Pratt and Whitney announced, in early June 2013, the successful first flight of the PW1100G Engine: The very engine family (PW1124G,PW1127G, PW1133G) that will be one of two engine options to power the Airbus A320NEO aircraft family (A319NEO, A320NEO, A321NEO). The other engine option is offered by GE-SNECMA’s CFM International’s LEAP engine series.
While CFM’s LEAP engine family is an improvisation of conventional turbofan engines, the Pratt and Whitney PW1000G family, of which the PW1100G series is a member, uses a rarely adopted technique, to promise high fuel burn saving by adopting a bypass ratio hitherto unheard of: 12:1. In this article, we explore, at a high level, the design of the PW1100G family, how it compares with existing A320 engines, the differences, on a high level, of this family with the competitor’s offering: the LEAP 1A, and why the Boeing 737MAX family doesn’t need such a large turbofan engine.
Getting Technical
The fuel efficiency of an aircraft is dependent upon three factors: the drag contributed by the airframe, as parasitic drag from the fuselage, wings, horizontal and vertical tail planes, and induced drag from aerodynamic effects such as wing tip vortices; the drag contributed by the engine, by virtue of its shape and size, and effects due to the exhaust gases; and third and most importantly, the efficiency of the engine itself.
From a high level, there are two factors that determine the efficiency of an engine: The thermal efficiency, and the propulsive efficiency. In this article, we disregard mechanical losses in transmissions, gearboxes, and all forms of inter-mechanical conversions. Thermal efficiency deals with how efficiently the engine extracts mechanical work from a unit mass of fuel that is burnt. For all forms of jet engines, including turbojets, turbofans, turboprops, and turboshafts, this is the energy conversion that takes place inside the core of the engine, which includes the compressor, the combustion chamber, and the turbines that extract mechanical energy from the hot, expanding gases.
Propulsive efficiency deals with how effectively the extracted mechanical work is used to generate thrust. In the case of turbojet engines, the mechanical energy generating thrust is the hot expanding gases that exit from the engine at a high speed. For a given size, a “pure” jet engine such as a turbojet can deliver significantly more thrust than other forms of subsonic air breathing propulsions. This is achieved by accelerating a small mass of air to high speeds.
In all other derivatives of a turbojet, such as a turbofan, turboprop, and turboshaft, the propulsive efficiency is determined by the mechanism that extracts energy from the turbines in the engine, and how that is used to move a mass of air. In a turbofan engine, the energy of the hot expanding gases is used to drive a set of turbines, which turn a large ducted fan at the front of the engine. The fan moves a large mass of air, at a good speed. In a turboprop engine, the turbines turn a shaft, which moves to a reduction gearbox that turns slower but with greater torque. The high torque and low speed drives a large propeller, which moves a larger mass of air at a slow speed. In a turboshaft engine, the energy in the shaft is, much like a turboprop engine, passed through a reduction gearbox that drives a large rotor, such as that in a helicopter, which moves a very large mass of air at a slower speed.
In all three derivatives of the “pure” jet engine, the actual jet engine is relegated to the compressor, combustion chamber, and the turbines that drive the compressor. The energy extracted from the hot gases, via an extra set of turbines, drive either the fan, the propeller, or the rotor attached to a shaft, leaving very little energy in the hot gases that leave the turbines. The hot exhaust contributes to little (turbofan), or no thrust (turboshaft, turboprop).
This allows the creation of an engine that marries the best of both worlds: the advantages of a jet engine (such as high reliability, low failure rates, the ability to operate at high altitudes, and the high energy conversion efficiencies), with the benefits of methods that deliver high propulsive efficiency. Watch this video to understand how a jet engine works:
Propulsive Efficiency.
A heavy truck travelling at the same speed as a light car will inflict a greater damage in a collision. A car travelling at a higher speed than an equally lightweight but slower car will inflict a greater damage in a collision. Mass and speed determine momentum: larger of either, or both, will result in a greater momentum. Larger the momentum, more are the forces associated with a collision. Theoretically, the lightweight car mentioned in this example, if travelling fast enough, can inflict as much damage as a slow moving heavy truck.
If the same slow moving heavy truck is brought to a gradual halt, the damage caused is lower. If on the other hand, the truck is brought to a sudden halt, such as hitting a concrete wall, the damage can be far severe. The rate of change of momentum determines the forces associate with, as in this example, a collision.
