Jetflyer From Netherlands, joined Aug 2009, 0 posts, RR: 0 Posted (10 years 1 month 2 weeks 4 days 15 hours ago) and read 13068 times:

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:

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?

Jetflyer From Netherlands, joined Aug 2009, 0 posts, RR: 0
Reply 2, posted (10 years 1 month 2 weeks 4 days 4 hours ago) and read 13012 times:

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!

David L From United Kingdom, joined May 1999, 9644 posts, RR: 42
Reply 3, posted (10 years 1 month 2 weeks 4 days 3 hours ago) and read 13002 times:

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 climb angle or the angle of attack (or something else) is 5.33^{o}? 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?

Vikkyvik From United States of America, joined Jul 2003, 12081 posts, RR: 24
Reply 4, posted (10 years 1 month 2 weeks 4 days 3 hours ago) and read 12997 times:

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

I'm watching Jeopardy. The category is worst Madonna songs. "This one from 1987 is terrible".

Jetflyer From Netherlands, joined Aug 2009, 0 posts, RR: 0
Reply 6, posted (10 years 1 month 2 weeks 4 days 1 hour ago) and read 12986 times:

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.

Pihero From France, joined Jan 2005, 5145 posts, RR: 78
Reply 9, posted (10 years 1 month 2 weeks 3 days 17 hours ago) and read 12919 times:

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.

Jetlagged From United Kingdom, joined Jan 2005, 2620 posts, RR: 25
Reply 10, posted (10 years 1 month 2 weeks 3 days 7 hours ago) and read 12879 times:

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.

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

Doktor From Germany, joined Mar 2006, 29 posts, RR: 0
Reply 11, posted (10 years 1 month 2 weeks 3 days 2 hours ago) and read 12853 times:

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

Liedetectors From United States of America, joined Jul 2005, 360 posts, RR: 0
Reply 14, posted (10 years 1 month 2 weeks 2 days 16 hours ago) and read 12793 times:

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.