X-Plane manual

Designing Airfoils in Part-Maker

Every airfoil ever designed has it's own characteristics, which are its coefficients of:
lift, (how much the airfoil wants to lift up)
drag, (how much the airfoil wants to pull back), and
moment (how much the airfoil wants to pitch up).

What you see on the Screen

You'll see a big black box dominating the screen with green, red, and yellow lines on it.

The left edge of the chart corresponds to an angle of attack of  20 degrees, and the right edge corresponds to an angle of attack of +20 degrees.

The center of the chart represents an angle of attack of zero degrees. (Remember the angle of attack is the angle of the wing to the air. It is the angle at which the wing hits (or "attacks") the air).

The green line is the coefficient of lift. The red line is the coefficient of drag. The yellow line is the coefficient of moment. We'll look at the behavior of each of these lines.

You will see a number labeled "Reynolds number" on the upper left. The "Reynolds number" is simply the air density times the speed of the airplane times the chord of the wing divided by the viscosity of air (Wow!). Experiments have shown that the coefficients of lift, drag and moment of wings vary somewhat with Reynolds number. For recreational purposes, you can probably neglect any change in performance with Reynolds number, so you can just ignore this setting altogether. The number entered in the Reynolds number box may have some impact however on the simulation. For highest realism you can generate 2 different airfoil files for the same airfoil, each file at a different Reynolds number, and assign them both to your wing! X-Plane will figure out the Reynolds number on each piece of the plane at least 10 times per second and interpolate between the two airfoil files to give the most realistic coefficients for that flight Reynolds number.

Pilots should realise: very good accuracy can be obtained without messing with the Reynolds number at all, and without generating two airfoil files for each airfoil. You can ignore the above paragraph and the "Reynolds number" slot in the airfoil generation screen without sacrificing a good simulation.

What the coefficients are:

Coefficient of lift
Look at the green line. It is the coefficient of lift.
Notice that at zero degrees angle of attack (center of screen) the coefficient of lift is fairly low. (It is close to the thin white line, which represents zero). As the angle of attack increases, the coefficient of lift increases right along with it, until you get to around 16 degrees angle of attack, at which point the coefficient of lift falls abruptly... That is the stall! If you go to negative angles of attack, you see that the coefficient of lift actually gets negative. If you go to a large enough negative angle of attack, the airfoil stalls then, too. It is possible to stall upside down! A good wing will have a decent coefficient of lift (maybe 0.4) at angles of attack close to zero, and a nice high coefficient of lift (maybe 1.6) at the maximum angle of attack. A safe airfoil will also have a stall that is not too abrupt. In other words, the coefficient of lift will fall off gradually at the stall, rather than sharply.

Coefficient of drag
Look at the red line. It is the coefficient of drag.
Notice that the coefficient of drag is lowest close to zero degrees angle of attack. The drag gets higher and higher as the wing goes to larger and larger angles of attack. That is not surprising, is it? The higher the angle you offset the wing from the airflow, the greater the drag!
It doesn't matter much whether you are going to positive or negative angles of attack (aiming the wing up or down)... moving the wing away from it's most streamlined position increases it?s drag. A good airfoil will obviously have the lowest drag possible. (Notice that this drag coefficient does NOT include the drag due to the production of lift. X-Plane will figure this drag out automatically).

Coefficient of moment
Look at the yellow line. It is the coefficient of moment.
The coefficient of moment is the tendency of the wing to pitch up about its axis, or rotate upwards about the spar. Most wings actually want to pitch down, so the coefficient of moment is usually negative. The moment varies a bit with angle of attack, often in ways that are a little bit surprising. Typically the moment will be negative for all normally-encountered angles of attack, getting especially large in the negative direction as the angle of attack is increased, until the stall, at which point the moment heads back to zero. A desirable characteristic of an airfoil is usually to have a low coefficient of moment.

