Hi, I have a project underway to calculate the rates of climb of airliners based on various factual information about the airfoil.

This includes calculating the lift quantity and induced drag in pounds. Thew following data will go into a spreadsheet, you see, and these are the requirements:

The input example below is of a generic aircraft.

INPUT

AIRFOIL

Wing Area: 980sq/ft

Mean Chord: 11.06ft

Wing Span: 88.6ft

Aspect Ratio: 8.01

Camber - %C: 4.05

Thickness - %crd: 9.25

OUTPUT REQUIRED

Total lift at a given airspeed (lbs) at sea level

Induced drag at a given airspeed (lbs) at sea level

AIRCRAFT INPUT

Gross takeoff weight: 107,456lbs

Gross Thrust: 25,120lbs

Parasite Drag

OUTPUT desired:

Some form of calculation that gives me the rate of climb based on the excess thrust/lift taking into account drag levels and aircraft weight.

Is this asking too much maybe? I know there are many "experts" in this area on the board, so I'd be happy to know a way of calculating these parameters to give me some intelligible data on Microsoft Excel.

That aircraft above is an example, with most of the data required to calculate this stuff, can anyone show me the formulas to create the lift/drag quantity and work out the rate of climb?

This includes calculating the lift quantity and induced drag in pounds. Thew following data will go into a spreadsheet, you see, and these are the requirements:

The input example below is of a generic aircraft.

INPUT

AIRFOIL

Wing Area: 980sq/ft

Mean Chord: 11.06ft

Wing Span: 88.6ft

Aspect Ratio: 8.01

Camber - %C: 4.05

Thickness - %crd: 9.25

OUTPUT REQUIRED

Total lift at a given airspeed (lbs) at sea level

Induced drag at a given airspeed (lbs) at sea level

AIRCRAFT INPUT

Gross takeoff weight: 107,456lbs

Gross Thrust: 25,120lbs

Parasite Drag

OUTPUT desired:

Some form of calculation that gives me the rate of climb based on the excess thrust/lift taking into account drag levels and aircraft weight.

Is this asking too much maybe? I know there are many "experts" in this area on the board, so I'd be happy to know a way of calculating these parameters to give me some intelligible data on Microsoft Excel.

That aircraft above is an example, with most of the data required to calculate this stuff, can anyone show me the formulas to create the lift/drag quantity and work out the rate of climb?

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Thanks, I've been using that site as reference, but if only calculates the climb angle, for my aircraft I calculated that with 10,000lbs of excess thrust and a weight of 107456lbs the climb angle equals 5.33 degrees approximately. However, that doesn't tell me much, because it could be anything from 1fpm to 5,000fpm.

I was told once that in order to climb the amount of excess thrust and lift in pounds put together had to exceed the aircraft's weight. Therefore I need to use those properties of the wing above to work out the rate it would climb at a 5.33 degree climb angle, based on a given airspeed and and a certain quota of lift based on aspect ratio, wing loading etc... There must be a way!

I was told once that in order to climb the amount of excess thrust and lift in pounds put together had to exceed the aircraft's weight. Therefore I need to use those properties of the wing above to work out the rate it would climb at a 5.33 degree climb angle, based on a given airspeed and and a certain quota of lift based on aspect ratio, wing loading etc... There must be a way!

Quoting Jetflyer (Reply 2):Therefore I need to use those properties of the wing above to work out the rate it would climb at a 5.33 degree climb angle, based on a given airspeed |

I'm not sure I follow. Are you saying the

Of course, as usual, I may be missing something.

I think what you're missing is the property of the airfoil. Yes, chord, thickness, etc. are the physical properties, but to know the lift, you need to know the plot of lift coefficient vs. angle of attack. Once the lift coefficient is determined, you can calculate the total lift, and therefore, the induced drag. I don't believe you can calculate the lift just with the figures you gave, because you don't know what the incidence angle of the wing is in the first place.

Unfortunately, many modern commercial airplanes tend not to use one airfoil over the whole wingspan. The airfoil will change as you go from fuselage to wingtip.

God, it's been about 3 years since I went over this stuff, but if you need the formula for the lift or induced drag (both for a given lift coefficient), then I can supply those.