The same principles apply to thrust generation, though not for destructive purposes. The rate of change of momentum determines thrust. The mass here is the mass of air that the engine ingests and spews out. The velocity is the speed of the exhaust gases, or air. These two determine the momentum. The rate of change of momentum is defined, in the case of an aircraft engine, the time taken to impart the exhaust velocity to the exhaust gases/air. This is the time taken between the engine ingesting a given mass of air, and expelling the air at a higher speed. The mass of air over time is known as mass flow rate.
Additionally, as the speed of the aircraft through the air approaches the speed at which the gases or air are expelled from the engines, the propulsive efficiency starts approaching 100%. If an aircraft engine’s exhaust airstream has a speed of 300kts, then the engine will perform best when the aircraft to which it is attached flies at close to 300kts. Above and below this speed, the propulsive efficiency decreases.
Supposing that an aircraft designed to fly at not more than 300kts through the air, needs a certain amount of thrust, say, 30,000 lbs (136,000N). The thrust can be delivered either by an engine that exhausts gases at 600kt (308 m/s), and has a mass flow rate of 440 kg/s, or by another engine that exhausts gases at 300kt (154m/s), and has a mass flow rate of 881kg /s. The second engine will bode well for the 300kt airplane, but will require a larger mass flow rate.
Ingesting a larger mass of air per second, and keeping the exhaust velocity low, will need a larger larger fan, or a larger propeller.
A large fan has three problems: First, it offers more drag to the oncoming air, which can offset the gains in propulsive efficiency. Second, a larger fan makes the engine heavier, adversely affecting fuel burn. Third the larger the fan gets, the faster the tips of the blade travel through the air, for a given rotational speed. Here is an example.
The fan of an IAE V2500 engine (that power the Airbus A320) spins at a maximum speed of 5,650 RPM. The fan diameter is 63.5 inches (1613mm). The distance covered by the blade tip, in one revolution is 1.613m X π = 5 meters. 5 meters X 5650 RPM = 28,630 m/minute = 477m/s, which is about 1.4 times the speed of sound at sea level. If the PW1100G’s 81 inch diameter fan (2057mm) is spun at the same speed, the blade tips travel at 608m/s, which is almost twice the speed of sound! (This is responsible for the characteristic chainsaw noise that can be heard from an Airbus A320’s engine when taking off at close to full take off power). A large amount of energy will be needed to overcome the drag associated with the high speed of the blades (this is different from the drag that the blades pose to the oncoming air stream). In addition, the airflow becomes more complex.
The Bombardier Q400, for instance, is a slow moving turboprop airplane (in relation to a jetliner) that has a 6 blade propeller, spinning at a 100% speed of 1020 RPM. The diameter of the massive propeller is 13.5ft (4115 mm). The tip speed at 1020RPM is 220 m/s.
In short, a turbofan engine suited for a subsonic airplane must have a larger diameter fan, to produce the same thrust at low exhaust velocities but higher propulsive efficiencies. However, the fan must spin slower, to keep the tips from attaining very high velocities.
Conflicting requirements
The core of the engine, which is the “true” jet engine relegated to the role of a gas generator, is made of many significantly smaller diameter bladed disks. These bladed disks, some of which form the low pressure compressors, and others the high pressure compressors, ingest only a part of the total air sucked in by the large diameter fan. That air is compressed to a significant amount before mixing with fuel and being ignited in the combustion chamber.
For example, on the IAE V2527-A5, 17.24% of the air ingested by the fan enters the compressor. 4, smaller diameter disks with blades serve to initially compress the air ingested by the compressor. These compressors disks run at the same RPM as the fan at the front of the engine, which is a maximum of 5650 RPM.
The air is further compressed by 10 disks spinning at a speed different from those of the fan and low pressure compressor stages. These disks spin at a much higher speed, at a maximum of 14,950 RPM. It is this high rotational speed, that allows the 10 high pressure compressor stages to effectively compress the air to required levels. At the end of the tenth high pressure stage, the air is compressed to 32.8 times that of the ambient air. The small diameter, high RPM, high compressor stages contribute to most of the compression.
Because of the low contribution of the four low pressure compressors (attached to the fan) to the overall compression, the high pressure compressor must incorporate 10 stages. If however, the low pressure compressor could deliver a higher compression, the high pressure section could be reduced in stages. For the low pressure compressor to perform better, the disks must spin at a higher speed. But since the fan is attached to the same shaft that spins the low pressure compressors, and as demonstrated earlier the need for a slow spinning fan, the rotational speed of the low pressure compressors are limited.
Reducing the number of compression stages in an engine reduces weight, decreases system complexity, reduces the overall length of the engine, saves cost, and improves efficiency.
