Stall Speed at Altitude and Sea level
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From: Sydney, Australia
Stall Speed at Altitude and Sea level
Hi guys,
Can anyone spare some time for my amateur questions on Airspeed and Stall speeds.
I am aware that IAS is the airspeed shown by the ASI. Am I also correct to say that IAS will reduce with altitude? I am trying to find the stall speed for a Q400 (I can't find it published anywhere). To be rough about it, I read in a flight test evaluation article that it stalled at 15000 feet at 102 knots. I would like to therefor find the stall speed for it at sea level so used TAS = IAS sqrt p at sea level on row at altitude. I got round about 128 knots at sea level. I am however surprised that this is higher than the stall speed at altitude. Obviously I am interpreting the information incorrectly. Therefor, is stall speed somehow not related to this velocity at sea level?
Any help is much appreciated (I am in first year Aircraft Engineering)
Cheers
Can anyone spare some time for my amateur questions on Airspeed and Stall speeds.
I am aware that IAS is the airspeed shown by the ASI. Am I also correct to say that IAS will reduce with altitude? I am trying to find the stall speed for a Q400 (I can't find it published anywhere). To be rough about it, I read in a flight test evaluation article that it stalled at 15000 feet at 102 knots. I would like to therefor find the stall speed for it at sea level so used TAS = IAS sqrt p at sea level on row at altitude. I got round about 128 knots at sea level. I am however surprised that this is higher than the stall speed at altitude. Obviously I am interpreting the information incorrectly. Therefor, is stall speed somehow not related to this velocity at sea level?
Any help is much appreciated (I am in first year Aircraft Engineering)
Cheers
Joined: Jun 2008
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From: Brisbane
G'day jhurditch,
I'm no engineer but I think I can answer this one.
Let's say:
The lift equation is:
L = CL˝ρV2S
Where:
CL = the coefficient of lift of the wing. This is determined by the designer.
ρ = (rho) the density of the air. We can find this in ICAO Doc 7488. 1.225 kg/m3 at sea level.
V = the true airspeed. According to my Jepp CR-3 (Same as an E6-B) 102 KIAS is equal to 128 KTAS.
S = the surface area of the wing. This is set by the designer (64.1 m3 for the Q400)
Since CL and S are fixed we can ignore them and just say that Lift is proportional to ˝ρV2 which happens to have it's own name: q, or "dynamic pressure". q is the total air pressure at a point minus the static pressure or, in other words: indicated airspeed!
So this is all a roundabout way of saying that Lift is proportional to IAS, therefore the wing at the stalling angle (or any other angle) will produce the same lift at the same IAS. Therefore the stall IAS must be the same at all altitudes!
The 128 knot figure you calculated would be the TAS which, in nil-wind conditions, would equal the groundspeed. So the Q400 was indicating 102 knots but it was actually moving over the ground at 128 knots. So the answer to your question is: if the pilot maintains the same TAS as s/he climbs then yes, the IAS will decrease as the air density decreases. However, the beauty of it is that if you maintain the same IAS up to higher altitudes you will get a higher TAS! The decreasing air density also affects the engine's ability to maintain the same IAS, but that's another issue...
Like I said: I'm no engineer. This is just me figurin'
I'm no engineer but I think I can answer this one.
Let's say:
- the Q400 in question weighs the same at sea level as at FL150
- lift is equal to weight in level, unaccelerated flight
The lift equation is:
L = CL˝ρV2S
Where:
CL = the coefficient of lift of the wing. This is determined by the designer.
ρ = (rho) the density of the air. We can find this in ICAO Doc 7488. 1.225 kg/m3 at sea level.
V = the true airspeed. According to my Jepp CR-3 (Same as an E6-B) 102 KIAS is equal to 128 KTAS.
S = the surface area of the wing. This is set by the designer (64.1 m3 for the Q400)
Since CL and S are fixed we can ignore them and just say that Lift is proportional to ˝ρV2 which happens to have it's own name: q, or "dynamic pressure". q is the total air pressure at a point minus the static pressure or, in other words: indicated airspeed!
So this is all a roundabout way of saying that Lift is proportional to IAS, therefore the wing at the stalling angle (or any other angle) will produce the same lift at the same IAS. Therefore the stall IAS must be the same at all altitudes!
Am I also correct to say that IAS will reduce with altitude?
Like I said: I'm no engineer. This is just me figurin'
Last edited by Aviast; 21st April 2010 at 07:36.
Do a Hover - it avoids G
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From: Chichester West Sussex UK
All that is true if you ignore compressibility effects which are related to mach number.
However wing lifting efficiency is adversely affected by mach number and because the mach number increases as you climb at constant IAS, the stall speed will be higher at altitude. How big an increase is type dependant but it is likely to be a few knots.
However wing lifting efficiency is adversely affected by mach number and because the mach number increases as you climb at constant IAS, the stall speed will be higher at altitude. How big an increase is type dependant but it is likely to be a few knots.
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From: La Belle Province
All that is true if you ignore compressibility effects which are related to mach number.
