What airfoil parameter should I focus on to achieve the maximum service ceiling?

Question

I am working on a design competition project where I have to design a low subsonic aircraft for high altitude performance.

From what I have studied service ceiling will depend on excess power available so I should focus on minimising drag or minimising cl/cd.

Problem is I haven't seen any aircraft in service that uses high lift airfoils, why is it so? What parameter should I focus on to design a better wing?

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Some information about the aircraft and the constraints set by the rules of the competition:

The aircraft has a 3.5 metre wingspan, weighing in the ballpark of 25 kg, although maximum weight allowed is 50 kgs. It has to carry 5 kg payload to >6000m altitude on electric propulsion.

Answer 1

Given that you must use electric propulsion and have a size constraint, the usual criterium of selecting for the maximum product of lift coefficient and flight Mach squared will not apply here. Instead, this can be dealt with using incompressible aerodynamics.

Therefore, the figure of merit is to maximize $\frac{c_L^3}{c_D^2}$ in order to stay airborne with the smallest possible power. Given that you not only need to lift your airplane's mass up but also need to accelerate it so dynamic pressure can stay constant, you will achieve the highest altitude when you can make most of the limited energy stored at take-off.

Since induced drag at the optimum operating point will be three quarters of total drag and span is limited, you need to save mass where ever you can. This includes the selection of a capable airfoil because a thick wing and a high lift coefficient will improve structural efficiency and reduce wing area. However, make sure to use all the mass allowance to pack as much battery capacity as possible. If the rules allow, consider dropping empty battery packs enroute.

Without having performed an exhaustive comparison, I would start with the Daedalus airfoil set (DAE-11 at the root to DAE-31 at the tip) and adapt those to the Reynolds numbers you will meet in flight.

Answer 2

Don't worry about the airfoil at the start. You need to focus first on getting the wing area right.

Do you have a speed requirement at altitude? For example, do you have to be able to maintain a spot over the ground in some expected headwind?

Your airplane will fly at CL=W/(q*S).

q=0.5rhoV^2 is determined by your flight condition. If you are free to choose V, then things are more complex.

You want to fly at CL for best L/D. For a given drag polar, that is a specific value of CL.

If you know your target CL, flight condition, and weight -- you can solve for the wing area S that will make it happen.

You want as much span as you can get -- do you have a span constraint? If not, you will be limited by structures. Keeping S constant, increasing span will shrink the chord and also the thickness of the airfoil. Such a thin spindly wing will not support the load.

Once you know chord and flight condition, you will know the cruise Reynolds number. You also know the cruise lift coefficient. For a straight, high aspect ratio wing, your sectional lift coefficient will be close to the cruise lift coefficient. You should start by selecting an airfoil with a design cl that matches your cruise CL.

This process is iterative -- you make a guess about the drag polar, calculate everything through, and then update your drag polar estimate based on the real wing area, the airfoil selected, etc. Then iterate again. A few times through and you'll converge on a good wing loading and airfoil for your point.