Sharp vs rounded leading edge

Any dialogue would be appreciated.
I am scratch building a plane and right now the leading edge is very sharp-but I remember someone on this forum stating that a stunt airplane would turn better with a rounded leading edge.
Is this true?
If so how does a plane with a sharp leading edge turn compared with one with a rounded leading edge.
Would appreciate any feedback on the subject.
BudS of the Tampa Bay Line flyers.
"Your flying field doesn't have to be level, but it helps!"

Hi Bud,
I'm only going to give you a generalization as I have never made a real comparision. The few stunt ships I have built all have a rounded or blunt leading edges and they turn excellent. I only know from experience that the RC ships I have built with a sharp leading edge on the wing and on the stab do not track as well as the rounded type, they tend to be contantly seeking a natural position in the sky and are harder to trim to fly level on their own. Also there are many older stunt ships as in old time and classic don't have the thick rounded leading edge as do the more modern stunt ships of today, Yet many of them are excellent turners if their built light. The thicker airfoils will carry a little more weight and still give an excellent turn.

HI Bud. If I remember my reading correctly, the sharp leading edge has a tendency to cause the air flow to separate (and get very turbulent) from the wing at certain angles of attack and the wing then loses lift. A blunt airfoil allows the air flow to remain over the wing for a longer period of time (and much smoother) and at a greater angle of attack and thus loses lift much later or not at all. Not very scientific, but our engineers might all be getting warmed up for the NATS and Internats. Maybe one of them will come in here with a more technical explanation. Al Rabe? Ted Fancher? Brett Buck?

Bud,

Certainly not an expert in aerodynamics but from many years experience and playing with different designs. I would agree with the others that blunt leading edges are better. I played with a combat design back in the 70s and built several different airfoils for this plane before making a final decision. The rounded leading edge was much more "solid" in level flight and allowed the plane to turn without stalling. I have built all my planes with rounded leading edges since that time.

Good luck with your project. Hope it comes out a "killer" plane.

Unless the "sharp" edge is laser perfect, the plane will tend to "hunt". Smooth edges are the only way to go, IMHO.

you have to define what you mean by "blunt" and "sharp". The only explication I've ever seen was from Wild Bill. According to WB(going form memory) a "sharp" leading edge has a diameter of 2-3% of thickness, a medium 5-7%, and a blunt LE is has a dimeter over 10-12% of thickness. Those are pretty self-explanatory. Something like a NACA 0018 has a LE diameter around 20% of thickness. Current fashion makes a blunt LE around 30% of the thickness.

Personally, a NACA 0018 works just great, but I've never had any trouble going down to a LE diameter of 10-15% of thickness, as long as the wing loading is lower rather than higher. Heavy planes(anything over 10 oz/sq.ft without flaps, or 12 oz/sqft with flaps) will have trouble regardless of the LE radius, unless you have GOBS of power.

The only direct back-to-back experience I've had was with a friend's Nobler. It was very sloppy on corners and stalled quite easily. It took me a couple of days before I noticed the leading edge was made from triangular stock that had barely had the sharp edge taken off it so I suggested he sand it back further and blend it in. We did it then and there at the field by sanding back <1/8" and used some 5 minute epoxy to seal it. Flight performance was immediately very much improved.

Hi

I started this yesterday and have not had opportunity to review other posts which may cover this anyway.

All other things being equal, a profile with a sharp LE will stall in both positive and negative lift before one with a large nose radius (blunt). That is the blunt profile will have a wider incidence range, before stall. As the relationship between lift and angle of attack is very nearly the same for all profiles, this translates to a wider lift range.

This a well know aerodynamic principle and is easily verified by studying a range of L/D polars from various profiles.
A thick profile will have a larger nose radius and be capable of higher angles of attack and greater lift than a thin profile of the same type. Also comparing two profiles of the same thickness, the one with the most forward maximum thickness point will generally have the largest nose radius and will therefore have a greater incidence range and maximum lift.

Look at Eppler 472, This is just 12% thick but the max thick point is at 17% of chord rather than the more usual 30%. This profile therefore has a very blunt nose for it's thickness and is capable of max lift better than most 18% profiles.

On the other hand laminar flow profiles tenf to have rearward max thick points around 50 or 60% of chord. While the drag is lower, the opperational lift range tends to be reduced also. Blunt airfoils tend to have more profile drag than sharper ones but in the bigger picture, this is not very significant to teh total drag budget and not as important as lift range especially for stunters and gliders.