However wing lifting efficiency is adversely affected by mach number and because the mach number increases as you climb at constant IAS, the stall speed will be higher at altitude. How big an increase is type dependant but it is likely to be a few knots.
However wing lifting efficiency is adversely affected by mach number and because the mach number increases as you climb at constant IAS, the stall speed will be higher at altitude. How big an increase is type dependant but it is likely to be a few knots.
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From: La Belle Province
I believe LM is referring to the fact that the usual conventions for forces on an aircraft are:
Lift - acts normal to the flight path
Drag - acts along the flight path
Weight - acts earthwards
Thrust - if net thrust, acts along the engine axis;
- if ram drag/gross thrust, ram acts along the flight path (as drag), gross thrust acts along the engine axis
Therefore, except at zero angle of attack and zero engine incidence (or an appropriate offsetting combination) there will be a thrust component which resoles normal to the flightpath and thus contribute to (or detracts from) aerodynamic lift.
Lift - acts normal to the flight path
Drag - acts along the flight path
Weight - acts earthwards
Thrust - if net thrust, acts along the engine axis;
- if ram drag/gross thrust, ram acts along the flight path (as drag), gross thrust acts along the engine axis
Therefore, except at zero angle of attack and zero engine incidence (or an appropriate offsetting combination) there will be a thrust component which resoles normal to the flightpath and thus contribute to (or detracts from) aerodynamic lift.
Joined: May 2005
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From: France
I once flew in a CJ (C.525 twin jet) with the owner who had asked about 'coffin corner'. We did stalls at FL100, 200, 300 and 400. The stalling IAS was pretty much the same at all altitudes (110 I think).
The CJ has a straight wing, cruises at about 360kts/M0.63. The weight was obviously a bit less at the higher altitudes due to the fuel burned off in the climb.
I was expecting 10 to 15 knots increase at FL400 compared with FL100, due to the effects mentioned, but I guess the actual TAS is just too low to have much compressibility?
The CJ has a straight wing, cruises at about 360kts/M0.63. The weight was obviously a bit less at the higher altitudes due to the fuel burned off in the climb.
I was expecting 10 to 15 knots increase at FL400 compared with FL100, due to the effects mentioned, but I guess the actual TAS is just too low to have much compressibility?
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From: La Belle Province
I once flew in a CJ (C.525 twin jet) with the owner who had asked about 'coffin corner'. We did stalls at FL100, 200, 300 and 400. The stalling IAS was pretty much the same at all altitudes (110 I think).
The CJ has a straight wing, cruises at about 360kts/M0.63. The weight was obviously a bit less at the higher altitudes due to the fuel burned off in the climb.
I was expecting 10 to 15 knots increase at FL400 compared with FL100, due to the effects mentioned, but I guess the actual TAS is just too low to have much compressibility?
The CJ has a straight wing, cruises at about 360kts/M0.63. The weight was obviously a bit less at the higher altitudes due to the fuel burned off in the climb.
I was expecting 10 to 15 knots increase at FL400 compared with FL100, due to the effects mentioned, but I guess the actual TAS is just too low to have much compressibility?
(On some sections) the Mach effect manifests as a lower stalling AoA with increased Mach; however, that applies to the "natural" stalling AoA i.e. in the absence of trips or artificial devices. If you had a situation where, say, the wing stalls at 16 deg at low Mach, but 14 deg at higher Mach, but you have a stall dressing or system which is designed to activate/be effective at 12 deg AoA, you'll struggle to see the mach effect.
Something like that may have been happening. (Afraid I don't know the details of the CJ wing and stall system, if any)
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I was expecting 10 to 15 knots increase at FL400 compared with FL100, due to the effects mentioned, but I guess the actual TAS is just too low to have much compressibility?
Joined: Sep 2002
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From: La Belle Province
@trim_stab
Sorry, but a freestream Mach number of 0.3 can easily show up as a localized Mach number approaching, if not in excess of, sonic, depending on the section. There are some aerofoils out there with VERY definite Mach-dependency, even at M0.2 and below.
Sorry, but a freestream Mach number of 0.3 can easily show up as a localized Mach number approaching, if not in excess of, sonic, depending on the section. There are some aerofoils out there with VERY definite Mach-dependency, even at M0.2 and below.
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From: Hampshire
Just a quick addendum to Aviast's otherwise excellent post:
1/2 rho V^2 has rho decreasing with altitude and V = TAS
1/2 rho0 Vi^2 assumes constant sea-level density and indicated airspeed (Vi), where rho0 = 1.225 kg/m^3)
One reason ASIs use a constant density (ISA sea level density) is that density cannot easily be measured and is also quite difficult to calculate, especially when using a mechanical instrument. Also, it has the handy side effect that for some aircraft (i.e. those that do not experience compressibility to any significant extent), scheduled speeds (e.g. Vs, Va, Vne, etc) are not density dependent and thus not temperature dependent or altitude dependent. Imagine how much harder it would be to operate an aircraft where all the safety-related speeds changed with temperature and altitude!