“Meshing” two speeds
The industry has long been aware of the need for slow spinning fans and fast compressors. By employing the same mechanism that is employed in a turboprop engine, which is a reduction gearbox, the fan essentially gets attached to a third shaft: one that is mechanically linked to the low pressure compressors, but spins at a lower speed. With the addition of a gearbox, the turbofan engine becomes a geared turbofan engine: a multi bladed, shrouded turboprop, in essence. The low pressure compressors run at a high speed, and the gearbox spins the fan at a much lower speed, allowing both sections to run at their optimum speeds.
The Geared turbofan isn’t a new concept. The British Aerospace BAe 146, a regional airliner that first flew in 1981, and produced till 2002, was fitted with four Textron Lycoming ALF 502R-5 geared turbofan engines. The Bombardier Challenger 600 originally were fitted with the ALF 502L geared turbofans. The TFE731, a geared turbofan engine, first ran in 1970, and its variants power popular airplanes such as the Learjet 35, 40,45 and 55, Dassault falcon 900DX, Hawker 800,850XP and 900XP, and a few Cessna Citations.
Pratt and Whitney PW1000 Geared Turbofan Engines
Pratt and Whitney’s PW1000G series of geared turbofan engines are a result of over 15 years of development that started with the work on the PW8000, which was planned as a V2500 and a CFM56 replacement for narrow body airliners such as the Airbus A320 and the Boeing 737. The PW8000 was announced in 1998, almost 11 years before Airbus announced the A320 NEO. A 30,000lb thrust demo geared turbofan engine flew during the conceptual phase for the first time, in 2008, on Pratt and Whitney’s Boeing 747SP Flying Test Bed, and later on an Airbus A340 Flying Test Bed.
The PW1000G “Geared Turbofan Engine” Family will power the Mitsubishi Regional Jet, the Bombardier C Series, the A320NEO family, and the Russian Irkut’s MC-21 series. The family spans across a 15,000lbs to 33,000lbs thrust range, but the architecture remains unchanged throughout.
The PW1000G architecture comprises one fan upstream, followed by a reduction gearbox that allows the low pressure compressors to run faster than the fan, two low pressure compressor stages, 8 high pressure compressor stages, 2 high pressure turbines (to keep the high pressure compressors running) and 3 low pressure turbines that keep the low pressure compressors and the fan running. A listing of the number of bladed disks is referred to as the “stage count”, and in the case of the PW1000G family of engines is:
1-G-3-8-2-3, with the “G” referring to the gear system that links the fan(1 bladed disk) with the low pressure compressor (3 bladed disks).
The reduction gearbox is planetary gear system, with the low pressure shaft driving a sun gear, and five planetary gears enmeshed between the sun gear and a ring gear non-rotating relative to the engine nacelle. A carrier cage, holding the planetary gears, drives the fan.
The PW1000G series have different variant numbers, even for seemingly similar engines. For example, the engines on offer for the A320 and the MC-21 are very similar, the difference lying in the numbering. Pratt and Whitney number their engines as PW-[Generation]-[Customer]-[Thrust Class in thousands of pounds of thrust]-[Specific]. This makes a 27,000lb (12.2kN) engine for the Airbus A320 NEO as PW-1-1-27-G, and the same thrust engine for the MC-21 as PW-1-4-27-G.
The PW1100G series, which will fly on the A320NEO family, feature an 81 inch (2057mm) diameter fan, and an astounding bypass ratio of 12:1, the highest ever. The high pressure compressor is expected to spin at around a maximum of 20,000RPM, and the low pressure compressor is expected to spin at around 10,000RPM. The gear ratio of the gearbox is 3:1, implying the low pressure compressor bladed disks spin at three times the speed of the fan, which is expected at around 3500RPM maximum. This results in optimal performance of both the fan and the low pressure compressor. We’ll see how this engine compares with the existing, highest thrust power plants for the A320: the IAEV2527-A5, and the CFM56-5B4.
At such rotational speeds, the tip of the 81 inch fan is expected to touch close to 370m/s, which is just 1.08 times, or 8% over, the speed of sound. Watch this video to understand more about the PW1100G:
Further Innovation
Pratt and Whitney had planned to implement Variable Area Fan Nozzle (VAFN) to ensure optimal efficiencies across the flight spectrum. The VAFN, also known as the fan variable area nozzle (FVAN), attempted to adjust a flap assembly (seen colored in red in the image on the left) that would vary the fan exit area through which the fan air is discharged. This would vary the speed of the exhaust gases, optimising thrust and fuel economy at each flight regime. Pratt and Whitney patented a system, which is the first of its kind for high bypass turbofan engines: a simpler, and inexpensive system of variable area nozzles, as compared to those seen in military jets.
