Sorry it was a typo by auto-correction ... wanted to say "Just as a reminder that the ..."
Chinese Engine Development
- Thread starter jackbh
- Start date
Sorry it was a typo by auto-correction ... wanted to say "Just as a reminder that the ..."
The US also start with military engines first even it's not even reliable. Just put WS-20 engine into C919 first and do more tests in this desperate situation. It will become more reliable over time.
Army is not passenger? For me just do more tests. Don't afraid of failure
I am sorry but airliners are a business that needs to make profits. With fuel and maintenance the major components to their cost, you can't compete if your planes runs less efficient engines. Why should they put up with the higher cost, or risk the lives of their passengers to become the Guinea pig? It aint a game. with that attitude you might as well work for boeing since you fit right in with the 737 Max crew.....The US also start with military engines first even it's not even reliable. Just put WS-20 engine into C919 first and do more tests in this desperate situation. It will become more reliable over time.
Army is not passenger? For me just do more tests. Don't afraid of failure
This isn't about efficiency. China's own HB engine manufacturing tech has not even reached the level we can discuss competitive efficiency. It's a question of reliability and safety. How long the engines can run for x amount of cost and how well they can run with well quantified and observed performance relating to safe operation.
WS-20 is presumably not a great option in this regard because it is a military HP turbofan and so was not designed or planned to be mass manufactured with the same tolerance a commercial passenger one should be. It also places other qualities before economics and safety, qualities that certainly cater towards military purposes like a production more suited for planned Y-20 acquisition rates (they don't want WS-20 production to cap Y-20 especially since Y-20 is almost certainly going to be branched into EW, tanker, AWACS etc roles in future). I'm sure there are many other factors that separate it from an engine designed for commercial applications. While it's true that plenty of military engines are derived from commercial ones and vice version, those cases do not prove that WS-20 was designed to be this way. It could be but that's not the rule.
Hendrik_2000
Lieutenant General
Well I found this explanation
In almost every large turbofan I see, the LP turbine stages outnumber the HP turbine stages by a factor of at least 2. Here's a photo of the RR Trent 900:
The Trent 900 (like most Rolls-Royce turbofans) is a 3-spool turbofan, rather than a 2-spool. But even here, my observation holds. There are far more LP turbine stages than the middle or high pressure stages.
Why is this?
Because the LP turbine extracts power for the fan, which requires the most power. The HP and IP turbine only extract power for their connected compressors - the LP turbine extracts power for the fan and the LP compressor. The fan does work on all airflow through the engine, the compressors only on a fraction (10:1 for a high bypass like the Trent 1000). The fan produces up to 75% of the thrust of the engine.
From : the cross section of the Trent 1000. It shows that the LP, IP and HP rotors have different rotational speeds, but does not list them. Another (older) presentation lists them as 3600. 6800 and 10200 RPM. Rotational speed goes down as the volume of the mass stream goes up.
An old textbook of mine gives the power P extracted from a turbine stage as:
P=m˙⋅u⋅vax⋅[tan(α2)+tan(α3)]P=m˙⋅u⋅vax⋅[tan(α2)+tan(α3)]
with
So the tangential speed of the turbine blade is in the equation for power extraction, which is composed of rotational velocity and blade radius. The faster the turbine turns, the more power can be extracted per stage, and the fewer stages required. Why then does rotational velocity go down with pressure? (HP = 10,200; IP = 6,800).
The reason is the constructional limits of the turbine. As the gas stream expands, the turbine blades become larger and are mounted at a greater axial distance, which results in greater centrifugal forces which are proportional to blade mass, rotational speed and distance from the axis. In order to limit centrifugal forces, the rotational velocity of the rotor is reduced: each subsequent stage turns at a lower RPM. Notice that a lower RPM can be compensated by mounting the blade further away from the rotational axis.
The LP rotor of an un-geared engine runs at the same rotational velocity as the fan. Its optimal RPM for the turbine may be higher, and that can be accommodated by the geared fan. As the bypass ratio gets higher, the LP turbine will be extracting a higher fraction of total power from the airstream - turboprops and turboshafts have a gearbox in between the LP shaft and the propeller/rotor, and the high bypass fan approaches the relative dimensions of a propeller...
In almost every large turbofan I see, the LP turbine stages outnumber the HP turbine stages by a factor of at least 2. Here's a photo of the RR Trent 900:
The Trent 900 (like most Rolls-Royce turbofans) is a 3-spool turbofan, rather than a 2-spool. But even here, my observation holds. There are far more LP turbine stages than the middle or high pressure stages.
Why is this?
Because the LP turbine extracts power for the fan, which requires the most power. The HP and IP turbine only extract power for their connected compressors - the LP turbine extracts power for the fan and the LP compressor. The fan does work on all airflow through the engine, the compressors only on a fraction (10:1 for a high bypass like the Trent 1000). The fan produces up to 75% of the thrust of the engine.
From : the cross section of the Trent 1000. It shows that the LP, IP and HP rotors have different rotational speeds, but does not list them. Another (older) presentation lists them as 3600. 6800 and 10200 RPM. Rotational speed goes down as the volume of the mass stream goes up.
An old textbook of mine gives the power P extracted from a turbine stage as:
P=m˙⋅u⋅vax⋅[tan(α2)+tan(α3)]P=m˙⋅u⋅vax⋅[tan(α2)+tan(α3)]
with
- m˙m˙ = mass flow [kg/s]
- u = tangential blade velocity [m/s]
- vaxax = axial gas velocity [m/s]
- α2α2 and α3α3 angles according to the figure below.
So the tangential speed of the turbine blade is in the equation for power extraction, which is composed of rotational velocity and blade radius. The faster the turbine turns, the more power can be extracted per stage, and the fewer stages required. Why then does rotational velocity go down with pressure? (HP = 10,200; IP = 6,800).
The reason is the constructional limits of the turbine. As the gas stream expands, the turbine blades become larger and are mounted at a greater axial distance, which results in greater centrifugal forces which are proportional to blade mass, rotational speed and distance from the axis. In order to limit centrifugal forces, the rotational velocity of the rotor is reduced: each subsequent stage turns at a lower RPM. Notice that a lower RPM can be compensated by mounting the blade further away from the rotational axis.
The LP rotor of an un-geared engine runs at the same rotational velocity as the fan. Its optimal RPM for the turbine may be higher, and that can be accommodated by the geared fan. As the bypass ratio gets higher, the LP turbine will be extracting a higher fraction of total power from the airstream - turboprops and turboshafts have a gearbox in between the LP shaft and the propeller/rotor, and the high bypass fan approaches the relative dimensions of a propeller...
Is that showing a newly built Y-20 with four WS-20 engines?
Not newly built. I think this is the same bird spotted flying last year.
Do we know if this is a prototype, or has it already entered the production pipeline and from now on all Y-20s will use WS-20?