In a given jet engine, is the high pressure turbine more efficient at extracting work than the low pressure turbine (notwithstanding the effect of the HPT)?
IOW, is there a reason to prefer having just the HPT rather than just the LPT?
If a jet engine only used a LPT, what effect would that have on the engine’s operation and output? Would the LPT be unable to spin the compressor to a sufficient speed?
I’m not an expert, but…
The high- vs. low- designation comes from the fact that work is extracted from the exhaust by expanding it. You couldn’t have just an LPT, because by its very nature, the first expansion stage must be a HPT–after all, it is just after the combustion chamber, and the exhaust is at high pressure.
I guess you might ask why there are multiple stages at all. There are lots of reasons, but physical and thermal loading both play a part. A single stage would probably be adequate if you had unobtainium at your disposal.
The fan on a turbofan is driven by the LPT due to the way the driveshaft is nested. The LPT is rearmost and drives the fan and part of the compressor, which is frontmost. The HPT drives the rest of the compressor. No other way to do it without massively increasing the complexity.
More generally, turbines are at their most efficient when they expand all the way to ambient temperature and pressure. However, the optimal design of the turbine may change as these values are reduced. Combined-Cycle Power Plants use two independent turbines to extract the most energy from the heat source. They have different designs, and either one on its own wouldn’t be efficient.
The turbine doesn’t need to be *just after *the combustion chamber. You could put the turbine away from the combustion chamber by putting it some distance into the nozzle.
Sure. But the fluid pressure will be the same no matter how far back it is (unless it gets sent through a de Laval nozzle). It takes a turbine to have a pressure drop, so the first turbine the exhaust encounters will be the HPT.
First some terminology:[ul][li]A turbine “stage” is a single layer of fixed blades and an adjacent single layer of rotating blades. Conventions differ on whether the fixed set is immediately upstream or downstream of the rotating set. [/li][li]Multiple stages can be coupled together on a common shaft to rotate at a common RPM. Such a group is typically called a “turbine”. If there is more than one turbine in an engine they’re usually numbered 1, 2, 3 from upstream to downstream or labeled high-, medium- / intermediate-, or low-pressure = HPT, MPT/IPT, and LPT. [/li][li]The whole shebang, from the downstream end of the combustor to the upstream end of the exhaust pipe of the engine is a “turbine system.”[/ul][/li]
Many pure turbojets were/are single-turbine engines. Some, particularly in the early days or in low power applications, were/are even single-stage single-turbine engines. Recognize also that turbine systems are a necessary evil, not an inherent good. If we could get compression some other way, a turbojet wouldn’t have a turbine system at all. We call those ramjets. Our problems with those are mostly down to inadequate materials science. From a pure aero and mechanical simplicity perspective they beat the crap out of turbine-equipped engines.
The advent of turbofan engines changed the equations a bunch. Now the goal is to use the fire to drive the fan. Any thrust out the tailpipe is gravy, and inefficient gravy at that. The ideal turbine *system * in a turbofan engine will extract 100% of the thermal & pressure energy from the combustor exhaust stream. Such that the exhaust downstream from the turbine *system *is ambient air at ambient pressure just wafting gently out the back.
It is not efficient to try to do that in one stage. The input and output conditions are more radically different than can be converted by passage over just a single airfoil. And the hotter & more highly compressed the combustion scenario, the more that it true. Modern engines with modern compressors & combustors burn far hotter and at far higher pressures than the canonical science museum 1950s turbojet. And hence require more sophisticated turbine *systems *to extract the vastly greater work available per cubic inch of combustion exhaust gas generated.
From a steady-state and strictly aerodynamic POV it might well be possible or even advantageous to use a single *turbine *with multiple stages. But aero efficiency within the turbine is not the only consideration and steady-state is not the only operating point.
Prior to the advent of the geared turbofan, the fan RPM was exactly equal to the LPT RPM. And even with the latest in geared fans, the power extracted by the fan drive turbine (typically called LPT) is equal to the power applied to the fan, net of frictional losses. So the LPT has to be sized to deliver the right amount of power to the fan and at reasonable RPMs. And has to be able to accelerate and decelerate at reasonable rates, without producing excessive back pressure against the upstream turbine(s). All this leads in the direction of a large and relatively slow-turning fan-drive turbine.
Further upstream in the turbine system, the demands are different. The compressor system(s) turned by the other turbine(s) have different power, RPM, and acceleration needs. And as noted above ref ramjets, 100% of the work extracted by these turbines is “wasted” in that it’s used to run the engine, not to produce anything useful outside the engine. So the more efficiently we do this, the greater the overall combined efficiency of the overall engine. There’s really a twofer to be gained by doing this part right. As a separate matter, one or more of these turbines drive the mechanical accessory loads: fuel pump(s), hydraulic pump(s), electric generator(s), etc. Which each have their own mechanical power consumption needs and range of acceptable RPMs and accelerations. And which needs are mostly independent of the engine’s needs.
