Aces High Bulletin Board
General Forums => Aircraft and Vehicles => Topic started by: Stoney on October 17, 2008, 11:21:21 AM
-
Ok,
For my project, I went looking for the design lift coefficients for the two extremes of the flight envelope:
1) Straight flight at maximum speed; approximately 300 mph.
2) Sustained 3 G turn decelerating to approx. 200 mph.
Aircraft weight: 725 lbs
Wing Area: 66 ft^2
Dynamic pressure calculated for standard conditions at 6,000 ft.
Using the lift equation: L = Cl X q X V
Where L = lift (weight), Cl = lift coefficient, q = dynamic pressure, and V = wing area
I came up with .0571 for condition #1 and .3856 for condition #2. Is there anything I'm not considering for the turn condition, if 3G is sustained, and the speed doesn't drop below 200 mph?
-
V is usually velocity while A is usually area, although I'm way out of date with modern methods that formula as written is confusing, at least to me.
-
Could be my text book then. He has V (capital) listed as area and A (capital) is typically Aspect ratio in his formulas.
q = 1/2 p v^2 -or- q = 1/2 X p X v^2 ...I can't find a dot or better symbol for multiplication.
Where p = pressure and v (lower case) = velocity (in fps).
-
I thought the formula for Lift was
Lift = CL(1/2 pV^2s)
Where
CL = lift coefficient
P = air density (not really a 'p', actually a greek letter 'rho')
V^2 = velocity squared
s = wing area.
Basically it's rho times the square of the speed times wing area, collectively divided by two and then times the CL.
-
I'm solving for dynamic pressure, then putting it into a simplified equation:
1/2 p v^2 = dynamic pressure, or q
My lift equation is L = Cl q S where L = lift, Cl = lift coefficient, and S = wing area.
*Sorry* I typo'd on the V = area thing earlier and improperly identified p as pressure vs. density.
Regardless, we're getting away from my original question (now that I have my variables properly labled :o)
-
P = 0.5 # density rho
m = 725.0 # mass lbs
A = 66.0 # wing area ft2
V1 = 300.0 # V condition 1 mph
V2 = 200.0 # V condition 2 mph
g2 = 3.0 # g force in condition 2
# Cl in condition 1
Cl1 = m / ( 0.5 * V1**2 * A * P)
# Cl in condition 2
Cl2 = (m * g2) / ( 0.5 * V2**2 * A * P)
-
Ah, ok Stoney, must've missed that last nite. FSO took a lot out of me ;)
-
Can't edit the original post.
So, the density at 6000ft is something like a 0.0499 lbs/ft3.
-
Can't remember the actual value. I use a dynamic pressure calculator on the Stanford University AeroE website that removes all the decimal-place math. It also computes mach number, Reynolds number, etc. at the same time and is very useful for some shortcuts in the formulas.
-
Stoney
Sorry, I can't follow you...
What VALUE your are actually after?
-
Well, my original question, which apparently I failed to adequately articulate, was whether or not there were other factors I forgot to consider when determining the lift coefficients for each condition of flight.
-
The Cl is basically up to the angle of attack:
http://en.wikipedia.org/wiki/Coefficient_of_lift
-
It depends what you want; if you are looking for speed, concentrate to the Cl range at max speeds and which is reachable at high speed turns. Only you know what you exactly want
However, I'm wondering a bit your aproach because apparently you have allready selected the wing profile. Normal way is to determine general layout of the plane and assumed Cl range first and choose profile and fine tune the layout based on that.
-
Ok,
For my project, I went looking for the design lift coefficients for the two extremes of the flight envelope:
1) Straight flight at maximum speed; approximately 300 mph.
2) Sustained 3 G turn decelerating to approx. 200 mph.
Aircraft weight: 725 lbs
Wing Area: 66 ft^2
Dynamic pressure calculated for standard conditions at 6,000 ft.
Using the lift equation: L = Cl X q X V
Where L = lift (weight), Cl = lift coefficient, q = dynamic pressure, and V = wing area
I came up with .0571 for condition #1 and .3856 for condition #2. Is there anything I'm not considering for the turn condition, if 3G is sustained, and the speed doesn't drop below 200 mph?
I came up with the same numbers running them by hand. Can't think of anything you missed for a simple Cl calculation.
-
Can't edit the original post.
So, the density at 6000ft is something like a 0.0499 lbs/ft3.
Mass in most aerodynamics work is expressed in slugs. In an ICAO 6000' standard atmosphere it would be .001987 slugs/ft3.
-
However, I'm wondering a bit your aproach because apparently you have allready selected the wing profile. Normal way is to determine general layout of the plane and assumed Cl range first and choose profile and fine tune the layout based on that.
Well, this is my math used to determine the design lift coefficient. If I can determine the highest and lowest values needed during a lap, it will guide my selection. I haven't determined the airfoil to use yet--I'm still testing my choices. But, regardless of the airfoil, the design lift coefficient will be the same. I'm actually still changing the airfoil and mean lines to see which one will be the best compromise between the straight flight and turns.
I already know how much area the wing will have, the aspect ratio, the approximate weight of the aircraft, etc. My question was to check and see if I had missed anything in making my determination, especially for the lift coefficient necessary during the turn.