# If the pitch of a bypass fan blade in a turbofan is fixed how does it manage to provide thrust over such a large speed range?

A fixed pitch propeller is only good for aircraft that do around 100kts because the angle of attack on the propeller decreases as the forward airspeed of the aircraft increases, limiting the thrust available. Variable pitch propellers overcome this allowing the angle of attack to remain optimal at speed thus allowing faster cruise speeds.

However the pitch angle of the bypass stage of a turbofan is fixed but yet they do not seem limited by the same problem affecting a fixed pitch propeller. Turbofans provide sufficient thrust at take-off (very low forward speed) as well as much higher cruise speeds.

I'm trying unsuccessfully to understand the physics of this or more specifically: If the pitch of the turbofan is set optimally for cruise, why does it not stall at take-off due to a very high angle of attack? Conversely, if the pitch was set for high thrust at take-off (or some lower speed setting), shouldn't the angle of attack reduce as speed increases thereby limiting cruise speed attainable?

• Thanks for the responses. From what I understand there are 2 factors at play: a) The downwash is greater on a turbofan and this changes the local airflow around the blade causing an otherwise high angle of attack at low forward speed to be below stalling angle of attack b) The inlet design has the effect of reducing the range of air velocity that the blades actually see and in turn reduces the range of angles of attack. For example, at the start of the take-off roll, there is 0 aircraft airspeed but clearly there is a significant local forward airflow inside the duct.
– Ian
Dec 20, 2022 at 15:18
• appreciate any follow up posts! Thanks again.
– Ian
Dec 20, 2022 at 15:20

A large part of this is provided by the turbofan inlet design. Above pic is from Aircraft Gas Turbines by C.J.Houtman and was used earlier in this answer: the inlet shape slows the incoming air down at cruise speed, and trades velocity in for pressure.

The variables:

• At TakeOff thrust needs to be maximal, the fan blades need to suck all the air in, and the fan does so at maximum RPM.
• At cruise thrust is lower, the RPM is lower, and the air inflow is slowed down.

Indeed the variation in thrust requirements and airspeed create quite a difficult set of operational circumstances to tune, care must be taken to prevent stalling of the blade tips, and stalling of the inflowing air-stream around the inlet nose at lower velocities, which is why the inlet is optimised for these lower airspeeds.

Propellers operate at lower airspeeds than turbofans, and operate at constant RPM with variable blade AoA for optimal thrust efficiency. Turbofans would also benefit from variable blade AoA, and efforts have been made to develop them, but the cost/gain ratio seems to be prohibitive for now.

• You say that the inlet is designed to slow the airflow down; wouldn't this increase the AoA of the blades, thereby encouraging a stall? But then you say that the inlet is optimised for lower speeds, where a blade stall should be most likely. Dec 20, 2022 at 11:12
• @AdityaSharma The local AoA decreases with decreasing inflow speed. Stall is caused by an inflow with too high AoA, and a way of decreasing inflow speed helps avoid creating blade stall. Dec 20, 2022 at 12:34
• I respectfully disagree - decreasing the flow speed increases the local AoA. If you look at the figure in Sophit's answer, you can see that 'V' (incoming flow speed) induces an AoA onto the rotor blades, which reduces their effective angle of AoA. If this 'V' component got smaller with the same 'U' component, the angle 'β' will increase, and the angle between U and W will decrease - this angle is the induced AoA. For a given blade pitch, a decrease in induced AoA means an increase in effective AoA Dec 20, 2022 at 22:53
• @AdityaSharma Yeah the blade AoA reduces too much with increasing inflow velocity. Dec 21, 2022 at 1:26
• Indeed, and that reduction in AoA is what's preventing them from stalling. The intake design is only making it worse by reducing the inflow velocity, thereby increasing the AoA - the intake is not optimised for low speeds, it's optimised for high speeds where increase in blade AoA is actually needed. Dec 21, 2022 at 1:32

Compressor's (or fan or turbine) blades are designed to properly work in a quite ample range of angles of attack, more or less 15° around their design point.

The speed $$W$$ seen by the blade (and the relevant angle of attack $$\alpha$$) is given by the sum of two terms:

1. the speed $$U$$ due to the rotation of the blades;
2. and the speed of the incoming airflow $$V$$.

This can be seen for example in the following plot (slightly modified by me) taken from this report (blades are seen from above and rotate to left; flight direction is upward):

The rotating speed $$U$$ can be more or less controlled via the throttle while the speed $$V$$ of the incoming flow is driven toward the blades (of the first compressor's stage) by an extremely important component that lies in front of the engine and that most of the time passes unnoticed: the inlet.

The main job of an inlet is to reduced the speed of the incoming air to some Mach 0.4, whatever high the speed of flight is. The range from Mach 0 to 0.4 is the limit within which the compressor's blades (and especially their tips) work in a proper way.