General Info

Coefficient display box
One thing that you have probably noticed is that the axis are not labeled, and don't have numerical values to tell you exactly what the coefficients are. Look at the little box in the upper left-hand corner of the airfoil generation screen. The top number (white) is ?alpha? or the angle of attack of the wing. The next numbers are the coefficients of lift, drag, and moment at that angle of attack. Wiggle the mouse back and forth all the way across the monitor, and notice that the angle of attack display changes, and the coefficients with it. The coefficient display box is giving the angle of attack and coefficients of the airfoil at the angle of attack that the mouse is currently pointing at. Just point the mouse at the part of the curve you are interested in, and look at the exact coefficients in the coefficient display box! Easy!

One question you might be asking yourself is: How do I find what the coefficients are for the airfoils on my airplane?
First, you need to find what airfoil your aircraft uses, probably from the manufacturer. Then you need to see if that airfoil is included with our program. If you are flying a Cessna 182, for example, that aircraft uses the NACA 2412 airfoil, which is included, so you do NOT need to generate your own airfoil for that wing. If you do not know what foil to use, then just leave them as the defaults of Plane-Maker.

Airfoil selection is a fun and interesting process, because you will be looking for the best possible combination lift, drag, and moment characteristics for your particular airplane. If you will be experimenting with your own airplane designs, and are new to the matters discussed in this manual, we highly recommend:

R/C Model Airplane Design
A.G. Lennon
Motorbooks International Publishers and Wholesalers, Inc.

to get you started. The book is intended for radio control designs, but is very straightforward, easy to understand, and all of the principles apply to full-scale aircraft. General Info

Once you understand the basics of airfoil theory and nomenclature, we recommend

Theory of Wing Sections
Abbot and Von Doenhoff
McGraw-Hill, New York (1949)

...an oldie but goodie! This books has the lift, drag, and moment plots of many airfoils in it, so you can choose your favorite airfoil for your design and then enter it into the computer using the technique that is about to be explained.

In the following discussion, thin symmetrical, thick highly cambered, and "normal general aviation" airfoils will be discussed. These are three types of airfoils that are good for discussion purposes because they are so different.

Thin symmetrical airfoils are thin and have the same shape on both the top and bottom surfaces. They do not produce very much lift or drag. They typically are used for vertical stabilizers and often horizontal stabilizers as well because they are not called upon to produce a lot of lift, and are not expected to produce much drag, either.

Use thick, highly-cambered airfoils in the foreplanes of canards, or other applications where you want a LARGE amount of lift from a SMALL wing area. These foils are known for providing a large amount of drag as the penalty for providing a large amount of lift.

So-called "normal general aviation airfoils", like the NACA 2412, are compromises between the two, and are good candidates for the wing of a general aviation aircraft.

Supercritical, laminar-flow, and other possible groupings of airfoils exist, but for the purposes of our discussion we will concentrate on the thin symmetrical, thick and highly cambered, and "normal general aviation" airfoils just outlined.

Airfoil generation buttons

Now let's actually generate an airfoil. The first button to click on is the coefficient of lift intercept button, the green one labeled "intercept" in the upper left hand corner. To increase this number, just click right above the numbers that you want to increase, and below the ones that you want to decrease. For example, if the lift intercept on the screen is 0.25, and you want to change it to 0.33 to model your airfoil, just click right above the "2" in "0.25" and twice below the "5" in "0.25". You change all of your data that way for the entire design and simulation system. Easy! Now what exactly is a coefficient of lift intercept, anyway? Read on to find out!

Coefficient of lift intercept, "INTERCPT"
This is the coefficient of lift at an angle of attack of 0 degrees. For a symmetrical airfoil, this will always be zero, since the air is doing exactly the same thing on the top and bottom of the wing for a symmetrical airfoil at zero degrees angle of attack. Symmetrical airfoils are sometimes used for horizontal stabilizers, and are almost always used for vertical stabilizers. Sleek, skinny wings with low camber might have a lift intercept of 0.1. Fat, highly cambered foils have a value around 0.6. A typical airfoil like the NACA-2412 (commonly used in general aviation) has a value of about 0.2.