It turns out that for the current state of the art, using 2 or 3 distinct turbines is the best way to extract as much energy as possible as efficiently as possible. Efficiency also includes things like packaging diameter and weight of parts.
I suspect the ultimate in efficiency would be to have each turbine *stage *be separately freewheeling and have an embedded superconducting electrical generator in the hub. Each stage would be like this and computers would adjust the field currents to extract the appropriate increment of power from each stage according to its aerodynamics of the moment. Meantime the compressors and fans would be driven by similarly situated and controlled electric motors which consume all the energy extracted from the total turbine system, less some for ship’s services.
In effect today, the current HPT/LPT systems are a bit like a transmission, coupling a high speed high energy low volume combustion chamber to a low speed high energy high volume fan. There is an upper limit to how much gear reduction you can get in a single gear pair. Beyond that point it’s more efficient, albeit more mechanically complex, to use a two stage gear train.
Bottom line, it isn’t really sensible to talk about one stage of an existing two-stage engine without the other stage; the engine *as designed *can’t operate without both. e.g. In the current Toyota 2.0 liter 4 cyl, which is more efficient, the piston or the cylinder? Not really a logical question. The metaphor isn’t perfect, but it’s in the ballpark.
I wondered about putting the turbine further behind because I figured that this would allow the exhaust to expand, thereby increasing its volume while reducing its pressure and heat, which would allow either higher exhaust heat or turbine stages made of less expensive materials.
While ramjets are great for anti-air missiles, I was under the impression that for other purposes, they are lackluster. Their fuel efficiency below about Mach 2 is low because their speed doesn’t afford enough air compression. At and above Mach 2, they’re at their optimum fuel efficiency but still less fuel efficient than turbojets/fans/props for a given distance. This makes them quite suitable for anti-air missiles which need to go quite fast and greatly benefit from a light, cheaper engine. Is this accurate?
Turbofans more a larger amount of air slower than turbojets and are more fuel efficient for a given distance because of that, right? Do I understand correctly that this is because thrust is proportional to the momentum of the working mass whereas the fuel requirement goes up at the square of the working mass’ velocity?
Compressor:
This stage takes up a good chunk the exhaust’s energy, especially when idling, right? I haven’t been able to find any figures for what % of exhaust energy is used to drive the compressor and how much air it supplies at different power settings/speeds for a given engine.
While I know that compressing the air allows much more energy to be extracted out of the fuel, I am not clear on why. Is it because without a compressor, the combustion would either be oxygen-starved (low fuel efficiency) or have to use a tiny amount of fuel (low engine power-to-weight ratio)?
So, if the turbine is put some distance in the back and the nozzle widened, that wouldn’t give the exhaust time to expand and thereby cool? Or would the cooling be insignificant for any reasonable widening and distance between the combustion chamber and the turbine?
The entire point of the jet engine is to produce high-temperature/high-pressure stream of gas. This allows the same turbine wheel to produce more power. And since turbines are heavy, you want as few stages as possible.
Expanding the combustion products before they hit the HPT would result in larger and heavier turbine, and a larger diameter engine.
The whole engine is a system. There’s no reason to make the fire hotter than the turbine can withstand. If you’re constrained by turbine material or cooling, simply design the upstream engine sections make a less-hot fire. Adding additional engine length which simply takes the heat you just spent big $ creating and wastes it by trying to radiate or conduct it out into the surrounding structure is not smart or efficient.
Also, simple distance without an increase in cross-section wouldn’t do much for volume/expansion.
Pretty much. My comment about ramjets was mostly a red herring to emphasize that the turbine component of a jet engine, like the camshaft in IC engines, are a deadweight loss, something that ideally we could / would eliminate in the name of efficiency. For the mission of powering subsonic air vehicles below about 50,000 feet, the systemic efficiency of turbofans beats ramjets 8 ways to Sunday. For the much narrower question of aerodynamic efficiency within the flow stream of a jet, ramjets have massive benefits. Sorry to red herring.
That’s one factor. Another thing goes all the way back to Carnot. We’d like to extract 100% of the energy from the exhaust stream somehow. When it exits a traditional turbojet at high velocity & temperature, that couples poorly to thrust. A lot of the energy still present in the exhaust stream is wasted simply heating the surrounding air & injecting turbulence into it, which waste we perceive as the ear-splitting noise of a 707 or F-100.
Compare it to the physics of a rocket nozzle, where the ideal nozzle expands the exhaust stream all the way to ambient pressure. All real-world nozzles are huge compromises for size/strength reasons, packaging reasons, and the problem that a fixed nozzle has to be used across a wide altitude range. Part of the benefit of multiple rocket stages is the ability to swap out a sea-level-optimized nozzle for an exo-atmosheric-optimized nozzle.
