At sea level aircraft fly proportionally too fast to make this relevant, as Cl increase is linear, but Cd increase is more exponential: I couldn’t find 64A204 but lets take 64A215 airfoil, its similar to the F-16’s airfoil, but thicker and does not have LE flaps:
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(Though this graph made by to CFD modelling there are inaccuracies; there should be minimal CD increase between 0-5 deg AOA following the parabolic behaviour of the graph)
Lets take this airfoil as the F-16’s airfoil;
For an F-16 blk50 (27,87m2 wing area) flying with 50% fuel + 4 missiles (12690 kg) at sea level (air density = 1,2)
Level flight at M0,5 reqiures 0,23 CL. this (coindicentally) happens at 0 AOA, and with minimal drag coefficient. As drag is also multipled with wing area, having smaller wings is beneficial and will create less drag. This will help in level flight accelerations.
Turning at 9G at M0,85 requires: 0,72 CL. This happens at 5 deg AOA; still at very low Cd. Low Cd multipled with low wing Area results in low drag, which explains the impressive sustained turn performance of the F-16. If F-16 had bigger wings, it would only increase its drag, and lower its sustained turn performance.
However when same aircraft reaches 30k feet (air density = 0,35)
Level flight at M0,5 requires 0,98 Cl at 10 deg angle of attack; this corresponds to 0,02 Cd; Now the Wing Area becomes important: IF F-16 aircraft had 33% larger wings; it would have required ~0,75 cl and had 0,01 Cd: This 33% increase in wing area and 50% decrease in Cd would actually resulted in 33% LESS draggy aircraft (despite being physically larger), directly improving acceleration.
Turning at 4Gs at 30k feet would require 1,27 Cl at 15 deg AOA resulting in 0,06 Cd (note that this is not entirely accurate for the Real Life F-16 has LE flaps which would decrease the Cd). If F-16 had twice the wing area; it would have had only 0,01 Cd; and be 300% less draggy;
Wing Loading also tells about how much available lift aircraft can produce; At M0,85 Clmax= 1,5; F-16 can pull 5Gs at best, resulting in 10,3 deg/s instantenious turn rate at 30k feet. Greater the wing area, greater the instantenious turn rates.
This all comes to optimization and engineering choices: Current small wings of F-16, are optimal at Sea level, but inferior at 30k feet. Slightly enlarging wings would improve both maneuvering and acceleration performance at high altitude, at the cost of low altitude performance. Enlarging wings too much would horribly worsen the low altitude turn, acceleration and climb performance; worsen high altitude acceleration, but improve high altitude turn performance. And I am talking purely about aerodynamics; an increase in deadweight would also decreased payload and further increased the drag etc etc.
With ever increasing payloads; smaller winged and lighter aircraft have more drastic increases in wing loading;
When we fuel an F-16 and Su-27 at 50% fuel load; they will have 383 and 338 kg/m2 wing loading respectively. However if we are to add 3 tons of fuel/payload; they will become; 490 and 387 kg/m2 respecively. Remembering the formula above; this means, F-16 will require 28% greater lift coefficient to achieve same turn rate; Su-27 will only require 14%; Considering Cd increases exponentially, that means there will be huge performance drop for F-16 while both accelerating and turning; but -aerodynamically speaking- Su-27 wont be affected at all while accelerating; and only a slight performance drop while turning. That is if we are speaking about high altitude.
In simpler terms smaller the aircraft gets, greater the Cl requirement will become with equal payload. Wings stay same; so there will be greater AOA requirement; and lower L/D as in this graph.
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As to add to the discussion; let me try to explain how wing loading, T/W, and other factors can be used to *roughly* estimate and compere maneuverability: look at F-15; F-15 uses NACA 64A203 airfoil, very similar to the F-16’s 64A204.
At 50% fuel clean, F-16 has wing loading of 383 kg/m2, whereas F-15 has 285 kg/m2. Wing’s Low AOA behaviour is the same, but F-16 has Clmax of ~1.6 at M0,5 maneuvering conditions F-15 has 1,1; due to lack of LE flaps it stalls earlier.
Remembering this formula: Wing loading * aircraft G load = 1/2 * air density * Lift coefficent * airspeed^2
45% less Clmax + 26% less wing loading means F-15 will have less available G and less instantenious turn rate by 8% at sea level. Obviously this is no coincidence that FM data shows at M0,5 F-16 can pull 24,2 deg/s and F-15 22 deg/s, 9% difference.
As wing behaviour is similar due to similar airfoils; 26% less wing loading = 26% less necessary Cl at 9G; We can assume (albeit a little inaccurrately) that means 26% decrease in Cd. F-15 has 202% greater wing area (read = drag) of F-16, it has 63% greater static thrust. All relations are linear; so multiplying all relations will tell us; F-15 will have 91% Sustained turn performance of F-16; Real numbers are 20,5 vs 21,5 respectively; 5% drop in performance.
As Su-27 uses different airfoil such estimate about STR is not possible because we simply dont know how Cd changes; We can however estimate Instantenious turn rate:
Compared to F-16; it has 15% better Clmax; and 11% less wing loading. Knowing this; means we can estimate 24,2*1.11*1.15 = 30,8 deg/s ITR at M0,5. But that would result in greater than 9Gs; Truth is Su-27 achieves its highest ITR at M0,48 (30,2 deg/s); That is still an accurate estimate.
Here is the STR graphs of 4th gen fighters compared to F-4E at 30k feet when clean 50% fuelled:
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So 64A203 is only 3 % thick where as 64A204 is 4% and 64A215 is 15% thick..entirely different wing with entirely different flight caracteristic and drag etc. I think there is a reason why can’t find them…they are classified.
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