The best way to look at this is that thickness and nose radius relate to the operational lift range of the profile while camber determines the most efficient lift/drag trim and therefore the crusing speed of the aircraft.

It is not only leading edge (LE), it is whole curvature of the airfoil surface. The trick is, that airflow cannot stick at surface if the radius at that place is too small. Other thing is, that ability to stick is smaller and smaller (stability) as airflow goes down the chord. Thus the smallest radius is at very front point of airfoil and is known as radius of LE. The rest of curvature must have smoothly lower and lower “radius” – up to trailing edge where the surface is mostly flat. Those mentioned NACA airfoils coming from analytical equations are done exactly that way. Beside difficult to understand equation, that curvature is easy to explore from thickness and point of maximal thickness of airfoil. That is telling how the surface runs from LE arc and thus what it will do around stall point.

The LE radius means at what angle of attack the stream separates at LE and the rest of surface curvature (the thickness) says how it propagates from trailing edge.

So you want blunt LE to allow higher angle of attack and thus lift and thick airfoil to be smooth at stall.

Ok, I understand that part Igor, but when (In theory)does blunt become TOO blunt?

From my experience with my, and others Jamison Specials, it is possible to get carried away with blunting the leading edge, and you get great lift, great round loops, great tracking in rounds, but lousy wind penetration, weird things happened in square corners (like hitting a brick wall, bam, then pull out of corner), but not a stall, and need LOTS of power to get through the corners.
Thoughts?
Eric

OK, an example. NACA airfoil with thickness 12%, 18% and LE radius 1%, 3% and 5% at theoretically analyzed at RE=200 000:

12% x 1%:

NACA 0012-43

Re = 200000
á Cl Cd Cm 0.25 TU TL SU SL L/D
<°> <-> <-> <-> <-> <-> <-> <-> <->
0,0 -0,000 0,00870 -0,000 0,639 0,639 0,991 0,991 -0,000
1,0 0,120 0,00882 -0,002 0,576 0,690 0,991 0,991 13,570
2,0 0,239 0,00923 -0,003 0,486 0,733 0,991 0,991 25,914
3,0 0,358 0,01073 -0,005 0,250 0,772 0,991 0,991 33,360
4,0 0,476 0,01139 -0,006 0,206 0,806 0,990 0,991 41,757
5,0 0,591 0,01289 -0,008 0,160 0,837 0,989 0,991 45,816
6,0 0,700 0,01510 -0,009 0,043 0,861 0,978 0,990 46,345
7,0 0,802 0,01687 -0,010 0,011 0,882 0,962 0,990 47,511
8,0 0,893 0,01859 -0,012 0,008 0,902 0,936 0,990 48,042
9,0 0,963 0,02114 -0,013 0,007 0,916 0,860 0,990 45,543
10,0 0,855 0,06863 -0,007 0,006 0,930 0,017 0,990 12,459
11,0 0,901 0,07841 -0,007 0,005 0,939 0,011 0,989 11,497
12,0 0,936 0,08975 -0,007 0,005 0,947 0,008 0,989 10,433
13,0 0,951 0,10349 -0,007 0,005 0,957 0,008 0,989 9,186
14,0 0,943 0,11795 -0,008 0,004 0,961 0,007 0,990 7,992
15,0 0,920 0,13176 -0,008 0,004 0,966 0,007 0,990 6,980
16,0 0,884 0,15266 -0,008 0,004 0,970 0,007 0,990 5,794
17,0 0,840 0,17189 -0,009 0,003 0,976 0,006 0,990 4,888
18,0 0,790 0,19639 -0,009 0,003 0,978 0,005 0,990 4,025
19,0 0,738 0,21944 -0,010 0,002 0,980 0,006 0,990 3,362
20,0 0,684 0,25237 -0,010 0,002 0,982 0,006 0,990 2,712

12% x 3%:

