the reverse. If you understand that at a fundamental level you can understand precisely how every jet turbine works.

Thermodynamics is mentioned early in a design. New engines that rely on variable cycle technology are detailed thermodynamics analysis.

However practical aspects often result in severe compromises to the thermodynamics. For example, decades ago I was part if a large engine design team for an engine not taken up. In the design study, we found the costs skyrocketed for manufacturing the engine above a very specific fan diameter. This was due to the lowest cost vendor having automated equipment up to a certain diameter and there still isn't enough volume to automate as effectively at too much larger. It made for a cost difference of $1 million per engine which just destroyed the economic model of going only a little larger.

Now larger tools have been bought, but the PW1100G and LEAP set the standard. Pratt made vendors buy for the MoM, but not for GE9x or Txwb. e.g., at vendors, the Txwb, even for small parts I would have thought were fine in existing tooling, the vendors needed new tooling. So much that RR bought the automated tooling and the vendors pay rent for non Txwb work on the tools and no excuses, Txwb parts go before all others. The tools are so efficient that 40% of the work on them (my best estimate) is rented time.

Some vendors wised up and expanded factories to where all non-Txwb work is on vendor owned automated tooling.

Jet engines are squeeze-bang-blow. It is all about getting out inefficiencies and weight. Thermodynamics will play a key role in:
1. Variable turbine cooling (cut down during cruise, ~3% efficiency gain in LEAP and GE8x only so far).
2. Variable fan pitch. A 2% to 3% gain for a lot of weight. Bummer all those Hamilton-Standard engineers who developed constant speed props are long dead. There are several good WW2 concepts to model.
4. Variable fan nozzle. Less gain, takes 3+ hours if mission to pay for itself (fuel burn break even in a half hour of cruise or less, but then manufacturing and maintenance have to be paid for in fuel savings during cruise).
5. variable stators in the turbine. Obviously only low turbine due to heat.
6. More turbine stages allows higher pressure ratios. GE's foamed nickel helps reduce weight, allowing lighter blades. This shifts the thermodynamic optimum bypass ratio and pressure ratio by allowing more low turbine stages at less weight. For lighter blades mean a lighter rotor, lighter blade out containment,

It isn't just component efficiency, but weight and nacelle drag are always part of the design optimum.

A few examples might help. Please recall all engines must set many design parameters years before the fuselage.

For the PW1500G, the original mission was a 1 hour optimization. Sorry, no link, take this as rumor. This set the thermodynamics (bypass ratio and pressure ratio). Later Bombardier switched to a 90 minute mission. Pratt acheived this, in my opinion, by increasing high turbine cooling and using margin in combustor and low turbine life (increased maintenance costs). I believed they used up too much combustor margin and this is why they must improve overhaul intervals, but that is my speculation.


The A320NEO set a 2 hour mission. This allowed more weight and bypass ratio, which hurts climb fuel burn. GE broke part of the compromise by variable turbine cooling, but the mission sets optimization.

Lightsaber

Love the post. Wish we had a like button.

But I thought that the big gains were to be made in allowing ever higher temps in the core, which allowed for higher bypass ratios. Is that not a fair 12 word oversimplified summary of progress to date, and our best hope for the future too?

No, core temps and BPR are independent of each other, AFAICT. For example, the Rolls-Royce Trent XWB only has a bypass ratio of 9.6, even though it can handle temperatures over 2,000 K (1,727 C), and apparently has room for higher diameter (leading to increased BPR) on the Airbus A350. The Pratt & Whitney GTF has a higher bypass ratio of 12, but its max turbine temps don't even reach 1,100 C.

Limiting issues for turbofan BPR are

• clearance issues (e.g., the CFM LEAP variant having smaller BPR on the short-legged 737 than on planes with longer landing gear),
• mismatches in ideal rotational speed on fan vs. turbine for direct-drive engines (which forces the move from direct-drive to geared engines),
• gearbox reduction ratios (only 3:1 for PWGTF, but 4:1 to 5:1 being more desirable) for geared engines, and ultimately
• the cowling weight.

EDIT: The link says exotic materials are required for a higher gear reduction ratio. Turboprop engines reach reduction ratios of 10 or above using multi-stage gearboxes, so that point confuses me.

CowAnon wrote:

kitplane01 wrote:

lightsaber wrote:

Thermodynamics is mentioned early in a design. New engines that rely on variable cycle technology are detailed thermodynamics analysis.

However practical aspects often result in severe compromises to the thermodynamics. For example, decades ago I was part if a large engine design team for an engine not taken up. In the design study, we found the costs skyrocketed for manufacturing the engine above a very specific fan diameter. This was due to the lowest cost vendor having automated equipment up to a certain diameter and there still isn't enough volume to automate as effectively at too much larger. It made for a cost difference of $1 million per engine which just destroyed the economic model of going only a little larger.

Now larger tools have been bought, but the PW1100G and LEAP set the standard. Pratt made vendors buy for the MoM, but not for GE9x or Txwb. e.g., at vendors, the Txwb, even for small parts I would have thought were fine in existing tooling, the vendors needed new tooling. So much that RR bought the automated tooling and the vendors pay rent for non Txwb work on the tools and no excuses, Txwb parts go before all others. The tools are so efficient that 40% of the work on them (my best estimate) is rented time.