Pratt and Whitney however dropped the idea, “after the fan blade demonstrated better performance across the flight spectrum.” This indicates a method to further inch toward better propulsive efficiencies, but dropped in the light of the gains expected versus the loss in engine reliability and overall increase in weight associated with the addition of another mechanical system.
The alternate A320NEO engine: the CFM LEAP-1A
CFM hasn’t opted for a geared turbofan approach. Instead, CFM has stuck to conventional engine design: a twin spool engine, with the low pressure rotor connecting the low pressure compressor stages and the fan, making them spin at the same speed, and high pressure compressor stages that run at a different, higher speed.
CFM has endeavoured to target two birds with one stone: increased propulsive efficiency, and increased thermal efficiency. GE has long been working on raising the temperatures that can be withstood by its turbines. By ensuring high compression ratios, high temperatures, and a twin annular pre-swirl combustor contributing to a “lean burn”, the thermal efficiency of the engine is increased. However, the blades of the high pressure turbines, upon which the extremely hot gases impinge, must be able to withstand a higher temperature. CFM uses “advanced cooling” to keep the blades of the turbine cool, while employing a ceramic matrix composite (CMC) shroud to withstand the temperatures. Watch this video to understand more about the LEAP:
CFM’s fan for the Airbus NEO, measured at 78 inches in diameter is bigger than the 68.3 inches on the CFM 56-5B4 on existing A320s. The nearly 10 inch increase in fan diameter translates to a higher propulsive efficiency, but since the fan is attached to the low pressure compressors, it comes at a price. While the sum of the low pressure and high pressure compressor stages remain the same for the LEAP 1A and the CFM 56-5B4, at 13, the number of turbines differ significantly. While the -5B4 has 1 high pressure turbine + 4 low pressure turbines, the LEAP 1A features 2 HP turbines and 7 LP turbines.
Offseting the weight increase contributed by the significantly larger number of bladed disks are the “3D woven carbon fiber composite blades and case”, which are lighter, and promise greater durability.
CFM has been very cautious in the engine design, committing to “Proven Performance”, “Low Risk Execution”, and “Leading Technology”. The advantage that the LEAP will have over the Pure Power is the proven reliability of the conventional engine design, which a geared turbofan engine of such dimensions hasn’t had the chance to demonstrate.
Boeing 737MAX’s LEAP-1B vs A320NEO’s LEAP-1A: Engine Size and Thrust Requirement
The competition to the A319, A320, and A321 are the Boeing 737-700, 737-800 and 737-900, respectively. However, the thrust requirement for the Boeing 737-900, is 24,200-27,300lbs, which is much lower than the 30,000-33,000lbs thrust range required for the Airbus A321.
While the 737-800 can operate with a thrust range of 24,200-27,300 lbs, the A320CEO operates with a thrust range between 25,000lbs – 27,000lbs, in which case the ranges are similar. The Boeing 737-700 can fly with engines featuring a thrust in the range of 20,600 – 26,300 lbs, and the A319 can fly with 22,000 – 23,500lbs. In all cases, a Boeing 737NG can fly with lower thrust, with the largest gap seen between the Boeing 737-900 and the A321.
The design of a family of engines is determined by the highest thrust requirement. In the case of the PW1100G family, the size of the fan, 81 inches, was determined based on the need for 33,000lbs of thrust from the PW1133G to power the A321NEO. The PW1127G and the PW1124G share the same design and dimensions, but operate at less extreme conditions.
In comparison, the Boeing 737-9 MAX needs a thrust range of 27,000-28,000lbs. Since the 737 family and A320 family have similar cruise speeds, for the same speed of exhaust gases from the respective aircraft engine’s fan, the A321’s engine will need to be larger to move a larger mass of air, as opposed to the Boeing 737-9 MAX’s.
It is for this reason that the LEAP-1A that will power the A320NEO family, has a fan diameter of 78 inches, while the LEAP-1B that will power the Boeing737MAX family has a fan diameter of 69 inches, prompting Boeing to state, “The A320neo pays the economic price for the A321neo’s thrust needs”, with a larger, heavier engine with more turbine stages that offer more dag for the same thrust.
Every fan has two sides: perceived pros and cons.
The geared turbofan engine has, theoretically atleast, a reduced reliability in comparison to a standard twin spool turbofan design, because of the inclusion of an extra mechanical stage: a gear system. The engine now houses three shafts, all turning at different speeds. Further, the gear system adds weight.
However, countering both effects seem to be the reduced number of engine stages, which have many benefits: decreased part count, increased reliability, and reduced weight. How the two effects: increased mechanical complexity and decreased stages counter each other’s effects is to be seen.