NACA 0012-83

Re = 200000
á Cl Cd Cm 0.25 TU TL SU SL L/D
<°> <-> <-> <-> <-> <-> <-> <-> <->
0,0 0,000 0,00942 -0,000 0,636 0,636 0,989 0,989 0,000
1,0 0,120 0,00907 -0,001 0,577 0,686 0,990 0,989 13,265
2,0 0,240 0,00985 -0,002 0,498 0,729 0,989 0,989 24,361
3,0 0,359 0,01276 -0,003 0,094 0,766 0,986 0,989 28,105
4,0 0,476 0,01351 -0,004 0,062 0,800 0,982 0,987 35,234
5,0 0,591 0,01449 -0,005 0,045 0,829 0,978 0,985 40,792
6,0 0,702 0,01559 -0,006 0,036 0,852 0,969 0,985 45,006
7,0 0,808 0,01692 -0,007 0,031 0,874 0,956 0,983 47,763
8,0 0,905 0,01865 -0,008 0,026 0,890 0,921 0,982 48,506
9,0 0,975 0,02246 -0,009 0,022 0,907 0,791 0,981 43,403
10,0 0,922 0,05896 -0,007 0,019 0,917 0,155 0,981 15,628
11,0 0,975 0,07260 -0,007 0,017 0,929 0,089 0,980 13,434
12,0 1,028 0,08445 -0,007 0,016 0,937 0,063 0,980 12,171
13,0 1,069 0,09641 -0,007 0,015 0,942 0,049 0,981 11,090
14,0 1,100 0,10941 -0,007 0,015 0,950 0,039 0,981 10,057
15,0 1,123 0,12295 -0,007 0,014 0,956 0,036 0,982 9,130
16,0 1,133 0,13737 -0,008 0,013 0,960 0,030 0,982 8,249
17,0 1,134 0,15372 -0,008 0,012 0,962 0,026 0,983 7,377
18,0 1,126 0,17315 -0,008 0,012 0,964 0,024 0,984 6,504
19,0 1,110 0,19211 -0,008 0,011 0,967 0,023 0,985 5,779
20,0 1,089 0,21052 -0,009 0,010 0,971 0,027 0,985 5,172

12% x 5%:

NACA 0012-103

Re = 200000
á Cl Cd Cm 0.25 TU TL SU SL L/D
<°> <-> <-> <-> <-> <-> <-> <-> <->
0,0 -0,000 0,00943 -0,000 0,634 0,634 0,983 0,983 -0,000
1,0 0,120 0,01208 -0,001 0,092 0,683 0,985 0,983 9,967
2,0 0,240 0,01259 -0,001 0,067 0,725 0,983 0,983 19,060
3,0 0,358 0,01307 -0,002 0,056 0,762 0,980 0,982 27,432
4,0 0,475 0,01385 -0,003 0,048 0,796 0,978 0,981 34,314
5,0 0,590 0,01482 -0,003 0,040 0,822 0,972 0,980 39,800
6,0 0,700 0,01585 -0,004 0,035 0,846 0,961 0,979 44,186
7,0 0,802 0,01730 -0,005 0,032 0,867 0,928 0,978 46,375
8,0 0,756 0,04929 -0,004 0,030 0,883 0,122 0,978 15,335
9,0 0,829 0,05798 -0,004 0,028 0,898 0,088 0,977 14,290
10,0 0,897 0,06711 -0,004 0,026 0,911 0,067 0,977 13,372
11,0 0,962 0,07628 -0,004 0,025 0,920 0,058 0,977 12,610
12,0 1,021 0,08565 -0,004 0,022 0,930 0,059 0,977 11,923
13,0 1,072 0,09678 -0,005 0,020 0,936 0,056 0,977 11,080
14,0 1,114 0,10944 -0,005 0,018 0,941 0,052 0,977 10,183
15,0 1,148 0,12265 -0,005 0,017 0,946 0,046 0,978 9,362
16,0 1,175 0,13820 -0,005 0,017 0,951 0,043 0,978 8,499
17,0 1,193 0,15365 -0,006 0,017 0,957 0,041 0,978 7,764
18,0 1,204 0,17107 -0,006 0,016 0,960 0,038 0,978 7,036
19,0 1,208 0,18694 -0,006 0,015 0,961 0,036 0,979 6,460
20,0 1,205 0,20596 -0,006 0,015 0,963 0,035 0,979 5,852

18% x 1%:

