With again thanks to others writers who have inputted valuable information, it has been found that excessive aspect ratio, by way of Reynolds Number, can effect gliding efficiency of a smaller, slower glider negatively as follows:

Reynolds number = Vc/v. Slower speed and shorter chord reduce Reynolds number. Kinematic viscosity of air v does not change.

Using Airfoil Tools, Coefficient of Lift/Coefficient of Drag at various Alpha was evaluated using a range of Reynolds Numbers from 50,000 to 1,000,000 for the Bleriot Eiffel 428 (thin undercambered) and the Wortmann FX 60 126 (thicker, less under camber).

It was found that Clift/Cdrag is significantly greater, at higher Reynolds number, particularly in a "sweet spot" around 6 degrees AOA. In this region, drag seems to "plateau", while Clift continues to rise.

At very low Reynolds Numbers, the Eiffel 428 showed a slightly better Clift/Cdrag ratio, but in ALL cases higher Reynolds number improved Clift/Cdrag ratio up to around 500,000 (will vary depending on air foil type), after which the polars seemed much more consistant.

With again thanks to other writers who have inputted valuable information, it has been found that excessive aspect ratio, by way of Reynolds Number, can effect gliding efficiency of a smaller, slower glider negatively as follows:

Reynolds number = Vc/v. Slower speed and shorter chord reduce Reynolds number. Kinematic viscosity of air v does not change.

Using Airfoil Tools, Coefficient of Lift/Coefficient of Drag at various Alpha was evaluated using a range of Reynolds Numbers from 50,000 to 1,000,000 for the Bleriot Eiffel 428 (thin undercambered) and the Wortmann FX 60 126 (thicker, less under camber).

It was found that Clift/Cdrag is significantly greater, at higher Reynolds number, particularly in a "sweet spot" around 6 degrees AOA. In this region, drag seems to "plateau", while Clift continues to rise.

At very low Reynolds Numbers, the Eiffel 428 showed a slightly better Clift/Cdrag ratio, but in ALL cases higher Reynolds number improved Clift/Cdrag ratio up to around 500,000 (will vary depending on air foil type), after which the polars seemed much more consistant.