It seems unlikely, with the larger fan and the gear system, that the GTF will be lighter than existing IAE or CFM engines that power the A320 with the highest available thrust. Interestingly, the maximum thrust ratings for the PW1127G, as published by Pratt and Whitney, remain unchanged from what the IAE and CFMs offer for the A320: 27,000 lbs of thrust.
What changes, however, is the amount of drag that the new engine offers, due to its larger size. If there seems to be no significant weight difference between the existing engine options (2,500 kg) and the PW1100G series, there may be a performance penalty, marked by slower cruise speeds for the same thrust setting, possibly slower max cruise speeds, slower green dot (best L/D speed), slightly shallower climbs, and possibly degraded single-engine performance. The single engine performance difference is expected to be the most prominent, with a possibly larger yaw, and a definitely reduced climb gradient, consequently lowering obstacle clearances.
With a gear ratio of 3:1, the equivalent moment of inertia of the fan, as seen by the low pressure spool, is only 1/9th of the actual moment of inertia. The radius of the fan is 27% more than the IAEV2527-A5’s fan, making the volume of the fan approximately 2 times that of the IAE’s, implying very crudely that the mass is double that of the IAE’s, assuming the same material is used to make the blades. With double the mass and 27% greater radius, the moment of inertia of the PW1100G’s fan is about 3.3 times that of the IAE’s. This makes the equivalent moment of inertia of the fan, as seen by the low pressure spool, just 36% of the IAE’s, despite the larger mass and radius. Considering that the PW1100G’s low pressure spool has only 3 compressors and 3 turbines, as opposed to larger and heavier 4 compressors and 5 turbines on the IAEV2527-A5, the overall equivalent moment of inertia of the low pressure spool of the PW1100G’s is atleast around 25% that of the IAEV25257’s. With the PW1100G’s low pressure spool estimated to spin at twice the angular speed of the IAEV2527’s, and the overall moment of inertia around 25%, the spool up time may be reduced to around 50% of that in an IAEV2527-A5. This allows thrust to be made available faster, increasing safety margins through enhanced engine response times.
However, with the increased aerodynamic drag, the V1 will be lower, and the take off run is expected to be longer than A320s which feature high thrust IAE and CFM engines, with the Sharklets.
In all, while the PW1100G series engine is expected to contribute to around 11.5% of the proclaimed 15% fuel burn savings on the A320NEO, the A320 may not fare as well as the high thrust sharklet equipped A320CEO in high altitude, terrain challenged operations. However, neighbourhood noise is expected to be lower. The slow, large fan produces lesser noise, making the airplane quieter for both passengers and communities around airports.
Conclusion
The PW1000G series of engines, and especially the PW1100G family, are expected to be disruptive implementations, by employing the largest bypass ratio in the history of turbofan engines, and adopting a geared turbofan engine design of scales hitherto unmatched, promising double digit fuel burn savings. The slow speed of the fans, contribute to low noise, promising an enhanced passenger experienced, and reduced flight-related fatigue.
However, as with any new system, the reliability of such a huge geared turbofan engine isn’t known, casting initial doubts on dispatch reliability for airlines. Further, the PW1100G series focuses primarily on propulsive efficiency, forcing the engine to take on a large fan diameter of 81 inches, which will offer more drag than the competing LEAP 1A engine, which features a fan of diameter 3 inches smaller. This once sided effort towards better fuel savings increases drag, and may cost the Airbus A320’s takeoff, climb, and cruise performance, especially at areas that have short runways, and/or challenged by terrain. Although the spool up time of the GTF engine is expected to be lower, allowing the airplane to respond faster to a terrain alert, a penalty on climb performance is expected to exist, reducing, to an unknown extent, safety margins related to obstacle clearance.
The CFM LEAP 1A, on the other hand, with the reduced drag footprint, increased thermal efficiency, and optimised propulsive efficiency (although probably not as optimised as the PW1100G’s), may lead to similar fuel burn savings, with a lesser penalty on performance. However, the spool up time may be considerably longer than the PW1100G’s.
Either engine option will affect the 320’s performance, and may not be able to match upto the climb performance, safety and statistical reliability offered by today’s sharklet equipped A320 with either the IAEV2527-A5, or the CFM 56-5B4.
The Flying Engineer 
Excellent article..!
Great analysis TFE! Two questions though..
1. How is the LEAP-X able to achieve almost the same bypass ratio as the PW1100G despite not having the reduction gearbox and not having a larger (as compared to PW1100G) diameter fan?