NACA 0018-33

Re = 200000
á Cl Cd Cm 0.25 TU TL SU SL L/D
<°> <-> <-> <-> <-> <-> <-> <-> <->
0,0 -0,000 0,01159 -0,000 0,517 0,517 0,982 0,982 -0,000
1,0 0,124 0,01287 -0,003 0,293 0,568 0,980 0,984 9,659
2,0 0,248 0,01316 -0,006 0,270 0,611 0,976 0,985 18,862
3,0 0,371 0,01355 -0,008 0,249 0,649 0,969 0,986 27,394
4,0 0,493 0,01406 -0,011 0,231 0,681 0,962 0,987 35,098
5,0 0,614 0,01470 -0,014 0,216 0,713 0,952 0,987 41,793
6,0 0,734 0,01554 -0,016 0,202 0,742 0,940 0,987 47,218
7,0 0,847 0,01662 -0,019 0,185 0,767 0,923 0,987 50,992
8,0 0,954 0,01800 -0,022 0,172 0,794 0,901 0,987 53,036
9,0 1,044 0,02407 -0,024 0,010 0,815 0,830 0,987 43,371
10,0 1,128 0,02753 -0,026 0,008 0,837 0,779 0,985 40,975
11,0 1,195 0,03221 -0,028 0,006 0,853 0,718 0,985 37,087
12,0 1,239 0,03866 -0,030 0,005 0,872 0,646 0,984 32,051
13,0 1,259 0,04829 -0,031 0,005 0,885 0,552 0,984 26,073
14,0 1,134 0,10471 -0,015 0,004 0,899 0,011 0,983 10,829
15,0 1,142 0,11800 -0,016 0,003 0,910 0,010 0,983 9,680
16,0 1,136 0,13231 -0,017 0,003 0,920 0,010 0,983 8,585
17,0 1,116 0,14966 -0,018 0,002 0,930 0,010 0,983 7,457
18,0 1,085 0,16874 -0,018 0,002 0,937 0,009 0,984 6,428
19,0 1,044 0,18547 -0,019 0,002 0,942 0,008 0,985 5,628
20,0 0,997 0,21155 -0,020 0,001 0,950 0,008 0,986 4,711

18% x 3%:

NACA 0018-53

Re = 200000
á Cl Cd Cm 0.25 TU TL SU SL L/D
<°> <-> <-> <-> <-> <-> <-> <-> <->
0,0 0,000 0,01159 -0,000 0,523 0,523 0,979 0,978 0,000
1,0 0,125 0,01176 -0,002 0,459 0,571 0,977 0,979 10,622
2,0 0,249 0,01310 -0,005 0,276 0,612 0,973 0,981 19,039
3,0 0,373 0,01366 -0,007 0,238 0,649 0,966 0,982 27,302
4,0 0,496 0,01430 -0,009 0,210 0,681 0,959 0,982 34,661
5,0 0,616 0,01518 -0,011 0,183 0,712 0,948 0,982 40,593
6,0 0,735 0,01623 -0,014 0,157 0,740 0,935 0,983 45,284
7,0 0,849 0,01766 -0,016 0,131 0,766 0,916 0,983 48,092
8,0 0,959 0,01947 -0,018 0,105 0,791 0,891 0,982 49,251
9,0 1,061 0,02191 -0,020 0,080 0,812 0,856 0,982 48,445
10,0 1,154 0,02524 -0,022 0,058 0,834 0,808 0,981 45,733
11,0 1,236 0,02982 -0,024 0,041 0,849 0,744 0,981 41,456
12,0 1,307 0,03576 -0,025 0,034 0,868 0,672 0,979 36,539
13,0 1,358 0,04539 -0,026 0,025 0,881 0,574 0,980 29,916
14,0 1,392 0,05841 -0,027 0,022 0,893 0,469 0,979 23,833
15,0 1,416 0,07466 -0,027 0,019 0,907 0,372 0,979 18,972
16,0 1,435 0,09244 -0,028 0,016 0,915 0,296 0,979 15,524
17,0 1,450 0,10993 -0,028 0,015 0,924 0,245 0,979 13,192
18,0 1,458 0,12872 -0,028 0,013 0,934 0,203 0,979 11,323
19,0 1,456 0,14901 -0,029 0,012 0,939 0,166 0,980 9,773
20,0 1,446 0,17176 -0,028 0,011 0,944 0,131 0,980 8,419

18% x 5%:

NACA 0018-73

Re = 200000
á Cl Cd Cm 0.25 TU TL SU SL L/D
<°> <-> <-> <-> <-> <-> <-> <-> <->
0,0 0,000 0,01164 -0,000 0,526 0,526 0,976 0,976 0,000
1,0 0,125 0,01181 -0,002 0,469 0,572 0,974 0,977 10,606
2,0 0,250 0,01299 -0,004 0,288 0,613 0,970 0,978 19,240
3,0 0,374 0,01398 -0,006 0,205 0,649 0,963 0,979 26,718
4,0 0,496 0,01495 -0,008 0,158 0,680 0,955 0,980 33,158
5,0 0,616 0,01604 -0,010 0,122 0,711 0,942 0,980 38,395
6,0 0,733 0,01738 -0,012 0,094 0,739 0,926 0,979 42,185
7,0 0,847 0,01879 -0,014 0,080 0,764 0,907 0,979 45,062
8,0 0,955 0,02069 -0,016 0,064 0,789 0,879 0,979 46,157
9,0 1,057 0,02299 -0,017 0,055 0,809 0,842 0,979 45,964
10,0 1,150 0,02606 -0,019 0,049 0,831 0,793 0,978 44,132
11,0 1,229 0,03066 -0,020 0,042 0,847 0,723 0,978 40,082
12,0 1,291 0,03812 -0,021 0,037 0,864 0,625 0,978 33,869
13,0 1,334 0,05026 -0,022 0,034 0,878 0,497 0,977 26,549
14,0 1,361 0,06874 -0,022 0,031 0,890 0,352 0,977 19,799
15,0 1,392 0,08758 -0,022 0,028 0,903 0,257 0,977 15,898
16,0 1,428 0,10455 -0,022 0,026 0,911 0,204 0,977 13,657
17,0 1,458 0,12253 -0,022 0,025 0,919 0,162 0,977 11,901
18,0 1,489 0,13837 -0,023 0,022 0,928 0,145 0,977 10,764
19,0 1,512 0,15553 -0,024 0,020 0,935 0,127 0,978 9,723
20,0 1,529 0,17387 -0,024 0,020 0,940 0,113 0,978 8,792
21,0 1,539 0,19252 -0,025 0,018 0,945 0,103 0,978 7,995
22,0 1,542 0,21293 -0,025 0,017 0,950 0,091 0,979 7,243
23,0 1,532 0,23448 -0,025 0,016 0,956 0,084 0,979 6,535
24,0 1,513 0,25770 -0,026 0,016 0,959 0,075 0,980 5,873
25,0 1,489 0,27865 -0,026 0,015 0,961 0,066 0,981 5,343
26,0 1,461 0,30438 -0,026 0,015 0,963 0,064 0,981 4,801
27,0 1,430 0,32724 -0,027 0,014 0,966 0,061 0,982 4,369
28,0 1,394 0,35446 -0,027 0,014 0,968 0,056 0,982 3,934
29,0 1,357 0,38350 -0,027 0,013 0,970 0,053 0,983 3,539

So look at the 18% thick airfoil. You can see that:

1% LE has minimal cd=0,01159 if you go to extreme, you can see that 5% LE has on the same airfoil drag cd=0,01164. It is only microscopical difference, because the airfoil drag is mostly caused by the airfoil thickness.

If you look at that thin 12% airfoil, it is 0,00870 versus 0,00943. It is little more, but still not so bad.

But you must look also to lift. That 18 % thick airfoil with 1% LE can make lift cl=1,259 at 13 deg. The drag at that point is cd=0,04829. That is far more than minimal drag at 0 lift. If you go little over 13 deg, you will have problems, because of abrupt separation at LE caused by too sharp LE in comparison to the thick airfoil. That is very bad not only because that abrupt separation, it also makes “bump” on polar what is very bad thing for PA.

If you look at that blunter version, it can make at 20 deg AoA cl=1,529 you can go even further as the top of the lift polar is flat. The drag at 20 deg is cd=0,17387.

If you use that blunt wing and you will use only that small lift from sharp wing, you can see it on 12deg cl=1,291 cd=0,03812.
It means that if you really go to higher wing load and you will really use that higher lift, it means you will have smaller wing for the same model, that smaller wing really makes higher absolute drag even with smaller area. But advantage is smaller area its better resistance to turbulent air … however you pay it by stronger engine. That is why we use flaps and why modern models go to rather smaller dimensions (higher wing load) and stronger engines.