Coefficient of lift slope, "SLOPE"
This is the increase in coefficient of lift per degree increase in angle of attack. A thin airfoil has a value of about 0.1. A really fat airfoil has a value of about 0.08. Fatter airfoils have slightly lower lift slopes. (You will find, however, that lift slopes are almost always very close to 0.1).

Coefficient of lift curvature near the stall, "POWER"
As the angle of attack gets close to stall, the lift slope is no longer linear, but gradually ?levels off? as it approaches the maximum, or stalling, coefficient of lift. Just play with the power button until you find a power curve that connects the linear and stalling regions smoothly. Chances are a power of around 1.5 will work pretty well. Just play with it until the lift comes up smoothly, then gradually levels off to the stall, since that is what happens with a real airfoil.

Coefficient of lift maximum, "MAXIMUM"
This is the maximum coefficient of lift, or the coefficient of lift right before the stall. A very thin, symmetrical airfoil has a value of around 1.0. A thick, highly cambered airfoil has a value of around 1.8. A typical general aviation foil might have a value of around 1.6.

Coefficient of lift immediate drop at stall, "DROP"
This is the drop that immediately follows the stall. For thin airfoils, which tend to stall sharply, this value might be 0.2. For many airfoils, however, there is no immediate drop, but instead a more gradual one as the angle of attack is further increased. In most cases, this number will be zero or very close to zero.

Coefficient of lift curvature after stall "POWER"
Different airfoils have different lift slopes after the stall. For skinny, sharply-stalling airfoils the power should be fairly low, perhaps around 1.4. For fat airfoils (which usually have more gentle stalling characteristics) this number may be closer to 2.0. Just play with the power button until the data looks like the data you are trying to model from the airfoil chart in whatever book you are getting your airfoil data from.

Coefficient of lift drop from stall to 20 degrees "DROP"
This is the decrease in coefficient of lift from the stall to an angle of 20 degrees. This number might be in the 0.4 range for a thicker airfoil, 0.6 for a thinner one.

The NACA-2412 has a value of about 0.4 . (The coefficient of lift goes from around 1.6 to 1.2 as the angle of attack goes from around 16 to 20 degrees).

Coefficient of drag minimum "DMIN"
This is the minimum coefficient of drag of the airfoil. (Again, not including induced drag, which is determined automatically by the simulator ?X-Plane?). This minimum coefficient of drag also should not include the ?low-drag bucket? of a laminar flow wing. A thick or highly cambered airfoil has a value of about 0.01, a typical older general-aviation airfoil such as the NACA-2412 has a value of about 0.006, and a really thin, symmetrical airfoil has about a 0.005 value. Laminar flow airfoils can approach values of 0.004, but that number should not be entered here, because it will be addressed in the laminar drag bucket buttons soon to come...

Coefficient of lift at which minimum drag occurs "MIN D CL"
Enter the coefficient of lift at which the minimum drag occurs. This value is probably very close to the coefficient of lift at zero degrees angle of attack, which is the ?lift intercept?. The very first number you entered! If anything, the minimum coefficient of drag occurs at a coefficient of lift a little lower than the lift intercept coefficient of lift. This is because an airfoil usually has the least drag at an angle of attack of about zero degrees or just a hair lower.

Coefficient of drag at angle of attack of 10 degrees "D ALPH=10"
For a thin, symmetrical airfoil, this value might be around 0.015. NACA-2412 comes in with a surprisingly good 0.012. A really highly-cambered airfoil might be around 0.025, though.

Coefficient of drag curvature "POWER"
The power curve is simply the curvature of the drag curve as it changes with angle of attack. You will have to fiddle with the curvature until the curve looks like the experimental data, but theoretically this number will be around 2.

Laminar drag bucket location "CL LOCTN"
Some airfoils, called ?natural laminar flow? or "NLF" airfoils, have perfectly smooth airflow across a large part of the wing, a flow pattern called "laminar flow" (Where did you think this company got the name "Laminar Research"?) This super-smooth, low-drag flow can only happen at fairly small angles of attack, though, so there is a "low-drag bucket", or area in a small angle of attack range, that has lower-than-normal drag. The drag bucket location is usually thought of in terms of the coefficient of lift. In other words, the center of the drag bucket occurs at some coefficient of lift of the airfoil. This might happen at a coefficient of lift of around 0.6.

