Hi,Don't know yet what to do with it, but this formula should give lift force in pounds:(A:AMBIENT DENSITY,Slug/ft3) (A:AIRSPEED TRUE,feet/second) (A:AIRSPEED TRUE,feet/second) 3050 (wing area in sq ft) (A:INCIDENCE ALPHA,degrees) * * * * 2 / (>L:LIFT,pounds) Is this correct?(Values sometimes over 2.000.000 pound)Jan"Beatus ille qui procul negotiis..."

I believe you need acceleration, not velocity... but I could be wrong.

Hi,Found:L = (1/2) d v2 s CLL = Lift, which must equal the airplane's weight in pounds.d = density of the air. v = velocity of an aircraft relative to the airmass expressed in feet per second. s = the wing area of an aircraft in square feet. CL = Coefficient of lift , which is determined by the type of airfoil and angle of attack. Straight and level at Fl340:Gross Weight 297.000 pounds.The Formula gives 1.700.000 pounds.( didn't take into account the shape of the airfoil )What to say about the big difference,?Jan"Beatus ille qui procul negotiis..."

The formula is correct. But you don't give the values for density, speed, wing area, and lift coefficient you are using so it's not possible to comment on the difference. Anyway lift only equals weight in straight and level flight. Based on your values the aircraft could be pulling about 6g.Ignoring non-linearities, Cl can be expressed asCl = Cl0 + ClAlpha * Alpha whereCl0 = lift coefficient at zero incidenceClAlpha = variation of lift coefficent with incidence (lift curve slope)Alpha = incidenceAccording to simplified aerodynamic theory, the lift curve slope is 2PI/radian or about 0.11/degree for an infinite span conventional wing at low speed. For finite span wings the same simplified theory suggests that the lift curve slope should be factored down by A/(2+A) where A is the aspect ratio of the wing whereA = Span squared/Area.In reality the actual lift curve slope tends to be less than this, but these formula but give good enough values to check if estimates are of the right order.For straight and level flight (weight = lift) then you can calculate Cl for given density, speed, wing area.

Hi,The values are in the first post.Jan"Beatus ille qui procul negotiis..."

Well,Not sure it is 100 % ok, but finding Cl in straight and level flight by:(A:TOTAL WEIGHT,pounds) (A:AMBIENT DENSITY,Slug/ft3) (A:AIRSPEED TRUE,feet/second) (A:AIRSPEED TRUE,feet/second) 3050 (wing area in sq ft, aircraft.cfg) * * * / 2 * (>L:LIFTcoefficient,number) (about 0.44) Gives a reasonable result for lift (Weight = Lift): (A:AMBIENT DENSITY,Slug/ft3) (A:AIRSPEED TRUE,feet/second) (A:AIRSPEED TRUE,feet/second) 3050 (L:LIFTcoefficient,number) * * * * 2 / (>L:LIFT,pounds) Jan"Beatus ille qui procul negotiis..."

Hi,Another unuseful addition after some more investigation: (A:TOTAL WEIGHT,pounds) (A:AMBIENT DENSITY,Slug/ft3) (A:AIRSPEED TRUE,feet/second) (A:AIRSPEED TRUE,feet/second) 3050 * * * / 2 * pi 2 (A:INCIDENCE ALPHA,radians) * * - (>L:AIRFOIL,number) (A:AMBIENT DENSITY,Slug/ft3) (A:AIRSPEED TRUE,feet/second) (A:AIRSPEED TRUE,feet/second) 3050 (L:AIRFOIL,number) pi 2 (A:INCIDENCE ALPHA,radians) * * + * * * * 2 / (>L:LIFT,pounds) Jan"Beatus ille qui procul negotiis..."

You didn't give values for speed, or density, in your first post. It didn't even mention lift coefficient which was why you results were wrong,. The only value given was wing area.

Hi,All units are there, not the values, but erronously I assumed (A:INCIDENCE ALPHA,degrees) as Cl.3050 was from aircraft.cfg.Jan"Beatus ille qui procul negotiis..."

Units are not much use without values when trying to check the result.Whhich entry in the .cfg file with the value of 3050 are you referring to?

Hi,The results of the units are of course visible in a gauge.Like:weight: 380.000 lbsspeed: 250 ft/secdensity: 0.001 slug/ft^3aoa: 0.1 radiansIn the acft.cfg Wing Area: in sq ft (3050)Also from Boeing data. Btw. The formula gives now exact the same results as AFSD of Herve Sors, so it seems to be not so very wrong.Jan"Beatus ille qui procul negotiis..."

But, of course, we don't have any gauges visible in the forum so how could we have helped you?The density is 0.001 slugs/ft3 only at about 26,600ft. It's 0.00237 slugs/ft3 at sea-level for the International Standard Atmosphere.

For those of us who don't write RPN... this is rather difficult to follow.(A:TOTAL WEIGHT,pounds) (A:AMBIENT DENSITY,Slug/ft3) (A:AIRSPEED TRUE,feet/second) (A:AIRSPEED TRUE,feet/second) 3050 * * * / 2 * pi 2 (A:INCIDENCE ALPHA,radians) * * - (>L:AIRFOIL,number) This is: AIRFOIL = ((TOTAL WEIGHT * 2)/(AMBIENT DENSITY * AIRSPEED TRUE * AIRSPEED TRUE * 3050))-(pi*2*INCIDENCE ALPHA)(A:AMBIENT DENSITY,Slug/ft3) (A:AIRSPEED TRUE,feet/second) (A:AIRSPEED TRUE,feet/second) 3050 (L:AIRFOIL,number) pi 2 (A:INCIDENCE ALPHA,radians) * * + * * * * 2 / (>L:LIFT,pounds) This is: LIFT = ((((((((INCIDENCE ALPHA * 2)*pi)+AIRFOIL)*3050)*AIRSPEED TRUE)*AIRSPEED TRUE)*AMBIENT DENSITY)/2)Is this correct?

Hi,Cl=2*weight/(density*speed^2*surface)Cl=Cairfoil+2*pi*aoa,radianstranslated:Cairfoil=2*(A:TOTAL WEIGHT,pounds)/((A:AMBIENT DENSITY,Slug/ft3)*(A:AIRSPEED TRUE,feet/second)*(A:AIRSPEED TRUE,feet/second)*3050)-pi*2*(A:INCIDENCE ALPHA,radians)So:Lift=1/2*density*speed^2*Cl*surfaceLift=1/2*density*speed^2*(Cairfoil+2*pi*aoa,radians)*surfacetranslated:Lift=1/2*(A:AMBIENT DENSITY,Slug/ft3)*(A:AIRSPEED TRUE,feet/second)*(A:AIRSPEED TRUE,feet/second)*((L:AIRFOIL,number) + pi*2*(A:INCIDENCE ALPHA,radians))*3050Was this a joke or are you serious?..Jan"Beatus ille qui procul negotiis..."

Quite serious.Despite what ACES thinks... most of the civilized world does not do math via reverse polish notation (RPN), as the XML stack method is commonly refered to.