If you want stick at original wing load, you can benefit on blunt wing with even LOWER drag at SOME conditions, but it is not universal rule. The rule is as you mentioned, that blunter LE has more drag – example can be cl=0,6 – the drag is at 1% LE 0,014, while on 5% LE is drag 0,016 at the same lift.

But again, the biggest advangage of blunter LE is stability against separation and its “bump” free polar. It is much more visible on flapped wings.

>From my experience with my, and others Jamison Specials, it
>is possible to get carried away with blunting the leading
>edge, and you get great lift, great round loops, great
>tracking in rounds, but lousy wind penetration, weird things
>happened in square corners (like hitting a brick wall, bam,
>then pull out of corner), but not a stall, and need LOTS of
>power to get through the corners.

Eric, most of the problems you mention sound suspiciously like too much weight for the wing area, all except the wind penetration. If the plane is overweight it will do the first square corner pretty well, but use up most of its energy, and slow down. Then you are left hanging going straight up until you get to the next corner. Everything goes downhill from there. A lighter plane will turn the same corner but not slow down nearly as much. It takes less horsepower to keep it flying through the corners.

The JS(~500 squares)probably needs to weigh in the low 30's max to turn sharply. You can make it fly pretty good at higher weights, but it takes careful work on the handle not to turn it too hard and rub off too much speed.

>Ok, I understand that part Igor, but when (In theory)does
>blunt become TOO blunt?
>
>From my experience with my, and others Jamison Specials, it
>is possible to get carried away with blunting the leading
>edge, and you get great lift, great round loops, great
>tracking in rounds, but lousy wind penetration, weird things
>happened in square corners (like hitting a brick wall, bam,
>then pull out of corner), but not a stall, and need LOTS of
>power to get through the corners.
>Thoughts?
>Eric

Hi Eric

While the difference between blunt and sharp profiles in terms of profile drag is small, you have to consider the difference in induced drag. Induced drag had a squared relationship with lift coefficent (Cl). That means that if your blunt profile will go to twice the Cl of the thin one, induced drag will increase 4 fold near stall. So yes it can be overdone and that will be seen most in a square corner as if putting on the brakes.

If this is what you are observing, you will likely get better results with the same engine by reducing both profile thickness and wing loading. Reduce loading so the wing will not have to work so hard in a corner and reduce thickness so the pilot can't push it to hard either. Reduce wing loading by either reducing weight or increasing wing area.

It seems clear that I should round off the leading edge.
The airfoil that I am using is a Rabe airfoil I found in Stunt News-2nd Bearcat, I think. It does resemble the "Burger" airfoil illustrated in another post.
Very interesting discussion-thanks all.
Bud S

Faith makes happy
This airfoil thing really is a very interesting topic for me, so I’ve read all answers with great interest. Alas I cannot give much help. My own experiences are very contradictory and don’t allow to make conclusions.
For decades I’ve used airfoils very close to what is shown in the sketch as “Maikis 1”. I never have had any problems. When seeing those shapes from Ted Fancher and Victor Salenek, I thought: what is good for top level flyers should be good for me. So I switched to an airfoil shape as shown by “Maikis 2”. Two airplanes were built, one with a slightly thinner shape as shown in the sketch. The model with the thinner airfoil flew quite well, but it had severe stalls in rain ( only in rain ). It just wouldn’t do corners. Nothing helped ( sealing hinge lines, turbulators, CG shift to the rear, enlarged flaps ). The second model didn’t fly at all, it’s now hanging off the ceiling of my friends living room - it just doesn’t fly. And it was the lightest airplane I’ve ever built ( just under 1800 gramm ).
In the sketch there are also the airfoils of Beringer’s, Burger’s, and Windy’s airplanes. Talk about blunt noses and soft corners: there are not many airplanes which fly corners as sharp as Beringer’s Sukhoi! With such an airfoil and the CG only a few Millimeters from the leading edge! That last airfoil in the sketch is that of my latest creation. Apart from being much too heavy again, it flies quite well ( well, for me at least ). It has some kind of mixture between Nobler and Cardinal airfoil.
To sum it up: I’ve lost all faith in small or large leading edge radius concerning their influence on flight characteristics. As has been mentioned before: we have not yet defined what can exactly be called sharp or blunt. And so far we haven’t talked much about or explored curvature aft of leading edge. In my opinion these belong together inseparably. I’m afraid that - as long as we don’t have wind tunnel test results - the old shoe sole outline is not so far off of what Mr. NACA, Eppler, Selig, etc try to recommend.
At least as long our wrist is still the deciding factor .
Regards, claus

Claus,

Considering the profile aft of its high point, that Burger airfoil looks very similar to those used by Al Rabe for thirty plus years with a certain degree of success.