2. Your last sentence in the conclusion suggests that the new engines are taking a step back at performance characteristics while trading off for better fuel efficiency. Would airlines not look at the safety and performance aspect of the overall performance of the engine rather than just look at how to save costs?
Not the same bypass ratio. 12:1 and 11:1 is a big difference. The Gears do not affect the bypass ratio, in any way.
The comparison was made with engines with the highest thrust ratings. Not many operators choose the best performing engines, as it adds cost and boosts performance to unwanted levels for normal operations. But when operating at extreme conditions which may need low drag and close to full thrust, you’ll see the difference.
A degraded performance does not necessarily make the airplane unsafe. It only requires operators to appreciate where they can deploy their airplanes and still remain maintain safety margins.
Thank you for the comment, Kaushik.
There’s a rather serious mistake in the Propulsive Efficiency section, in the example of the airplane that flies at 300 kts and needs 30,000 lbs of thrust. In flight, net thrust is given by the difference in flow momentum of the engine exhaust (gross thrust, fan flow plus core flow) minus the engine inlet flow momentum (ram drag). The numbers of the example would only apply in a static situation, with no forward velocity, and zero ram drag. Clearly not the case when flying at 300 knots.
Diagram matters first in aeronautics, because without diagram we cannot come to know about the working of the minor part of the engines especially any engine might be… Great article because very good explaination with the help of the diagram with graphs… This will surely help any guy who is not only an aeronautical but the one who is also very much interested to learn the concepts of aeronautics… Good job… I like it…
would anyone help me with the exhaust temperature in PW110g …..?
Great article……..What is meant by moment of inertia ?
Moment of inertia is a property of a body describing the distribution of mass about an axis of rotation. In practical terms think about the moment one must exert when slinging a relatively small mass around at a large radius compared to this same mass at a small radius. Another example would be figure skating: By drawing in their arms/legs a figure skater brings their mass closer to the axis of rotation thereby reducing their moment of inertia. If the skater’s angular kinetic energy is constant their angular speed (rotation) will increase as their moment of inertia decreases.
For linear motion (translation): Force = mass * linear acceleration
For rotation: Moment = mass moment of inertia * angular acceleration
However, note that the author refers to the equivalent moment of inertia. Since a gearbox defines a relationship between the rotational speed of the fan and that of the low pressure spool, if one desires to ‘add’ the moment of inertia of the fan to the moment of inertia of the low pressure spool the speed ratio of the gearbox must be considered to obtain an effective moment of inertia of the fan to add to the moment of inertia of the low pressure spool.
Does PW or its partner design and manufacture the planetary gear system?
I love this article. Does the rpm slow down in cruise compared to take off, or is there just a change in pitch of the fan blades, compressor blades and turbine blades?
(The fan exhaust must always be faster than the aircraft speed, so perhaps an exhaust speed of Mach 0.9 at cruise would give optimal fuel efficiency for a plane to go at Mach 0.85…?)
There are no variable pitch blades in the engine. Thrust is related to engine RPM< which is a different relation at different atmospheric conditions. Fan exhaust equaling cruize speed is a theoretical statement only. Thanks for the nice comment! 🙂
The PW geared engines will have a higher drag due to their increased size, but what will be the drag on approach with reduced power? Will the blades not run close to air speed and hence behave neutrally, or will the additional drag require more thrust? What about the abilty to give full thrust on aborted landing?
Thank you.
Wow, amazing article! I am a senior mechanical engineering student, so I understood most of this. Great analysis, I hadn’t realized the larger fan diameter greatly affected the performance of the engines. Also, the gearing system is genius. Love this.
I’m very interested in the future of aircraft efficiency and I found myself reading this from beginning to end, understanding almost all of it – thank you so much. My hobby project is looking at the practical realities of an electric hybrid drive with a turbofan or turboprop engine, but I can’t do it alone. If you can shape or contribute to this then please let m know. Many thanks again.
Sorry to disappoint you, Austin. If you are talking about electricity in commercial transport there is no room for batteries on board aircraft, in the foreseeable future. Flying on electricity faces two major headwinds. Both are fatal, pick the one you like to kill the idea: cost or weight. I will pick weight, let’s check on some facts, related to energy storage.
One kilogram of Jet fuel stores 43.2 megaJoules of energy, which are equal to 12 kilowatt-hour.
Now: one kilogram of onboard batteries (Lithium-ion type) can store, at best, 0.2 kwh.
12 divided by 0.2 is 60, like in six zero. For a given amount of stored energy onboard, you need 60 times more weight with batteries.
