For subsonic aircraft, would adopting a higher wing loading decrease drag?

Question

First, let's assume that problems with approach/TO speeds can be fixed by high-lift devices.

Here we read that the wings create more drag than the fuselage. So, an improvement seems obvious; shrink the wings.

However, for a given weight, altitude, and airspeed, such a wing would need to fly at higher angle of attack, creating more induced drag. How does this compare against reduced parasitic drag?

If one wants to reduce induced drag, than one must fly at lower altitude. This means more parasitic drag - again. But does the former outweigh the latter?

Induced drag can be reduced further by the weight reduction from a smaller wing. How does this compare to the above?

Answer 1

Generally, you are right. Reducing wing area is reducing overall drag. Within limits.

Induced drag depends on speed and span loading. If the reduced wetted surface allows you to fly faster, induced drag will be lower, leaving more of the power budget to overcome viscous drag. However, if wing span is reduced, induced drag will be higher at the same speed, so it is better to reduce chord than to reduce span. This can be witnessed by the recent development in airliners where the reduced fuel flow of engines has allowed to build smaller wings with higher aspect ratios. A similar but older example is the Davis wing which combined extreme root airfoil thickness and an aspect ratio of 11 to minimize overall drag and keep wing weight reasonable. Other aircraft designed for long range even sported an aspect ratio of 14.5!

If one wants to reduce induced drag, than one must fly at lower altitude.

Don't focus on induced drag reduction alone! Please remember that the lowest overall drag is at the polar point where induced and zero-lift drag are equal. Move away from this point and overall drag for lifting the same weight increases.

Turbine engines profit from colder air. Flying higher means higher true air speed for the same dynamic pressure. Therefore, wing area should be sufficient to allow operation close to the tropopause. Going higher than that does not improve things, but staying well below the tropopause means to forego an obvious advantage. Only airplanes designed for low level missions will have a wing loading well above the 600 kg/m² which is typical for airliners. That's the limit mentioned above.

Answer 2

You might have heard of the "coffin corner" the U-2 pilots flew in. The U-2, at 90,000 feet, had only 9 knots between stall speed and Vmo. Airliners fly in a similar, but wider corner, with circa 15-30 knots between stall and Vmo. Vmo is where shockwaves start appearing on the upper wing surface, buffeting, engine surge and all kinds of bad stuff happen. Stall speed is relevantly high as low air density means the wing already operates at a high lift coefficient, also the wing has a lower CLmax value in transonic conditions.

The reason for this may not always be purely aerodynamic/stall related, aeroelastic behavior of the wing with high CL and transonic speeds is very complicated (even more so with complex wingtip devices) and I believe this will be the limiting factor.

For specifically designed wings, the wing can achieve higher transonic CL compared to subsonic, but I hugely doubt that an airliner wing is designed like that. Wings designed for high transonic lift are generally seen on fighter planes. As they have a lower aspect ratios and don't have wingtip devices, they are simpler from structural and aeroelastic standpoints.

Coming back to your question, the wing does not have much room to be shrunk, even in cruise conditions. Accommodating for the takeoff and especially landing is not as simple as slapping on high-lift devices, as in that case you'll still need to account for the failure of some high-lift devices and still be able to land the airplane.

The F-104 interceptor was designed with a very small wing (we had a F-104 gate guardian in front of our building in college, I was shocked when I first saw it with my own eyes) and resorts to flaps, slats and a boundary layer control system, which is air bled from the compressor to energize the boundary layer and delay the onset of stall. BLC obviously only worked when the engine was running, so with a flameout the approach speeds of the F-104 exceeded 200 knots. This was a contributing factor to the terrible safety record of the Starfighter, some air forces lost over 30% of the airplanes they acquired in accidents.