There is a lot of opinion here, most of it little better than guesses. So, for what its worth, here is my guess, or maybe, a bit more than a guess since I seem to be the only stunt flyer to have actually performed airfoil tests, even if they are thirty four years old. Based on those those I became convinced that moderate leading edge radius works better in the wind with little, if any, loss of performance potential in more favorable conditions. I particularly don't like blunt leading edges and fail to find convincing arguments as to their efficacy. Blunt leading edges, as commonly used on contemporary stunt ships, look, to me, like serious over-kill in the desire to obtain lift. Wandering stagnation points????? Really.

>There is a lot of opinion here, most of it little better
>than guesses. So, for what its worth, here is my guess, or
>maybe, a bit more than a guess since I seem to be the only
>stunt flyer to have actually performed airfoil tests, even
>if they are thirty four years old. Based on those those I
>became convinced that moderate leading edge radius works
>better in the wind with little, if any, loss of performance
>potential in more favorable conditions. I particularly
>don't like blunt leading edges and fail to find convincing
>arguments as to their efficacy. Blunt leading edges, as
>commonly used on contemporary stunt ships, look, to me, like
>serious over-kill in the desire to obtain lift. Wandering
>stagnation points????? Really.
>
>Al
Al
Thanks for the input. It is important to me because I did read about you testing your airfoils and saw the picture of how you did it. (A clever device)
To me that is the ultimate test-under actual conditions.
It is obviously the reason I used your airfoil (or close to it) on my plane.
Regards
BudS

>>There is a lot of opinion here, most of it little better
>>than guesses. So, for what its worth, here is my guess, or
>>maybe, a bit more than a guess since I seem to be the only
>>stunt flyer to have actually performed airfoil tests, even
>>if they are thirty four years old. Based on those those I
>>became convinced that moderate leading edge radius works
>>better in the wind with little, if any, loss of performance
>>potential in more favorable conditions. I particularly
>>don't like blunt leading edges and fail to find convincing
>>arguments as to their efficacy. Blunt leading edges, as
>>commonly used on contemporary stunt ships, look, to me, like
>>serious over-kill in the desire to obtain lift. Wandering
>>stagnation points????? Really.
>>
>>Al
>Al
>Thanks for the input. It is important to me because I did
>read about you testing your airfoils and saw the picture of
>how you did it. (A clever device)
>To me that is the ultimate test-under actual conditions.
>It is obviously the reason I used your airfoil (or close to
>it) on my plane.
>Regards
>BudS
PS I'm gonna make my leading edge like the one in the picture.

>There is a lot of opinion here, most of it little better
>than guesses. So, for what its worth, here is my guess, or
>maybe, a bit more than a guess since I seem to be the only
>stunt flyer to have actually performed airfoil tests, even
>if they are thirty four years old. Based on those those I
>became convinced that moderate leading edge radius works
>better in the wind with little, if any, loss of performance
>potential in more favorable conditions.

How about the included angle of the radius? How does that effect it?

> I particularly
>don't like blunt leading edges and fail to find convincing
>arguments as to their efficacy. Blunt leading edges, as
>commonly used on contemporary stunt ships, look, to me, like
>serious over-kill in the desire to obtain lift.

I don't think that's why anyone is using relatively large LE radii. It's certainly not why I do it, since I don't really care much about Clmax, within reason. Any airfoil that is likely to work from a handling characteristic and structural standpoint is going to provide more than enough lift.

I use a relatively large LE radius since I prefer the handling feel they provide.

>Wandering
>stagnation points????? Really.

Yes, really, assuming you mean moving vice wandering. Inevitable consequence of the physics of the situation, no tests necessary.

I guess I only saw the maximum absolute lift for the Sea Fury, etc. airfoils. Do you have pitching moment curves and flow pictures for them as well?