~Vik

Unfortunately, many modern commercial airplanes tend not to use one airfoil over the whole wingspan. The airfoil will change as you go from fuselage to wingtip.

God, it's been about 3 years since I went over this stuff, but if you need the formula for the lift or induced drag (both for a given lift coefficient), then I can supply those.

~Vik

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Yes, it's the *"5.33*^{o} climb angle" that's confusing me. What exactly is *"climb angle"*?

Quoting David L (Reply 3):
I'm not sure I follow. Are you saying the climb angle or the angle of attack (or something else) is 5.33o? If it's the angle at which the aircraft climbs and you have a given speed, wouldn't simple trigonometry give you the rate of climb? |

No, you add the angle of attack to the angle of climb (5.33*) to get the attitude, but the point is that the angle of attack depends on the airspeed, which in turn affects the drag quantity which allowed me to calculate excess thrust in the first place.

The angle of climb is trigonometry it is implie that if the aircraft had excess thrust to match its weight it would climb straight up, but because its excess thrust is only a fraction of its weight, it climbs at a certain angle depending on that amount.

However, it doesn't tell me the lift produced by the airfoil.

The calculation of the climb angle was sin-1(10000/107456) = 5.339748767 degrees.

How do I go about determining the lift co-efficient? I used the NASA foil simulator to create the wing to those specifications, and it shows a lift calculation in pounds and a co-efficient. But I'm wondering how it calculates them.

[Edited 2006-03-22 14:42:16]

Well for the standard 3º climb/descent path it's groundspeed x5

"They have lady pilots......... they're not that good, but they have 'em"

Jetflyer,

I don't think anybody on this site could help you if he/she wasn't a manufacturer's performance engineer.

Vikkyvik started with a glimpse of some of the parameters you would need.

Basically, a polar of the complete aircraft would be desirable and in all my years I haven't been able to procure one. The wing polar in itself is not enough.

On the other hand, most posters here could provide you with the basic aerodynamic equations and the only way you could fill up that spreadsheet of yours would be to plot a lot of climb performance situations from an AOM, but it will only give you fleet average, not the accurate data you need.

The same remark is also valid for engine performance : For instance, what is the thrust value of engine X at OAT = 35°c ? I have no idea, I only know that I could expect N1=98%....and so on.

Sorry couldn't help you more

I don't think anybody on this site could help you if he/she wasn't a manufacturer's performance engineer.

Vikkyvik started with a glimpse of some of the parameters you would need.

Basically, a polar of the complete aircraft would be desirable and in all my years I haven't been able to procure one. The wing polar in itself is not enough.

On the other hand, most posters here could provide you with the basic aerodynamic equations and the only way you could fill up that spreadsheet of yours would be to plot a lot of climb performance situations from an AOM, but it will only give you fleet average, not the accurate data you need.

The same remark is also valid for engine performance : For instance, what is the thrust value of engine X at OAT = 35°c ? I have no idea, I only know that I could expect N1=98%....and so on.

Sorry couldn't help you more

Contrail designer

If you aren't accelerating, and you have a climb angle gamma, then

Lift = Weight / cosine (gamma)

Not exact, but close enough for your purposes I would have thought.

Lift = Weight / cosine (gamma)

Not exact, but close enough for your purposes I would have thought.

The glass isn't half empty, or half full, it's twice as big as it needs to be.

Hi Jetflyer!

If you are really interessted I could send you an pdf file from my lecture notes of flight mechanics. You can definitly find your answers, but they are not that easy to understand...

Dok

If you are really interessted I could send you an pdf file from my lecture notes of flight mechanics. You can definitly find your answers, but they are not that easy to understand...

Dok

Hi there, that would be interesting, I'll send you my E-mail address, or actually, it should be on my profile.

Jetflyer,

I tried to send you a mail, but I couldn attach anything. You got to send an instant message or mail me your address!

I tried to send you a mail, but I couldn attach anything. You got to send an instant message or mail me your address!

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Hey Jetflyer, Any college level flight mechanics text book should be able to help. I would look at mine and help you out, but they are all at work.

If it was said by us, then it must be true.