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Guest Ron Freimuth

xml lift formula

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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..."

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

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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..."

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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.

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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..."

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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..."

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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.

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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..."

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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?

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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..."

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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.

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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?

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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..."

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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.

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>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.I agree, Ed. However, I believe RPN is the only type of scripting that can be parsed from an FS XML file with an acceptable performance. In fact, the way it mainly works is such #### simple (so to speak) that I can't say anything but kudos to ACE's team developers!Tom

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>Hi,>>The results of the units are of course visible in a gauge.>Like:>>weight: 380.000 lbs>speed: 250 ft/sec>density: 0.001 slug/ft^3>aoa: 0.1 radians>In 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 One needs to use AoAe, not AoA. Since CL = 0 at AoAe=0. One can find this in TBL 404. REC 1101, 'AoA at CL = 0' should agree. AoAe is then calculated as AoAe + alpha. One can multiply Weight by 'G' to get actual wing lift. That worked pretty good in my XML Test Gauges. Lift slope should be gotten from TBL 404, it is [delta CL]/[delta alpha]. Another complication is that lift slope varies with M. Set in TBL 401. Since I calculated Lift from G*W, I didn't have a problem with some of the other details. In any case, AFSD is very useful for checking that one's calculations are working correctly. AFSD3, for FS9 and FSX no longer requires the standard FS WX, any variation of temperature, pressure, etc. gives the correct values for everything. Ron

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Ok, something odd happening.I calculate lift correctly. I have correct values for Cl, Cdp. All confirmed by AFSD.I can not get the correct value for Cdi. I am calculating it with the following formula:Cl^2/(Pi*e*AR)It's off, not horrendously but off never-the-less. I have confirmed I have the same values for e and AR that are shown in AFSD. So why am I not coming up with the same value for Cdi that AFSD is calculating?

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Ed,At this time, I'm affraid nobody knows exactly how FS calculates induced drag. The classical formula you quote doesn't work correctly in many situations (it is the vortex drag associated with the untwisted wing). By now, AFSD Cdi is the residual drag obtained by substracting from total CD, all other drag components (parasite, compressibility drag,flaps, gear, spoilers,etc..). I named it CDi since, after numerous tests, we are rather on the safe side for all other drag components and the remainer "should" be the viscous and inviscid drag parts. This is still open to experimentation but clearly you will not fit the value with the classical formulaHerv

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Well, I'm even more confused at this point.Cdp is the summation of all the items you just listed. It is clearly less than Cd. So, where are you obtaining Cd from?I need to be able to accurately calculate drag, that's my ultimate goal.

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Ed,Total drag and Cd are obtained from 2DOF aerodynamical equations (acting forces along true air speed flight path), dynamic pressure and wing reference area. No assumption here about the participating drag componentsIf your aim is to calculate total drag and Cd during steady level flight only, you will need to assess simultaneously total thrust (and thrust angle in FSX), AoA, dynamic pressure and acceleration along the TAS axe and use the classical equations of motion for calculating total drag and therefore Cd.Calculation during climb and descent will request some additional data (slope angle and actual gross weight)Herv

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Hi,I feel like an intruder between the pro's, but i think i calculated Cdi (controlled by AFSD) with:(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 / 100000 / 100 * (>L:LIFT,pounds) (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) (L:Airfoil,number) 2 pi (A:INCIDENCE ALPHA,radians) * * + (Cl)(L:Airfoil,number) 2 pi (A:INCIDENCE ALPHA,radians) * * + * (Cl) pi 0.69 * (eff.fact)156.083 (span, ft)156.083 * (span, sq ft)3050 (wing area, sq ft) / * / (>L:Cdi,number) Sorry for RPN, but "Translation" above.Jan"Beatus ille qui procul negotiis..."

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Jan et al,May I intrude as well? :-)I am eager to know the reason you need such "technical" values like Lift, Drag, Airforil, etc within a flight...Do you use them in FMC calcs maybe?Tom

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Tom,Nothing of the kind?, of course a big bit!I am always busy on long haul flights and that are the only flights i like because:1.Listening to or playing Mozart.2.Looking at the Mythbusters.3.Drinking Coffee.4.Talking to the wife.5.On call for duty, work.6.And....wrestling with aircraft stuff to develop a reasonable Thrust System.I still didn't succeed to set N1 in take off, climb, cruise and descent.I tried all kind of formulas without succes.And testing those, lift, drag, Cl etc. came along.And because of failures the question was born, see upstairs.Do you guys have good thrust, performance etc. stuff, so you can set thrust and derated thrust etc.?????Jan"Beatus ille qui procul negotiis..."

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I'll continue.> One needs to use AoAe, not AoA. Since CL = 0 at AoAe=0. >One can find this in TBL 404. REC 1101, 'AoA at CL = 0' should>agree. >> AoAe is then calculated as AoAe + alpha. > .......... That should be AoAe = alpha - AoAo. Where AoAo is value where CL=0 (typ -0.033 rad). CL_alpha is slope of TBL 403 at low AoA. Typ 5.0 Cdi = [AoAe * Kf*CL_alpha]^2 * Cid Cid = 1/[Pi*AR*e; AR = sqrt(bw)/Sw Kf is a fudge factor, typically 0.985, ideally 1.00. I adjust it to get correct total drag at low speeds were Cdi predominates. One can also adjust Kf so Cdi agrees with AFSD. The above calculates Cdi very well, and automatically accounts for TBL 401, which modifies Oswald Efficiency as Mach changes. Add zero lift drag(s) to get total CD. Physical drag then equals CD * S * q. Dynamic pressure, q, should be calculated from V based on Mach, not 'True Airspeed'. FS9 and probably FSX give the wrong value for TAS, further, I had other problems using it to get q = 1/2 rho V^2. Ron

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