If the speed of flight is high subsonic (Mach 0.4 to 1), then the inlet simply works as a divergent inlet (diffuser): its section gets bigger and bigger, the pressure gets higher and higher and the speed gets lower and lower (the compressor thanks Bernoulli).

Finally, if the speed of flight is supersonic, in front of the inlet one or more shockwaves form, across which the speed reduces to high subsonic and afterward a divergent portion of the inlet finishes the work reducing again the airspeed till Mach 0.4.

This is how inlets of a modern jetliner, the Concorde or the Blackbird work(ed) to supply their compressors with a relative slow and stable airflow.

• Thanks Sophit. Would it be correct to assume that the enclosed nature of fan blades (preventing spill over the tips) as well as the overlap (or very small gap) between blades is the reason they have a higher stalling angle of attack compared to props? It is harder for air to separate from the upper surface of a blade when there is another blade just nearby to help direct airflow? Thanks again, this is all starting to make sense.
– Ian
Dec 21, 2022 at 1:10
• @Ian The exact opposite of that would be correct to assume. Without the duct, there will be blade-tip vortices which will actually reduce the effective AoA of the blades, and that will be beneficial at preventing a stall. So the duct is not helping with blade stall prevention at all - it's doing the exact opposite. However, the closely positioned blades are indeed helping in preventing the stall (by inducing a downwash onto the consecutive blades). Dec 21, 2022 at 1:44
• @Ian: the ratio between the distance $s$ between two blades and their chord $c$ is called solidity $\sigma = s/c$. It does has an impact on the efficiency of the rotor and normally has an optimal value slightly bigger than 1. Dec 21, 2022 at 2:36
• @Ian: if we analyze the throat between two blades, we have an area of high pressure under one blade and an area of low pressure over the other one. Since in general air moves away from high pressure and toward low pressure, passing through the throat air is "pushed" against the upper surface of the blade and that helps against stall. Obviously there are limits to this effect: if the throat is too large this effect becomes negligible and if the throat is too small the airflow get clogged, that's why the optimal value of 1. Dec 21, 2022 at 3:02
• all good. for anyone familiar with sailing, this reminds me of the "slot" between the front sail (jib) and the back sail (main). The slot plays a huge role in optimising the performance of both sails.
– Ian
Dec 21, 2022 at 4:14

The AoA of the fan blades varies throughout the speed range, very similar in principle to a fixed pitch propeller. However, the change in AoA also depends on the propulsive efficiency of the engine.

This variation in AoA is determined by the change in "down" wash at the fan blades (think of the fan blade like a wing). This downwash is for to two reasons: forward speed (TAS), and the inherent downwash (jet-wash) due to the production of lift (thrust) itself. On a turboprop engine with high propulsive efficiency, this latter downwash is low, but is relatively high on turbofans (and is even higher on turbojets). The net downwash as experienced by the fan blades is a sum of the two.

When a turbofan is operating on ground at zero TAS, the effective TAS at the fan blades is quite high. This is due to the aforementioned downwash at the fan blades due to the production of thrust itself. This downwash reduces the effective AoA of the blades enough to prevent it from stalling. As aircraft TAS increases, the net downwash increases, causing the effective AoA of the blades to decrease.

However, on a turboprop with a much higher propulsive efficiency, the initial downwash (due to thrust production) is relatively low, and it's influence on the effective AoA is much less. When the aircraft TAS increases through the same speed range as in the turbofan case, the percentage increase in the net downwash is much higher, and so is the change in AoA. The only way to maintain a reasonable AoA is through the use of variable pitch propellers.

In the turbofan case, the change in AoA as the speed increased was significant, but still low enough - a variable pitch mechanism would be able to provide much less improvement, meaning that it's use is not feasible.

If you find this difficult to accept, think of the extreme case: if the initial downwash was infinite (blade pitch would be 90°), an increase in TAS through any speed range would have no effect on the AoA, and so a variable pitch system will be absolutely useless in this case.

One solution I haven’t seen addressed in the existing answers is compressor bleed valves. The speed with which the air passes over the compressor blades is a function of the pressure differences ahead of, and behind, the blades. The whole point of multi-stage compressors is for the pressure inside the engine to keep increasing as the air flows from inlet to combustion chamber.

At low airspeeds, and increasing pressures as the engine RPM increases, this means pressure is building up behind the blades faster than the plane is accelerating to create increased inlet pressure. This is what typically causes compressor stall… pressure increasing behind the blade reduces the speed of the flow over the blades.

Multi-stage engines will often have a bleed valve in the rear compressor stages that opens to let some air escape until airspeed increases. This alleviates the pressure building up behind the compressor blades, allowing faster flow, reducing blade AOA, alleviating stall.

Thus the fixed blade angle can be optimized for cruise performance and the engine manages effective AOA by managing pressure gradients inside the compressor. The bleed valves are pneumatic and operate autonomously with no crew intervention.

• Good point, I addressed it en passant in my answer with the picture of the Concorde's engine but it's also important at takeoff. Dec 23, 2022 at 7:22