No clue on the former. A real powerplant engineer could give us some rules of thumb, but that guy ain’t me.
As to the latter, it’s mostly about how big a fire you can create. You have to burn pretty close to stoichiometric. So using ballpark numbers, if you want to extract the energy in 3 lbs of fuel, you need 100 lbs of air. The more air we can run through our machine per second, the more fuel we can run through it per second. More is better when we’re all about power-to-weight in our overall engine and we need a pretty high power number.
Once we know how much energy we need and therefore how much fuel we need to burn per second we can choose between a great big campfire, or a very small concentrated fiercely compressed fire-in-a-can. Lots of losses scale from surface area, so a small concentrated fire gains there too.
Finally, let’s remember our old friend Carnot. Ultimately, efficiency is constrained by the available delta T. One side of the delta is stuck at ambient, so the only way to increase delta T is to maximize T[sub]max[/sub], the temp of our fire. This is the same reason a 12-to-1 compression IC engine has more HP &/or torque per cubic displacement than does a 6-to-1 compression engine. It’s also why super- or turbo-chargers work on IC engines.
Not really. Remember that the flow inside the jet is subsonic and constrained at the end. So the pressure is going to be about the same everywhere. If you expand the exhaust, you’ll just slow it down to keep the flow rate the same. It might heat up a tad since the KE will turn to heat.
You need a nozzle if you really want to expand your flow. You constrict such that the flow becomes sonic, and then expand. Downstream gas can’t “communicate” upstream since it’s sonic, and so instead of slowing down to equalize pressure, it speeds up as it expands and collides with the nozzle wall.
You basically have a rocket at this point, and can efficiently use the gas itself for propulsion. Real rockets power their pumps with a secondary fuel stream instead of a downstream turbine. I’m not entirely sure what you’d get if you tried to put a turbine at the end of a rocket (other than slag, that is).
Points taken on my idea of putting a turbine further back. It’s a mistaken idea.
No reason to be, it was instructive.
Turbofans are much more often used than turboprops yet turboprops, while slower, are more fuel efficient for a given distance. Fuel efficiency is very important to airlines. Passengers are very price sensitive and freight shippers may well choose to ship their goods at Mach 0.5 rather than Mach 0.8 if it means lower fares. So, is noise the only reason for the predominance of turbofans in the civilian market?
In the civilian market, you usually get to fully react your fuel with the oxidizer. What happens when you’re running the engine below or above the speed and altitude it was optimized for?
A ramjet running at Mach 1 has lower specific impulse than when it’s running at Mach 2. Is that because at Mach 1 you’re getting less than 100lbs of air but still pumping 3lbs of fuel, thereby making the combustion far less than stoichiometric?
A ramjet running at Mach 4 will also have a lower specific impulse than one running at Mach 2. Is that because while you might be getting 200lbs of air, getting to Mach 4 requires pumping more than 6lbs of fuel, thereby making the combustion far less than stoichiometric?
Is this why and are there other reasons?
The most important reason is the cruising altitude. If you want to run reliable long-distance service, you need to fly above 30,000 feet. At this altitude, the most efficient combination is a turbofan flying at M0.8-M0.9. Turboprops are more efficient at lower speeds and altitudes and are only used on short routes (where the time and fuel lost climbing to 30,000 feet does not pay off).
Modern jets have very efficient combustion. You may get some unburnt fuel during startup, but not elsewhere. The control system constantly adjusts the fuel flow to prevent this (and keep the engine within its operating limits).
A rule of thumb for a gas turbine engine at full power:
[ul]
[li]Compressor adds 30W of work into the air stream.[/li][li]Combustion chamber adds 70W of heat.[/li][li]HPT removes 30W to drive the compressor (30 out of 100 = 30% efficiency with 1 stage)[/li][li]power turbine (in ground engines) or LPT (in turbofans and turboprops) removes another 30W (30 out of 70 = 40% efficiency with 2-3 stages)[/li][/ul]
At the simplest conceptual level, because without a compressor what you have is a ramjet, and ramjets produce no thrust when stationary or at low speeds. The compressor is the answer to the question, “when fuel is ignited in the combustion chamber, why does the exhaust not come out both ends?” In fact in the event of a compressor stall, it might.
Any other reason besides flying above the weather?
They do get different fuel efficiency based on speed, don’t they?
Like here we can see specific impulse go up for ramjets until Mach 3 and then it goes down:
https://upload.wikimedia.org/wikipedia/en/e/ef/Specific-impulse-kk-20050824.png
or this at page 2: http://citeseerx.ist.psu.edu/viewdoc/download?rep=rep1&type=pdf&doi=10.1.1.215.6041 (ACHTUNG PDF!) where the specific impulse curves go and then down. Why does that happen if all the fuel gets burnt?