By the way, "sharp" means different things to different people. Last time I checked, the Infinity has maybe 1/16" more LE radius than the Mustang airfoil (although the included angle is much less), and the Trivial Pursuit is almost the same. So you are comparing micro-tiny differences.

When most people say "sharp", they mean really sharp, as in an edge, as in just putting a peice of 1/4" square balsa as the LE wood and just letting the sheeting come to a point. I seem to recall an article in American Aircraft Modeler in about 1973 discussing (quite correctly) the evils of this arrangement, indicating that "sharp" was bad, and that sanding back to the 1/4 LE wood to ensure sufficiently large radius. Or, also quite correctly, "buffeting" would occur.

Your "implicit" argument with Ted is over LE radii of 1/2 vs. 5/8". I contend that's in the noise of other factors.

The internet questions are usually more like 1/32 vs 1/2" and that does make a huge difference - and 1/2" is A LOT better than 1/32, in general.

Hi there Bud,

How are things going down in AL? Right?

This subject has been hashed, smashed and rehashed about 200 times since about 1965. Recommended reading is all the old Round and Round articles from the era by Bill Netzband and other articles by Bob Baron, Bob Gialdini, the Netzband article C/L Aerodynamics made painless?, and everything ever written by Al Rabe, Ted Fancher, Bob Hunt, Paul Walker and Brett Buck. Tom Morris has copies of all the old magazines with these articles. I think he even has an index of them now! Great reading these articles!

>Hi Greg,
>
>So, how does one practically apply this to themselves?

Hi

This in estimation method.

First we need to know a few things about our expected design, some are drawn from experience.

Airspeed or velocity need only be a rough estimate but should be fairly well known as you want a particular lap time and given the engine capacity and line length you use, that is easy to calculate.

V = (Pi * 2 * R) / T

where:
V : Velocity
Pi: 3.142
R : radius (line length)
T : time per lap

Mostly I fly 15s these days on 15.9 m lines and 4 sec lap times
So I'll use that for my example.

V = (3.142 * 2 * 15.9) / 4

V = 25 m/s

Next, we want to find the lift required to turn our model around the tightest turn of a square manouver so we need to know 2 other things.

The mass of the model. A good robust model of the expected sizes would be around 0.57 kg Thats quite heavy for a 15 model and not a stunter so for the purpose of this demo, lets use a mass of .5Kg,
still on the heavy side.

The other thing is to decide the smallest radius turn you want to do.
Lets say 3m for example.

Use :

Fc = m * V^2 / R

where
Fc: Centripetal force in newtons
m : Mass .5kg
R : Radius of turn 3

Fc = .5 * 25^2 / 3

Fc = 104N

Now for some aerodynamics.

L = 1/2 * p * V^2 * S * Cl

Where:

L : Lift force = Fc above
p : density of air = 1.225 kg/m^3 at sea level
V : Velocity = 25m/s
S : wing area (what we want to know)
Cl: Lift coefficient

Cl can be obtained form the NASG polar site for a Renolds value of around 300000, appropriate for the expected speed and size of this model. For E472 Cl max is just above 1.2. This is only a 12% thick
section but has very good Cl max.

We need to rearange the formulae above to find S (wing area)

S = L / (.5 * p * v^2 * Cl)

S = 104 / (.5 * 1.225 * 25^2 * 1.2)

S = 0.226 m^2

Now the mean chord is given by :

Mc = (S / AR)^.5

AR: Aspect ratio assume 6 for the sake of this example.

Mc = (0.226/6)^.5

Mc = 0.194m

Span = AR * Mc

Span = 6 * 0.194

Span = 1.164

So we now have a flapless wing of span 1.165m and average chord of 0.194m

If we wanted to use a flapped wing, we could expect a Clmax around
1.5 times that of the flapless version and end up with a wing area
around 0.151 m^2 a span of 0.953m and mean chord of 0.159m

This is the minimum wing required for the turn you want and you may decide to go for something different for other reasons but at least you know the minimum wing size for the turn radius you want to achieve.

Hope I got all the calcs right, looks OK to me.

Definitelly. If you look at those Eppler and Goettingen data, you do not see they are 40 years old and which of them are calculated and not really measured (even 30 - 40 years ago)

The only difference is, that nowadays you can do that calculation home in 5 minute, while those days it was job for whole night for university computer which is used in night only by crazy guys like we are