With our Reynolds number determined, we could isolate a specific plot that would more accurately represent our airfoil. Since Airfoil Tools only supplies Xfoil predictions for set Reynolds numbers, we had to round ours to the closest value. By clicking the “details” link next to the plot, we accessed the exact Xfoil data. To know if our wing is properly sized and will generate sufficient lift, we examined the angle of attack and lift coefficient data. By plugging in our weight calculations, mission requirements, and wing reference area into the lift coefficient equation, we determined the minimum lift coefficient required to achieve a maximum stall speed of 24 knots (as specified in Part 103). Unfortunately, we realized the Clark Y airfoil would stall long before producing a sufficient lift coefficient to maintain level flight at 24 knots. This means that our minimum speed would be greater than 24 knots, which would compromise the certifiability of our aircraft. In other words, our wing was simply too small.
To increase the wing reference area, we could either extend the wingspan or extend the length of the chord. Extending the wingspan would increase the aspect ratio but would come at the cost of extending our heavy aluminum spars. Alternatively, we could increase the length of the chord which would decrease the aspect ratio and therefore increase lift induced drag. Since ultralights are inherently draggy airplanes, we decided to extend the chord to 6ft and save a little weight.
Our second round of calculations can be broken down into two categories: stall and cruise. According to AC 103-7 section 21, our stall speed must legally be no greater than 24 knots. We can make our cruise speed whatever we want below 55knots, but ideally, we’d like it to be the velocity at which the airfoil has the lowest zero-lift drag coefficient.
For our stall calculations, we plugged in our new airfoil thickness, stall velocity, and kinematic viscosity at 59°F (a rough temperature average for San Francisco Bay Area weather). This produced a Reynolds number of approximately 1.81 E^4. We then selected the Clark Y plot with a Ncrit value of 9 and a Reynolds number of 2.00 E^4. We scrolled down the Xfoil prediction data and determined that at an angle of attack between 11 and 11.25 degrees, the airfoil would produce a lift coefficient between 1.3494 and 1.3609.