Degraded Su-27 aerodynamic lift

[DummyCatz]

4 hours ago, GT400 said:

Lastly,  I get it that in the TsAGI report it lists 18920kg as design flight mass at 50% fuel reserve,  but does any one know the exact amount of fuel that is and the Su-27S operating weight?  (easier to write it as operating weight?+fuel?= 18920kg)

 

On 3/1/2018 at 1:11 AM, Vatikus said:

Manual says there are 3 options of refueling the plane.. full (forward, center, back & wings), basic (center & wings), partial (center).

So the graphs for example state.. 50% remaining of normal fueling, which could been interpreted as 50% of basic option.

 

tank forward - 3180 kg

tank center - 4160 kg

tank wings - 1060 kg

tank back - 1000 kg

It's actually 50% of normal fueling, which is (4160 + 1060) / 2.

[draconus]

8 hours ago, GT400 said:

what does 1.6, 1.08, 1.4 numbers represent for Su-27, F-15, F-16?

It's permissible lift coefficient for full AB and 50% of internal fuel load.

Edited January 4 by draconus

[Yo-Yo]

Full gross weight for calculation in this document can be directly found from the 

G/S = 305 kg/m2.  S is known - 62 m2. So, you can find the necessary fuel weight to have necessary GW including mentioned weapons.

[MA_VMF]

11 часов назад, GT400 сказал:

I have a few questions about Su-27, and my apologies in that I cannot understand Cyrillic.  But in TsAGI Su-27S paper, what does 1.6, 1.08, 1.4 numbers represent for Su-27, F-15, F-16? 

 

Also, so once Su-27S and SK are are equal or less than 19,000kg they can both pull 9G?  171000/9=19000kg and as long as it's less than mach .85?

Lastly,  I get it that in the TsAGI report it lists 18920kg as design flight mass at 50% fuel reserve,  but does any one know the exact amount of fuel that is and the Su-27S operating weight?  (easier to write it as operating weight?+fuel?= 18920kg)

 

Thank you.

 

This is the data for the T-10 prototype. For the Su-27S, a total weight of 20000kg is assumed. 2хR-27R 2хR-73 fuel = 2600kg. Судоп=1.85

2 часа назад, draconus сказал:

It's permissible lift coefficient for full AB and 50% of internal fuel load.

the permissible coefficient does not depend on the mass

Edited January 4 by MA_VMF

[Xhonas]

On 12/30/2024 at 10:08 PM, Yo-Yo said:

 

1. Given that reliable data on the hysteresis loop shape for the lift coefficient (Cy) is available for only a handful of aircraft, do you believe applying unsteady aerodynamics exclusively to these aircraft is fair to others? Alternatively, would it be valid to use data obtained, for example, for the Su-27 to model the F-18?

Yes please ! Given that DCS tries to be a super accurate flight simulator i don't think "fairness" should be a question. If you have the data to model the Su-27 more accurately, i think it is valid to use it.

[GT400]

So how is permissible clmax of 1.6 different than clmax of 1.85?  When the Su-27sk flight manual book 1 says 171,000kg is the max allowed static force put on the wing under mach .85,  which clmax are they using to get that figure?   Also in flight manual book 2 (again apologies in that I cannot read Cyrillic)  in the first column it lists at mach .5 a clmax of 1.85 at 24 AOA and then under that in a second column also at  mach .5 a clmax of 1.61 at  20AOA ?  Is this the permissible clmax here?  

 

If using permissible clmax 1.6 in equation I get:  

 (I figured for sea level and not 200 meters as TsAGI em chart) 

lift = 1/2*v^2*reference wing area*air density *cl

lift = 1/2* 170m/s^2*62*1.225*1.6

lift= 1755964/9.81=178,997

 

and if using clmax of 1.85 in equation I get 206,965

neither of those numbers at mach .5 (170m/s at sl) get you to the 171,000kg.

What am I doing wrong here? 

Edited January 9 by GT400

[Yo-Yo]

On 1/9/2025 at 7:54 PM, GT400 said:

So how is permissible clmax of 1.6 different than clmax of 1.85?  When the Su-27sk flight manual book 1 says 171,000kg is the max allowed static force put on the wing under mach .85,  which clmax are they using to get that figure?   Also in flight manual book 2 (again apologies in that I cannot read Cyrillic)  in the first column it lists at mach .5 a clmax of 1.85 at 24 AOA and then under that in a second column also at  mach .5 a clmax of 1.61 at  20AOA ?  Is this the permissible clmax here?  

 

If using permissible clmax 1.6 in equation I get:  

 (I figured for sea level and not 200 meters as TsAGI em chart) 

lift = 1/2*v^2*reference wing area*air density *cl

lift = 1/2* 170m/s^2*62*1.225*1.6

lift= 1755964/9.81=178,997

 

and if using clmax of 1.85 in equation I get 206,965

neither of those numbers at mach .5 (170m/s at sl) get you to the 171,000kg.

What am I doing wrong here? 

And what about 5-7 km altitude?
You mixed two limitations: the table you posted limits AoA and, thus, CL to prevent controllability issues. 171 kG limits wing loads.  As Mach number and air density are different at different altitudes different limits will play a role.

[GT400]

22 hours ago, Yo-Yo said:

And what about 5-7 km altitude?
You mixed two limitations: the table you posted limits AoA and, thus, CL to prevent controllability issues. 171 kG limits wing loads.  As Mach number and air density are different at different altitudes different limits will play a role.

I think I get it now.  171,000kg is total limit for wing loads of Su-27S/Sk.  I also forgot to remember as you increase altitude, air density decreases.  Thank you Yo-Yo.

[DummyCatz]

On 1/10/2025 at 1:54 AM, GT400 said:

If using permissible clmax 1.6 in equation I get:  

 (I figured for sea level and not 200 meters as TsAGI em chart) 

lift = 1/2*v^2*reference wing area*air density *cl

lift = 1/2* 170m/s^2*62*1.225*1.6

lift= 1755964/9.81=178,997

 

One thing to keep in mind though, is that the 171,000 kg limit is a force being applied perpendicular to the wing (body-fixed coordinate system, or body axes frame of reference), while the lift is perpendicular to the velocity vector (velocity coordinate system, or wind axes frame of reference).

That means, the force being applied perpendicular to the wing = total lift * cos(AOA) + total drag * sin(AOA).

Ny is also in the direction perpendicular to the wing, and is not aligned with the lift. So Ny * weight = total lift * cos(AOA) + total drag * sin(AOA), which has a 171,000 kg limit.