The percent RPM is design value specified by the engine manufacturer. Let's consider this - the fan stage for example has a mechanical limitation on the maximum RPM it can withstand - that is NOT affected by altitude/aircraft speed/moon-phase... it is a function of physical strength. Beyond some RPM the thing will come unhinged and pieces fly all over the place. Now that of course is not to say 100%N1 is the fail speed, but it is related.

My specialism is in aircraft sound simulation - therefore I am very aware of the audio signature of engines, which not surprisingly is mostly related to RPM (engine whine tones), and I can assure you an engine turning at any given N1% RPM will turn at the same RPM irrespective of altitude. At least on all the modern era aircraft I have worked on (737NG/A320/lots of military jets...).

- GY :8

Garageyears

Well the percent RPM may be a design value specified by the engine manufacturer but it is not directly related to "redline speed" which is directly related to physical strength. Thus the same mechanical design used in multiple engine models may have different 100 % RPMs on their data sheets

Maybe we can get barit1 to cite some GE CF6 examples like how do you justify operating at 105-110% N1 on takeoff etc.

lomapaseo;

These are fixed numbers for a particular engine type and model, different per spool, for example (see link posted in Mechta's post#4):

RB211-535E4 on 757

100% N1 = 4500 LP RPM
100% N2 = 7000 IP RPM
100% N3 = 10611 HP RPM

But, as I wrote above, these numbers should not be confused with the 'red line' limits, which are also fixed but generally not 100%, and differ per rating.

regards,
HN39

lomapaseo digs me in the ribs:
Well - there's some history here. The CF6-50 fan is physically the same as the CF6-6, and GE in their wisdom did not change the arbitrary 100% N1 gage calibration point. Thus the fan spins faster to create more airflow => more thrust.

Also note that the CF6 (like other airline engines) is flat-rated up to a "corner point" OAT. This means that on cold days, the fan doesn't turn as fast, because the cold dense air doesn't need as much pumping to create the same thrust. Conversely on a hot day - up to that corner point OAT - the fan must spin faster. (Beyond that corner point temp - typically ISA+15 - fan speed cuts back to provide EGT protection). So at that sea level corner point OAT, fan speed reaches a peak around 110% IIRC.

NOW THEN - let's do a takeoff at DEN or MEX or a similar field elevation. At a given N1, less thrust is available. The airline doesn't want to take a huge payload penalty in the less dense air, but the engine manufacturer knows that maintaining thrust takes more N1. So some compromise is reached, the engine gets run a bit faster, but probably not all the airline would like. IIRC the CF6-50 mught run up the 118% N1 in this case. :uhoh:

minimumunstick
 

In a word yes. The size or type of the jet engine is not a factor in this statement.

I trust there is no hint of ambiguity in what I wrote.

One of the engine limits most likely to cause a reduction in RPM with increasing altitude is the engine RPM divided by the square root of the absolute temperature or N/ √ Θ as it is known in the trade. This is more likely to happen at very high throttle settings which if not reduced could cause the engine to approach the steady running surge boundary.

If you are looking for a simple intuitive explanation of why this (and other aspects of jet engines) should be so then I am afraid there is not one. The only way to understand what is behind the paragraph above for example (which is the PRL or pressure ratio limiter function of the engine fuel system) is to have a basic understanding of the aerodynamics of the compressor components.

I am not trying to blind you with science, the basics behind many engine issues are not difficult but they are complex in their interaction. People spend years studying the stuff.

GarageYears;

Please note that between the two conditions you gave in your post #19, thrust changes because atmospheric pressure and temperature, and Mach are different. They don't show that the FADEC maintains constant N1 in climb, do they?

EDIT:: Because of the difference in SAT and Mach the thrust values are not directly comparable. If you want to verify the 'aerodynamic similarity' rules I gave, then you must find two conditions with the same Mach, and either the same SAT, or with two RPM's that give the same N/ √ Θ (RPM corrected for temperature).

regards,
HN39

BTW: There is a gross overspeed criterion that must be demonstrated during certification testing. The case I'm familiar with (and the criteria may have changed over the decades) used the FAR propeller overspeed demo of 141% of the max normal operating RPM.

To get the fan to turn that fast, the fan nozzle will be butchered so it doesn't create much back pressure; thus the normal N1:N2 relationship is highly distorted. :eek:

John Farley

Thanks, I guess you are right.

Anyway reading the replies in this thread helps me grasp it better. I have also just re-opened my ATPL books and relearning stuff there which helps too.