Laminar drag bucket width "WIDTH"
This refers to how ?wide? the bucket is, or what range of coefficient of lift the drag bucket covers. 0.4 is a decent guess.
Laminar drag bucket depth "DEPTH"
This is the all-important variable: how much do you reduce your drag by going to laminar flow? Answer: 0.002 if you;re lucky. (But that is actually quite a bit. That might turn a cd of 0.006 to 0.004. Quite a large percentage difference.).

Laminar drag bucket curvature "POWER"
The power curve is the simply the curvature of this low drag bucket. You will have to fiddle with the curvature until the curve looks like the experimental data, but chances are this number will be around 3 to 5.

Coefficient of moment low-alpha change point "ALPHA 1"
The coefficient of moment is usually linear across the non-stalled angle of attack range. In other words, if the airfoil is not stalled, the moment curve is usually a straight line. After the stall, however, the moment coefficient tends to change direction. For the NACA-2412, the moment coefficient has it?s low angle of attack moment-change at 10 degrees, a point corresponding to roughly +4 degrees before the stall.

Coefficient of moment high-alpha change point "ALPHA 2"
The NACA-2412 airfoil has it?s high angle of attack moment-change right at the positive stalling angle of 16 degrees.

Coefficient of moment at 20 degrees "CM 1"
For the NACA 2412, this number is about 0.075. Notice that this is a positive number. This means that if the airfoil is at a clear negative angle of attack, it will stall and try to pitch back up to an angle of attack closer to zero. This is a nice effect, because the airfoil tends to try and recover from the stall automatically.

Coefficient of moment at low-alpha change point "CM 2"
For the NACA 2412, this number is about -0.05, which is a light pitch-down. A wing with a higher camber will have a value of around -0.10, perhaps even 0.13. A symmetrical airfoil will have no pitch tendency at all here, so 0.0 should be entered for that type of airfoil.

Coefficient of moment at high-alpha change point "CM 3"
For the NACA 2412, this number is about 0.025, which is a very light pitch-down. A wing with a higher camber will have a value of around 0.10, perhaps even 0.13. A symmetrical airfoil will have no pitch tendency at all here, so 0.0 should be entered for that type of airfoil.

Coefficient of moment at 20 degrees "CM 4"
This is the coefficient of moment well into the stall. For the NACA 2412, it is about -0.10. This is a moderate pitch-down, which is desirable because this pitch-down will help recover from the stall.

Finishing Up
Change all of the parameters we just discussed around a bit, and select "Save As" from the "File" menu. Now type in an airfoil name and hit return. Congratulations! You have just generated your own airfoil! Drop it in the "Resources:Airfoils" folder in your X-System folder (to be usable by ALL planes) or a folder that you make called "Airfoils" in the same folder as your airplane designs to be used only by that airplane.

(next section)

