Anyone know why does the low aspect ratio achieve higher angles of attack which stalling?

few definitions:
Wing loading is, of course, the area of the wing divided by the weight that it must lift. An 1800 lb airplane with a wing area of 100 square feet has a wing loading of 18 lbs/ft2.
Span loading is calculated by dividing the wingspan by the weight. An airplane with a weight of 1800 lbs and a span of 25 has a span loading of 72 lbs/ft.
Power loading is the weight of the airplane divided by the power propelling it. An airplane weighing 1800 lbs being pulled along by 200 hp has a power loading of 9 lbs/hp. (By some twisted semantic reasoning, a "high" power loading is a lower number卆n airplane with a power loading of 15 lbs/hp is said to have a "higher" power loading than one with 20 lbs/hp. This was probably dreamed up by the same guy who decided that larger AWG drill sizes would have smaller numbers.


First, consider climb rate. Basically, climb rate is linked most closely to power loading. It is determined by the engine horsepower or thrust available in excess of that required to maintain level flight. But span and span loading come into play as well. Everything else being equal, a shorter, lower aspect ratio wing has a higher induced drag, especially at low speeds, than a wing with a greater span and lower span loading. The shorter wing requires more horsepower just to maintain level flight at low speed. Since the power available is finite, this means less power remains for climb.

Low aspect ratio, "thick" wing, does not impair wing drag or lift. Good climb rate with short wings is achieved because of a low wing loading. Such a wing would also provide better spar depth and strength, the wider chord would mean less sensitive C.G. limits and it might achieve a greater angle of attack at stall.

A factor that combines with aspect ratio to affect lifting characteristics and stall angle is the thickness ratio of the selected airfoil. The thickness ratio is the ratio of camber to chord length. That is, the width of the airfoil at its thickest point divided by its tip to tail length. Thickness ratio effects lift by changing the nose shape of the airfoil. For a wing with a high aspect ratio and a moderate sweep, a large nose radius will increase the coefficient of lift. This arrangement will also increase the wing stall angle. For a wing with a low aspect ratio and a higher amount of sweep, the opposite is true. In this case, a sharp nose will produce leading edge vortices, which will counteract stalling and contribute to a greater maximum lift.

thanks eagle hannan.

why would the sharp nose leading edge vortices counteract stalling and contribute to a greater maximum lift.?

Buddy, I got the following from a document. I am posting the content related to your question along with the link to that page..





Design PLan

isnt anyone interested in why doesnt the LOW AR aircrafts stall at higher AOA

Aspect ratio has more to do with the induced drag that is due to the wing tip vortices - it is also called drag due to lift and is significantly larger than the zero-lift drag. The induced drag can only be minimized by increasing the AR (CDi = CL*CL/Pi*AR*e) - as you can see that it is in the denominator of the equation for induced drag in which e is the span efficiency factor - an empirical number. So for large aspect ratio wings, the total drag is low as is the case with sailplanes/gliders.

Low AR wings do stall at AoA that is actually lower than the large AR wings because pockets of separated flow overwhelm the lift producing capability.

I hope you got the idea - for deeper info check "Prandtl's Lifiting Line Theory" and "Drag Polar" on Google...

Good luck.

But i have read in many books that the AOA for Lo AR aircrafts to stall in much higher then for Hi AR wings?

Aerodynamics actually is slightly different. There are two other elements at play and they are camber (think of it as a curvature of the wing) and thickness distribution. Most short AR wing in use are actually symmetric airfoils and are therefore rarely used for lift of an actual airplane - and are mostly used for elevators, control surfaces and/or short fins on missiles. General aviation including including home built, military and commercial aircraft - all have well defined camber that helps produce lift even at zero AoA.
I very highly recommend a book by John Anderson - Theory of Flight. It is an introductory level book with good historic backgrounds. You will enjoy reading it...

Good luck.

Thanks for the insight. Have you seen the Mirage that the PAF flies? Its highly lo AR aircraft. I have met a couple of fliers of the mirage they have told me that this baby can fly at higher angles of attack and still not stall. The thing to do with the airflow i read in books but why nobody could answer me not even air force engineers?

Perhaps my earlier reply was not very clear - F15 has an AR of 3, F18 has AR of 3.5 close to Su27 etc... and yes you are correct - the short AR will not stall easily - some go higher than than 40 degrees - but here is the catch. Beyond 15 degrees AoA, pitching moment creates havoc - military fighter airplanes are slightly unstable for improved maneuverability but what actually happens is that in addition to the pitching moment problems, there is a steep drop in L/D ratio.

Thrust required is actually equal to the airplane weight divided by L/D - so as the L/D drops, thrust requirements increase tremendously - military airplanes have powerful engines and can therefore handle the thrust requirements but it is not true for all other airplanes - and by the way, AR less than 5 is called the short AR.

If you have indulged in RC model airplanes, you will notice that any shape can be made to fly as the weight of balsa wood models is very less compared to the thrust power of the engines (It is a fun hobby to get into) and therefore L/D does not matter.

Also note that power available from engine minus power required to overcome drag is called excess power which determines the rate of climb...

I hope this will help clarify my previous points.
Good luck.

Thanks Camber for the insight.

All the information on the pitching moment was greatly appreciated. But I still dont understand; my point is that I read in many books that the Airflow over the upper surface of the wing doesnt separate. Why? Is it because the airflow is bending inwards because of the hi pressure on the wing tips? and is it because of that it doesnt separate even at higher angles of attack?

OK Visionary - let me explain it this way:

The classical description of stall is a sharp drop in lift coefficient i.e. flow separation. Now there are two types of lift - potential lift and lift due to vortices (very common on delta wings). In case of short AR, the tip vortex, that becomes stronger with the angle of attack, covers most of the span and therefore even though the flow is not exactly two dimensional locally, wing continues to produce lift (this also explains the shift in aerodynamic center and pitching moment behavior). In other words although the potential lift drops because of the flow separation, the vortex lift increases due to proximity of the tip vortex. On large AR wings, this tip vortex is far from say mid-span and therefore when the flow separates on the wing, the tip vortex simply cannot help in producing lift. (Note that wing sweep helps a lot)

In all cases, pressure on the lower surface of the wing is higher than the pressure on the upper surface.

Keep in mind the airplane design is a very iterative and synergistic process - a lot of give and take even before a mission profile is established. I encourage you to pursue this career..

Hope this explanation helped you.

Good luck.