Some vendors wised up and expanded factories to where all non-Txwb work is on vendor owned automated tooling.

Jet engines are squeeze-bang-blow. It is all about getting out inefficiencies and weight. Thermodynamics will play a key role in:
1. Variable turbine cooling (cut down during cruise, ~3% efficiency gain in LEAP and GE8x only so far).
2. Variable fan pitch. A 2% to 3% gain for a lot of weight. Bummer all those Hamilton-Standard engineers who developed constant speed props are long dead. There are several good WW2 concepts to model.
4. Variable fan nozzle. Less gain, takes 3+ hours if mission to pay for itself (fuel burn break even in a half hour of cruise or less, but then manufacturing and maintenance have to be paid for in fuel savings during cruise).
5. variable stators in the turbine. Obviously only low turbine due to heat.
6. More turbine stages allows higher pressure ratios. GE's foamed nickel helps reduce weight, allowing lighter blades. This shifts the thermodynamic optimum bypass ratio and pressure ratio by allowing more low turbine stages at less weight. For lighter blades mean a lighter rotor, lighter blade out containment,

It isn't just component efficiency, but weight and nacelle drag are always part of the design optimum.

A few examples might help. Please recall all engines must set many design parameters years before the fuselage.

For the PW1500G, the original mission was a 1 hour optimization. Sorry, no link, take this as rumor. This set the thermodynamics (bypass ratio and pressure ratio). Later Bombardier switched to a 90 minute mission. Pratt acheived this, in my opinion, by increasing high turbine cooling and using margin in combustor and low turbine life (increased maintenance costs). I believed they used up too much combustor margin and this is why they must improve overhaul intervals, but that is my speculation.


The A320NEO set a 2 hour mission. This allowed more weight and bypass ratio, which hurts climb fuel burn. GE broke part of the compromise by variable turbine cooling, but the mission sets optimization.

Lightsaber

Love the post. Wish we had a like button.

But I thought that the big gains were to be made in allowing ever higher temps in the core, which allowed for higher bypass ratios. Is that not a fair 12 word oversimplified summary of progress to date, and our best hope for the future too?

No, core temps and BPR are independent of each other, AFAICT. For example, the Rolls-Royce Trent XWB only has a bypass ratio of 9.6, even though it can handle temperatures over 2,000 K (1,727 C), and apparently has room for higher diameter (leading to increased BPR) on the Airbus A350. The Pratt & Whitney GTF has a higher bypass ratio of 12, but its max turbine temps don't even reach 1,100 C.

Limiting issues for turbofan BPR are

• clearance issues (e.g., the CFM LEAP variant having smaller BPR on the short-legged 737 than on planes with longer landing gear),
• mismatches in ideal rotational speed on fan vs. turbine for direct-drive engines (which forces the move from direct-drive to geared engines),
• gearbox reduction ratios (only 3:1 for PWGTF, but 4:1 to 5:1 being more desirable) for geared engines, and ultimately
• the cowling weight.

EDIT: The link says exotic materials are required for a higher gear reduction ratio. Turboprop engines reach reduction ratios of 10 or above using multi-stage gearboxes, so that point confuses me.

I wonder if I was not clear.

I believe mathematical optimum BPR is a function of both core temp, which raises optimum BPR, and aircraft speed, which lowers optimum BPR, and other stuff. And of course there are practical issues which can lower the actual BPR used in real engines.

My claim is not that every engine with a higher max turbine inlet temp has a higher bypass ratio.

Never heard about aircraft engines adding an additional turbine to harness the hot exhaust. I've read about recuperator engines that use the exhaust to preheat the compressor outlet (combustor inlet) air, but apparently they're mostly useful for small engines. The Saturn/Lyulka AL-34 engine (an abandoned Soviet engine from the 1990s) was supposed to incorporate that feature. https://www.researchgate.net/figure/Tur ... _269400720

CFM Details Open-Fan Plan For Next-gen Engine (Aviation Week, 6/25/2021)

The RISE open fan will include a new compact high-pressure core to boost thermodynamic efficiency, as well as a recuperating system to preheat combustion air with waste heat from the exhaust.

That's for an open-rotor engine, but I don't think there's anything that prevents a recuperator from being added to turbofan engines.

A second media source confirms the CFM RISE will have a waste heat harvesting system:

Safran waiting for airframers before any new engine launch, says chief Andries (FlightGlobal, 12/2/2021)

Aside from the 4m (13ft)-diameter composite open fan – twice as large as the fan on the Leap – key technologies aiding the performance of RISE will be a “lighter, more compact” high-speed booster and low-pressure system, hybrid-electric assistance, and an “advanced waste heat recovery system”, says Delphine Dijoud, executive manager, CFM RISE programme, systems engineering at Safran Aircraft Engines.

The hybrid-electric system will “optimise engine performance” by “providing additional electric thrust” when needed, and also generate electricity both for itself and the rest of the aircraft, she says.

Considering the earlier discussion above combined-cycle efficiency, maybe the waste heat system could have an equal or larger impact on SFC reduction than even the open fan does? Narrowbody planes would have huge range if the CFM RISE ever gets adopted.