What you see on the Screen You'll see a big black box dominating the screen with green, red, and yellow lines on it. The left edge of the chart corresponds to an angle of attack of 20 degrees, and the right edge corresponds to an angle of attack of +20 degrees. The center of the chart represents an angle of attack of zero degrees. (Remember the angle of attack is the angle of the wing to the air. It is the angle at which the wing hits (or "attacks") the air). The green line is the coefficient of lift. The red line is the coefficient of drag. The yellow line is the coefficient of moment. We'll look at the behavior of each of these lines. You will see a number labeled "Reynolds number" on the upper left. The "Reynolds number" is simply the air density times the speed of the airplane times the chord of the wing divided by the viscosity of air (Wow!). Experiments have shown that the coefficients of lift, drag and moment of wings vary somewhat with Reynolds number. For recreational purposes, you can probably neglect any change in performance with Reynolds number, so you can just ignore this setting altogether. The number entered in the Reynolds number box may have some impact however on the simulation. For highest realism you can generate 2 different airfoil files for the same airfoil, each file at a different Reynolds number, and assign them both to your wing! X-Plane will figure out the Reynolds number on each piece of the plane at least 10 times per second and interpolate between the two airfoil files to give the most realistic coefficients for that flight Reynolds number. Pilots should realise: very good accuracy can be obtained without messing with the Reynolds number at all, and without generating two airfoil files for each airfoil. You can ignore the above paragraph and the "Reynolds number" slot in the airfoil generation screen without sacrificing a good simulation.			What the coefficients are: Coefficient of lift Look at the green line. It is the coefficient of lift. Notice that at zero degrees angle of attack (center of screen) the coefficient of lift is fairly low. (It is close to the thin white line, which represents zero). As the angle of attack increases, the coefficient of lift increases right along with it, until you get to around 16 degrees angle of attack, at which point the coefficient of lift falls abruptly... That is the stall! If you go to negative angles of attack, you see that the coefficient of lift actually gets negative. If you go to a large enough negative angle of attack, the airfoil stalls then, too. It is possible to stall upside down! A good wing will have a decent coefficient of lift (maybe 0.4) at angles of attack close to zero, and a nice high coefficient of lift (maybe 1.6) at the maximum angle of attack. A safe airfoil will also have a stall that is not too abrupt. In other words, the coefficient of lift will fall off gradually at the stall, rather than sharply. Coefficient of drag Look at the red line. It is the coefficient of drag. Notice that the coefficient of drag is lowest close to zero degrees angle of attack. The drag gets higher and higher as the wing goes to larger and larger angles of attack. That is not surprising, is it? The higher the angle you offset the wing from the airflow, the greater the drag! It doesn't matter much whether you are going to positive or negative angles of attack (aiming the wing up or down)... moving the wing away from it's most streamlined position increases it?s drag. A good airfoil will obviously have the lowest drag possible. (Notice that this drag coefficient does NOT include the drag due to the production of lift. X-Plane will figure this drag out automatically). Coefficient of moment Look at the yellow line. It is the coefficient of moment. The coefficient of moment is the tendency of the wing to pitch up about its axis, or rotate upwards about the spar. Most wings actually want to pitch down, so the coefficient of moment is usually negative. The moment varies a bit with angle of attack, often in ways that are a little bit surprising. Typically the moment will be negative for all normally-encountered angles of attack, getting especially large in the negative direction as the angle of attack is increased, until the stall, at which point the moment heads back to zero. A desirable characteristic of an airfoil is usually to have a low coefficient of moment.




General Info Coefficient display box One thing that you have probably noticed is that the axis are not labeled, and don't have numerical values to tell you exactly what the coefficients are. Look at the little box in the upper left-hand corner of the airfoil generation screen. The top number (white) is ?alpha? or the angle of attack of the wing. The next numbers are the coefficients of lift, drag, and moment at that angle of attack. Wiggle the mouse back and forth all the way across the monitor, and notice that the angle of attack display changes, and the coefficients with it. The coefficient display box is giving the angle of attack and coefficients of the airfoil at the angle of attack that the mouse is currently pointing at. Just point the mouse at the part of the curve you are interested in, and look at the exact coefficients in the coefficient display box! Easy! One question you might be asking yourself is: How do I find what the coefficients are for the airfoils on my airplane? First, you need to find what airfoil your aircraft uses, probably from the manufacturer. Then you need to see if that airfoil is included with our program. If you are flying a Cessna 182, for example, that aircraft uses the NACA 2412 airfoil, which is included, so you do NOT need to generate your own airfoil for that wing. If you do not know what foil to use, then just leave them as the defaults of Plane-Maker. Airfoil selection is a fun and interesting process, because you will be looking for the best possible combination lift, drag, and moment characteristics for your particular airplane. If you will be experimenting with your own airplane designs, and are new to the matters discussed in this manual, we highly recommend: R/C Model Airplane Design A.G. Lennon Motorbooks International Publishers and Wholesalers, Inc. to get you started. The book is intended for radio control designs, but is very straightforward, easy to understand, and all of the principles apply to full-scale aircraft.			General Info Once you understand the basics of airfoil theory and nomenclature, we recommend Theory of Wing Sections Abbot and Von Doenhoff McGraw-Hill, New York (1949) ...an oldie but goodie! This books has the lift, drag, and moment plots of many airfoils in it, so you can choose your favorite airfoil for your design and then enter it into the computer using the technique that is about to be explained. In the following discussion, thin symmetrical, thick highly cambered, and "normal general aviation" airfoils will be discussed. These are three types of airfoils that are good for discussion purposes because they are so different. Thin symmetrical airfoils are thin and have the same shape on both the top and bottom surfaces. They do not produce very much lift or drag. They typically are used for vertical stabilizers and often horizontal stabilizers as well because they are not called upon to produce a lot of lift, and are not expected to produce much drag, either. Use thick, highly-cambered airfoils in the foreplanes of canards, or other applications where you want a LARGE amount of lift from a SMALL wing area. These foils are known for providing a large amount of drag as the penalty for providing a large amount of lift. So-called "normal general aviation airfoils", like the NACA 2412, are compromises between the two, and are good candidates for the wing of a general aviation aircraft. Supercritical, laminar-flow, and other possible groupings of airfoils exist, but for the purposes of our discussion we will concentrate on the thin symmetrical, thick and highly cambered, and "normal general aviation" airfoils just outlined.