Note to the more technically advanced readers of the forum:
I think I am correct in stating that for a given weight and configuration, an aircraft will stall at the same Equivalent Airspeed at any altitude attainable within the flight envelope, but I didn't want to confuse our young questioner with discussions of IAS/CAS/EAS, pressure error corrections and scale altitude corrections! Anyway, for most light aircraft Vi is roughly Ve given the likely instrument errors and residual pressure errors, and provided the mach number is low enough then the Bernoulli's equation for non-compressible flow applies and thus IAS and TAS can be related simply by root sigma. However, I hope the university teaches its students about Bernoulli's equation for compressible flow and all the additional complexity that comes with it!
1/2 rho V^2 has rho decreasing with altitude and V = TAS
1/2 rho0 Vi^2 assumes constant sea-level density and indicated airspeed (Vi), where rho0 = 1.225 kg/m^3)
One reason ASIs use a constant density (ISA sea level density) is that density cannot easily be measured and is also quite difficult to calculate, especially when using a mechanical instrument. Also, it has the handy side effect that for some aircraft (i.e. those that do not experience compressibility to any significant extent), scheduled speeds (e.g. Vs, Va, Vne, etc) are not density dependent and thus not temperature dependent or altitude dependent. Imagine how much harder it would be to operate an aircraft where all the safety-related speeds changed with temperature and altitude!
Note to the more technically advanced readers of the forum:
I think I am correct in stating that for a given weight and configuration, an aircraft will stall at the same Equivalent Airspeed at any altitude attainable within the flight envelope, but I didn't want to confuse our young questioner with discussions of IAS/CAS/EAS, pressure error corrections and scale altitude corrections! Anyway, for most light aircraft Vi is roughly Ve given the likely instrument errors and residual pressure errors, and provided the mach number is low enough then the Bernoulli's equation for non-compressible flow applies and thus IAS and TAS can be related simply by root sigma. However, I hope the university teaches its students about Bernoulli's equation for compressible flow and all the additional complexity that comes with it!
Joined: May 2003
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From: Canada
Mach has an effect on maximum lift coefficient but Reynolds number is often the key player.
Actually in the case of the CitationJet, its airfoil is derived from the well studied HS-NLF213. At 100 knots calibrated between sea level and 41000ft you will see the wing Reynolds number go from about 5 million down to 3 million.
If you look at graph 28 in this report you will see that the full flaps 2d maximum lift coefficient drops from 3.7 to almost 3.4 due to the lower Re so this will explain most of the 10-15 knots stall speed increase between those altitudes. A similar effect would be present with flaps up.
As John Farley pointed out, the effect varies considerably between aircraft types. The example given here is on the high side.
As to why you would not see it by yourself in a 525 at 41000ft it probably has to do with airspeed calibration, if the lower fuel weight did not explain it. Certification agencies would be satisfied with a system that indicates constant stall speed as function of altitude even if it is not the case in reality, as long as what is shown to the pilot is the actual margin above stall speed for whatever weight is considered. In other words if your stall speed measurably increases with altitude, you probably want to have your avionics to indicate an airspeed that is such that the pilot does not have to worry about this effect.
Actually in the case of the CitationJet, its airfoil is derived from the well studied HS-NLF213. At 100 knots calibrated between sea level and 41000ft you will see the wing Reynolds number go from about 5 million down to 3 million.
If you look at graph 28 in this report you will see that the full flaps 2d maximum lift coefficient drops from 3.7 to almost 3.4 due to the lower Re so this will explain most of the 10-15 knots stall speed increase between those altitudes. A similar effect would be present with flaps up.
As John Farley pointed out, the effect varies considerably between aircraft types. The example given here is on the high side.
As to why you would not see it by yourself in a 525 at 41000ft it probably has to do with airspeed calibration, if the lower fuel weight did not explain it. Certification agencies would be satisfied with a system that indicates constant stall speed as function of altitude even if it is not the case in reality, as long as what is shown to the pilot is the actual margin above stall speed for whatever weight is considered. In other words if your stall speed measurably increases with altitude, you probably want to have your avionics to indicate an airspeed that is such that the pilot does not have to worry about this effect.
Last edited by Machdiamond; 29th April 2010 at 01:07. Reason: Clarification
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From: Australia
I should know the answer, but I don't.
In the 737 Classic, a clean approach to the stall at high altitude such as 35-37,000 ft with power at idle, is preceded by really strong buffet well before the stick shaker actuates. The buffet certainly cannot be confused with the type of turbulence experienced with CAT.
A landing flap approach to stall in the circuit with approach thrust, produces no buffet and the stick shaker is the first tactile indication. Why the quite different approach to stall, characteristics?
In the 737 Classic, a clean approach to the stall at high altitude such as 35-37,000 ft with power at idle, is preceded by really strong buffet well before the stick shaker actuates. The buffet certainly cannot be confused with the type of turbulence experienced with CAT.
A landing flap approach to stall in the circuit with approach thrust, produces no buffet and the stick shaker is the first tactile indication. Why the quite different approach to stall, characteristics?