S1210(2.25), S1223(2.3), S1223RTL (2.5).
This is a link to the airfoils and polars:http://www.kaiyipboxing.com/WIG/INDEX.HTM

Yes. But those polars were calculated for a Reynolds number of 2.5
10^6, which is way too high for your application. I did a run with
XFOIL on the S1210 and the S1223RTL with Re=1 10^6, which is closer to
your application. With a 1 m chord, you will get Re=7.5 10^5 at 25 mph
(around 11 m/s). The max c_L is still close, but from comparing XFOIL
results to wind tunnel tests on other profiles I know that XFOIL is
too optimistic. Usually even flapped airfoils don't go beyond a c_L of
about 2. But a normal flap on any of the airfoils you mentioned won't
give you more lift as you would have immediate flow separation on the
flaps upper surface at high angles of attack.

I would look for wind tunnel test data on flapped or other high-lift
airfoils and go from there. If the weight of your airplane and the
intended wing surface don't allow for a normal airfoil, try:

Investigation into high-lift systems such as fixed slats (such as used
on the Fieseler Fi 156 "Storch"), slotted flaps, or, maybe the best
option as a compromise between efficacy and simplicity is a Junkers
flap, which is just another small "wing" installed at a small distance
behind the main wing (Wikipedia and Google should give you some
insight on how this system is laid out and on which airplanes it was
used).

HTH,
Oliver