Passengers are price sensitive. But that’s not everything. Passengers are also time sensitive. And if we flew at 1/2 the speed our crew costs and capital costs would double. After all, people don’t buy a 1 hour ride, they buy a 400 mile ride. So for us to move X hundred people Y hundred miles we’d need twice as much time, or twice as many airplanes, or some combo in between to get the job done.
Compared to turbofans, turboprops have higher vibration and interior noise, and are perceived by the public as low-tech and dangerous. One of the big future directions in turbine propulsion is the prop-fan AKA “unducted fan” (a term I despise). One of the big question marks over their future is whether they’ll pass the marketing test. Will the public shy away? Nobody knows.
The engines’ control mechanisms always adjust the quantity of fuel to match the quantity of air so as to burn at stoichiometric, net of e.g. perhaps burning a bit lean for internal engine cooling, or a bit rich to avoid creating excessive NoX. Being off altitude isn’t a factor in this phase of the optimization.
The engine has a pretty flat efficiency curve near the high end. So we climb at the maximum possible thrust considering longevity and thermodynamic limits. As we climb and the air gets thinner the thrust drops off: mass (not volume) of air input per second is proportional to thrust out per second. And ideally we climb until the max thrust is just enough to maintain altitude at a speed that is ideally efficient for the air vehicle. So now we have the airplane and the engine *both *operating at their design optimal point. All is rainbows and unicorns.
In the real world we have constraints that make for fewer happy prancing unicorns. At heavy weights we need to fly faster than the aerodynamic ideal. At light weights we run into the altitude ceiling before the thrust decays low enough that we can’t climb any more. Sometimes the wind is better lower, so even though our true airspeed (TAS) is better higher the increasingly adverse wind gives a faster ground speed (GS) lower. GS is what affects total time and hence total burn to get there. Sometimes it’s turbulent at the optimal altitude and smooth higher or lower. On busy routes, there’s simply not enough airspace to accommodate everybody at their most optimal altitude. If we’re late, or if this flight is planned fast so our schedule is advertised as being 2 minutes quicker than Brand X, we’ll operate at speeds above optimal. Etc., etc., etc.
The whole thing is operated with an eye to fuel efficiency. But that means the manufacturers try to build machines with a fairly broad plateau of 90%-of-possible efficiency surrounding the tiny pimple of 100%-of-possible efficiency. Airline HQ’s job is to trade off all the other factors du jour to find the best spot on the plateau for todays conditions and goals. And my job is to operate at that chosen spot, and to apply sound tactics along the way to tweak towards greater goal-attainment when evidence shows reality isn’t matching the planning factors used. Sometimes goal-attainment means “save fuel & screw the schedule”; other times it means “save schedule and screw the fuel”.
The answer to your question is quite complex, so I am simplifying here quite a lot. To recap, specific impulse (Isp) is the ratio of thrust and fuel consumption. For an air-breathing engine, the thrust is the difference between the engine inlet and outlet speeds. Engine inlet speed is equal to the aircraft air-speed, and engine outlet speed is usually called the exhaust velocity.
At supersonic speeds, the kinetic energy of air is significant. This energy is used to compress the air at inlet. While a turbojet relies primarily on a mechanical compressor, a ramjet has to rely only on ram-air for compression. So as airspeed increases, the compression ratio increases, which in turn increases the exhaust velocity. The relationship is non-linear, but below Mach 3, it helps to increase Isp with increasing air-speed.
However, compression also heats up air. At higher air-speeds, the compressed air gets quite hot, even before you add fuel. Therefore, you can add less fuel (less energy) to avoid melting the engine, and thrust decreases accordingly. Above Mach 3, this effect dominates, decreasing Isp as air-speed increases. At some very high speed, the compressed air would be so hot, you would not be able to add any fuel, and would produce zero thrust.
Which loops back to my point in post #5 that the main challenge to ramjets (within the speed regime where they work at all) is materials science. *If *we could build engines from lightweight unobtainium and have 5000K combustion chamber inlet temps, plus add lots of fuel and combust it thoroughly *then *we’d have some truly insane exhaust energy levels. And pretty impressive Isp as well.
What kind of Isp do you think could be attained? Anything in the 3000 neighborhood?
Have I got it right that the turbine tends to be one of the pricier parts of a jet engine?
Also, opinions on the roles of air-augmented rockets (Air-augmented rocket - Wikipedia) and turborockets? ( Air turborocket - Wikipedia )
The former is pretty much only suited to anti-air missiles but turborockets look like they could have much broader applications.
All these things are pretty speculative. Some have existed as sorta-functional models. The details of them are beyond my ability to comment on.
This SABRE (rocket engine) - Wikipedia is the latest industry darling. Which may in fact have some real life in it. We’ll know for sure in about 5 years.