Rated thrust and engine limits
 

This sketch compares schematically the thrust rating schedule to the engine never-exceed limits for N1, N2, EGT and P3. The full line shows the max. thrust setting recommended by the manufacturer that should not be exceeded intentionally and is normally implemented in the EEC of modern engines. It is selected so that it does not exceed any of the physical limits of the engine that are shown by the the dashed lines. If the rating is ignored and the engine is run 'on the limit', it would be P3-limited at ambient temperature up to 'A', N1-limited between temperatures 'A' and 'B', N2-limited between 'B' and 'C', and EGT-limited above temperature 'C'.

regards,
HN39

John Farley,

And I guess this has to do with the mach number changing with temperature?

No, it's air density changing with temperature. Since a fan or compressor tends to be a constant-volume pump at a given speed, it implies that as temperature changes at e.g. 100% rpm, the mass pumped is larger at low temperature, and lesser at high temperature.

By trimming rpm according to N/ √ Θ we hold mass airflow constant.

In my view it's not air density either. The surge boundary occurs at a certain N/ √Θ for any air density. I think the N/ √ Θ of a compressor is somewhat analogous to the AoA of an airfoil. Just as an airfoil stalls when driven to a too high AoA, a compressor stalls when driven to a too high N/ √ Θ. Isn't 'surge' the stalling of compressor airfoils?

As John Farley writes, high N/ √ Θ is associated with high throttle settings. It is also associated with low ambient temperature, because at higher temperature N1, N2 and EGT limits will prevail, as illustrated in my sketch (at a given altitude and Mach, constant thrust is constant N/ √ Θ, so the 'P3' limit shown in the sketch is almost constant N/ √ Θ).

I agree with barit1 that mach number has little to do with it.

regards,
HN39

N/ √ Θ isn't greatly involved with stall margin on a single-stage fan, which after all is pumping into a fixed-area fan exit nozzle at a moderate pressure ratio.

But N/ √ Θ has a great deal to do with stall margin in a multi-stage axial compressor. If you design the compressor so each stage has about the same alpha at design rpm, then each stage is sharing the pumping equally and all is lovely.

The problem occurs at off-design rpm - starting, idle, and especially acceleration. Here the front end is trying to pump more than the rear stages can handle, and there's a large alpha mismatch (too high in front, too low or negative in back). The problem is aggravated with HIGH inlet temperature (LOW N/ √ Θ). Thus the front stages are stall-prone.

Historically three remedies are available:
1) Bleed off the excess air (but quite inefficient - wasted pumping energy)
2) Multiple independent compressor spools which can change speed to re-match pumping.
3) Variable stator vanes to directly control alpha.

Each system had its advocates in history, but now all OEMs use all these devices.

HazelNuts39
After giving this some thought, I think barit1 is correct. For example, in the past, some jet engines used water injection to gain added thrust:

The maximum power a turbine engine can output depends largely upon the density or weight of the flow of the gases through the engine. Therefore, when the atmospheric pressure decreases or ambient air temperature increases, there is a loss in thrust. The power output can be boosted or restored by cooling the airflow with water or coolant.

Turbine D

Turbine D;

Yes, at a given RPM, but ... I thought we were discussing surge at this point.

Prior to that, we were discussing what causes RPM to change at constant throttle setting (essentially the OP).

regards,
HN39

"People spend years studying the stuff."

You know when you think of all the variables it becomes mind numbing. In some of the supersonic jets the guys must have a hell of a faith to push that lever forward.

Imagine the stress say at 5k feet on the turbine blades when you ask it to GO.

If anyone has a link how a supersonic jet fighter performs over M1 I would be grateful.

In simple terms - the gas turbine becomes less and less important, the afterburner (reheat) takes over. However - with improved aerodynamics, so-called "supercruise" (> m1.0 sans burner) is practical.

Returning to OP: - In my graph I sketched the limiting parameters quite arbitrarily. It so happens that the way they are drawn implies that when decreasing OAT at constant N2 (as occurs in climb), N1 increases, and EGT decreases.

Any ideas if that is generally so?

regards,
HN39

P.S. In any case, the changes of N1 and EGT are strictly due to the change of ambient temperature, whereas ambient pressure has nothing to with it. Therefore it would be wrong to attribute the changes to 'density' effects.

Jane
 

In answer to your question - no.

If it was like that I would have said so.

Please read all of my post.