Airfoil generation buttons Now let's actually generate an airfoil. The first button to click on is the coefficient of lift intercept button, the green one labeled "intercept" in the upper left hand corner. To increase this number, just click right above the numbers that you want to increase, and below the ones that you want to decrease. For example, if the lift intercept on the screen is 0.25, and you want to change it to 0.33 to model your airfoil, just click right above the "2" in "0.25" and twice below the "5" in "0.25". You change all of your data that way for the entire design and simulation system. Easy! Now what exactly is a coefficient of lift intercept, anyway? Read on to find out! Coefficient of lift intercept, "INTERCPT" This is the coefficient of lift at an angle of attack of 0 degrees. For a symmetrical airfoil, this will always be zero, since the air is doing exactly the same thing on the top and bottom of the wing for a symmetrical airfoil at zero degrees angle of attack. Symmetrical airfoils are sometimes used for horizontal stabilizers, and are almost always used for vertical stabilizers. Sleek, skinny wings with low camber might have a lift intercept of 0.1. Fat, highly cambered foils have a value around 0.6. A typical airfoil like the NACA-2412 (commonly used in general aviation) has a value of about 0.2. Coefficient of lift slope, "SLOPE" This is the increase in coefficient of lift per degree increase in angle of attack. A thin airfoil has a value of about 0.1. A really fat airfoil has a value of about 0.08. Fatter airfoils have slightly lower lift slopes. (You will find, however, that lift slopes are almost always very close to 0.1). Coefficient of lift curvature near the stall, "POWER" As the angle of attack gets close to stall, the lift slope is no longer linear, but gradually ?levels off? as it approaches the maximum, or stalling, coefficient of lift. Just play with the power button until you find a power curve that connects the linear and stalling regions smoothly. Chances are a power of around 1.5 will work pretty well. Just play with it until the lift comes up smoothly, then gradually levels off to the stall, since that is what happens with a real airfoil. Coefficient of lift maximum, "MAXIMUM" This is the maximum coefficient of lift, or the coefficient of lift right before the stall. A very thin, symmetrical airfoil has a value of around 1.0. A thick, highly cambered airfoil has a value of around 1.8. A typical general aviation foil might have a value of around 1.6. Coefficient of lift immediate drop at stall, "DROP" This is the drop that immediately follows the stall. For thin airfoils, which tend to stall sharply, this value might be 0.2. For many airfoils, however, there is no immediate drop, but instead a more gradual one as the angle of attack is further increased. In most cases, this number will be zero or very close to zero. Coefficient of lift curvature after stall "POWER" Different airfoils have different lift slopes after the stall. For skinny, sharply-stalling airfoils the power should be fairly low, perhaps around 1.4. For fat airfoils (which usually have more gentle stalling characteristics) this number may be closer to 2.0. Just play with the power button until the data looks like the data you are trying to model from the airfoil chart in whatever book you are getting your airfoil data from. Coefficient of lift drop from stall to 20 degrees "DROP" This is the decrease in coefficient of lift from the stall to an angle of 20 degrees. This number might be in the 0.4 range for a thicker airfoil, 0.6 for a thinner one. The NACA-2412 has a value of about 0.4 . (The coefficient of lift goes from around 1.6 to 1.2 as the angle of attack goes from around 16 to 20 degrees). Coefficient of drag minimum "DMIN" This is the minimum coefficient of drag of the airfoil. (Again, not including induced drag, which