You may argue that electric engines are much more efficient than internal combustion, Brayton cycle turbofan engines. So let’s assume 90% overall efficiency for an electrically-driven turboprop engine (electrically-driven core at 100% efficiency and 90% prop efficiency) and also 33% overall efficiency for a current generation turboprop energy. Our 60 times number is now divided by 20. It would be around 30 for a turbofan.
But 20 – or 30 – times more weight of stored energy onboard is still absurd.
Cost-wise, batteries are also – and will remain – surreally expensive, as the slow sales of electric cars clearly demonstrate. What these numbers show is a dark scenario for the future of commercial aviation – it cannot easily get rid of fossil Jet fuel.
Thanks very much for your reply. I agree with your maths on energy density but I take a different and more optimistic view. The world is changing and battery technology is slowing getting better and cheaper whilst fossil fuel is not. Litium batteries have limited capacity, as you illustrate very clearly, but they do have a high power output for their weight.
Electric cars are a great analogy – 100% electric cars with real like-for-like value are only just emerging (Tesla’s Model S and the Formula E motor racing class) and they’re very expensive, but they go like the wind. By comparison, hybrid electric cars are popular- there’s barely a street here in the UK without a Prius – and they’re of equivalent performance to other cars. So, surely it is worth having electric components on board a vehicle, be it a car or an aircraft, when the weight and cost of the components is outweighed by the benefits of having them there. The challenge with aircraft is safety, space and of course weight, weight and weight. Drawbacks include having to land with the same weight as you took off with and the speedy turnaround time i.e. limited time to recharge a battery. For me, all of this points to a ‘thin’ hybrid solution until technology advances much further. Even a few% points reduction in fuel usage would interest many airlines more than the average driver.
Having read the article on the Q400 vs ATR72 I started thinking- how could we increase the power of the ATR72 without increasing fuel usage significantly? Various articles suggested that a turboprop loses a significant amount of power as altitiude increases. One factor is the use of bleed air for providing cabin air at pressure. P&W literature suggests a maximum of 8% of low pressure air can be bled. Imagine having an electrical air system in place of the bled air for the phases of the flight where maximum power is required i.e. take off and climb.
Where do you get the energy from? The hybird car gets much of its benefit form regenerative braking. This is not so easily done on a turbofan but perhaps a turboprop could be engineered to use the decent phase to partly recharge the batteries from the propeller to a small generator. Of course the flight plan would have to be adjusted so that jet fuel is not then used to cover the lost distance. Without ground power to complete the charge then benefits would be lower.
Does anyone have access to a FMC for the ATR72 ? Could you find the extra fuel required for, say, 500kgs additional payload (for batteries etc) on a 90 minute flight? On top of a normally loaded flight? Having such information would be really, really valuable.
Three good arguments here, worth treplicating: weight / bleed / descent.
Battery weight is not that much of a big deal in ground vehicles, provided there is regenerating braking in the system. If deceleration energy is not lost to surreally stupid friction brakes, no problem. Not so in airplanes, where there is a continuous battle against gravity. The cost of that fight is drag. For a constant aerodynamic efficiency (or lift / drag ratio), more weight => more lift => more drag => more thrust => more power => more energy consumption, of whatever energy type. Let’s see if someone helps us with ATR72-600 fuel burn data for the 500 kg exercise you proposed. It will dramatically show that electrons will not fly any time soon.
Bleed extraction is one area worth exploring. The B787 all electric ECS (Environment Control System) is a good example, utilizing electricty to drive ECS compressors gotta be more efficient than extracting bleed from the engines. Again, cost and weight are huge issues, and scaling the B787 technology down to turboprop size levels is also a question mark.
A final word on descent regeneration. Again, there’s a major difference here between cars decelerating and airplanes descending. A decelerating car has kinetic power to give away ‘for free’, either blown away in the form of heat in surreally stupid friction brakes, or collected and stored in on board batteries by regenerative braking systems. Not so in descending airplanes. The power that is given away during descent (rate of airplane potential energy being ‘lost’) is needed to overcome drag, like in gliding sailplanes. Engine power is reduced to essentially zero, and pitch angle is lowered to get a piece of the weight vector to produce thrust. If you add drag to produce some sort of electricity during descent, the pitch angle has to be lowered proportionally to compensate for that. Consequence? Descent profile is steeper, your top of descent point goes further closer to the destination, gotta stay longer in cruise, using cruise power for a longer time. The energy balance here is not one to one. Energy is lost by doing this, after all the efficiencies are factored in.
Let’s stay tuned on the 500 kg exercise and see what comes out of that, should be interesting.
Its a very much informative news about P&W engine and its contents of the function for specially the AME’s
Wouldn’t the relative size differences of the nacelles cause the increase in parasitic drag of the PW1100G? Neither the fan blades nor the core are drag-causing elements of the engine during phases of flight where the aircraft accelerates or maintains speed. In those phases, the fan blades and core are adding momentum to the airstream, and are not an obstacle to the passage of air. If true, then only the nacelle (and perhaps a tiny contribution from the spinner and the core/fan splitter) should be considered when examining the drag caused by the engine. Thank you, Joel
That is true for flight with both engines operating Joel. However, the limiting takeoff preformance and v speeds are calculated for the aircraft to safely climb to MSA in the event of single engine failure at or above V1. This is when the large fan will produce alot of drag, leading to lower climb gradients in the first and third segments. However, I am of the opinion that V1 will have to be increased to cater for increased Vmcg due to the greater yawing moment. Great article Flying Engineer.
One thing i want to know is that, according the text and the clip movie, i found that the lower speed of the fan and larger and larger is much much better , as there is ”Ingesting a larger mass of air per second, and keeping the exhaust velocity low, will need a larger larger fan, or a larger propeller”, what does it exactly mean? t means, for example PW1000G which spins at 3500 rpm, if its was 3500mm and rotates at 1800 was better? if so, the low speed can bear the pressure in rear at exhaust as pressure increase?
second, why keeping the exhaust velocity low is good? don’t you think giving more speed to the exhaust velocity is not better as we know the thrust is the difference speed between incoming and exiting air’s velocity?
Relative to Hido’s comment (09/October):
Let’s define Wa as mass flow of air through the engine, Vo as the flight velocity (true airspeed), and Vj as the velocity of exhaust gases.
Assuming a simple turbojet example, net thrust (FN), as Hido rightly noted, is mass flow times the difference beween incoming (flight velocity) and exhaust flows:
FN = Wa * (Vj – Vo)
Now: propulsive efficiency (let’s call it Eta P) defines how good the engine is when transforming the difference of incoming and exhaust kinetic power into the raison d’etre of aeronautical engines, thrust horsepower (let’s call it THP, which is net thrust times flight velocity, FN * Vo).
Eta P = THP / (Kinetic Power Out – Kinetic Power In)
Eta P = FN * Vo / { ( Wa * Vj * Vj / 2 ) – ( Wa * Vo * Vo / 2 ) }
Eta P = Wa * (Vj – Vo) * Vo / { Wa / 2 * ( Vj * Vj – Vo * Vo) } where (Vj * Vj – Vo * Vo ) = (Vj + Vo) * (Vj – Vo)
We’re almost there…
Eta P = 2 * (Vj – Vo) * Vo / { ( Vj + Vo ) * (Vj – Vo) }
Eta P = 2 * Vo / ( Vj + Vo )
In the equation for Eta P above, it can be seen that propulsive efficiency is ideal (1 or 100%) when Vj = Vo, id est, when the velocity of exhaust gases is equal to the flight speed. But what happens to net thrust when Vj = Vo?
FN = zero, no net thrust.
Propulsive efficiency gets better the closer Vj (velocity of exhaust gases) gets to Vo (flight speed).
But that also makes net thrust get closer to zero.
What is the best compromise here, not considering other aspects such as fan size, nacelle drag, engine weight, mechanical complexity, etc?
Produce net thrust with the highest possible Wa (large mass flows, large fans, high by-pass ratios) and the smallest possible Vj – Vo (average delta velocity across the engine).
Good Article followed by some good discussion. I interested to know how much the increased fan size will affect aircraft take off at altitude say (5500ft) and if runway length will become more of a factor.
Excellent article! I am wondering if you knew the outer dimensions of both the CFM and P&W engines (with Nacellle attached)?
I am looking for length, width, height and ground clearance.
TY.
I am wondering if you know the outside dimensions of both the P&W and CFM engines above with the Nacelle on? I am talking about width, length, height and ground clearance.
TY.
I think this vedio is one of the best
Vedio for jet engine
This article is extremely informative, and full of practicle mechanical language, that keeps a person from getting mired down in the Mathematical part of it.
I really dont mean to simplify the related issues of fielding this new engine, when I say that the new engine might be utilized in a three engine design.
Two conventional types being wing mounted and the new design center line at the rear. All three engines are used to achieve cruise altitude, and once attained, the conventional engines could be throttled back, to economize, while the new design would of course be spooled up to provide its own particular brand of “Magic”.
I’ll trade you this thought of mine, straight across for yours.
Please can you advice what is the purpose of the ring in the Turbo Fan on the PW engine?