is determined automatically by the simulator ?X-Plane?). This minimum coefficient of drag also should not include the ?low-drag bucket? of a laminar flow wing. A thick or highly cambered airfoil has a value of about 0.01, a typical older general-aviation airfoil such as the NACA-2412 has a value of about 0.006, and a really thin, symmetrical airfoil has about a 0.005 value. Laminar flow airfoils can approach values of 0.004, but that number should not be entered here, because it will be addressed in the laminar drag bucket buttons soon to come... Coefficient of lift at which minimum drag occurs "MIN D CL" Enter the coefficient of lift at which the minimum drag occurs. This value is probably very close to the coefficient of lift at zero degrees angle of attack, which is the ?lift intercept?. The very first number you entered! If anything, the minimum coefficient of drag occurs at a coefficient of lift a little lower than the lift intercept coefficient of lift. This is because an airfoil usually has the least drag at an angle of attack of about zero degrees or just a hair lower. Coefficient of drag at angle of attack of 10 degrees "D ALPH=10" For a thin, symmetrical airfoil, this value might be around 0.015. NACA-2412 comes in with a surprisingly good 0.012. A really highly-cambered airfoil might be around 0.025, though. Coefficient of drag curvature "POWER" The power curve is simply the curvature of the drag curve as it changes with angle of attack. You will have to fiddle with the curvature until the curve looks like the experimental data, but theoretically this number will be around 2. Laminar drag bucket location "CL LOCTN" Some airfoils, called ?natural laminar flow? or "NLF" airfoils, have perfectly smooth airflow across a large part of the wing, a flow pattern called "laminar flow" (Where did you think this company got the name "Laminar Research"?) This super-smooth, low-drag flow can only happen at fairly small angles of attack, though, so there is a "low-drag bucket", or area in a small angle of attack range, that has lower-than-normal drag. The drag bucket location is usually thought of in terms of the coefficient of lift. In other words, the center of the drag bucket occurs at some coefficient of lift of the airfoil. This might happen at a coefficient of lift of around 0.6. Laminar drag bucket width "WIDTH" This refers to how ?wide? the bucket is, or what range of coefficient of lift the drag bucket covers. 0.4 is a decent guess. Laminar drag bucket depth "DEPTH" This is the all-important variable: how much do you reduce your drag by going to laminar flow? Answer: 0.002 if you;re lucky. (But that is actually quite a bit. That might turn a cd of 0.006 to 0.004. Quite a large percentage difference.). Laminar drag bucket curvature "POWER" The power curve is the simply the curvature of this low drag bucket. You will have to fiddle with the curvature until the curve looks like the experimental data, but chances are this number will be around 3 to 5. Coefficient of moment low-alpha change point "ALPHA 1" The coefficient of moment is usually linear across the non-stalled angle of attack range. In other words, if the airfoil is not stalled, the moment curve is usually a straight line. After the stall, however, the moment coefficient tends to change direction. For the NACA-2412, the moment coefficient has it?s low angle of attack moment-change at 10 degrees, a point corresponding to roughly +4 degrees before the stall. Coefficient of moment high-alpha change point "ALPHA 2" The NACA-2412 airfoil has it?s high angle of attack moment-change right at the positive stalling angle of 16 degrees. Coefficient of moment at 20 degrees "CM 1" For the NACA 2412, this number is about 0.075. Notice that this is a positive number. This means that if the airfoil is at a clear negative angle of attack, it will stall and try to pitch back up to an angle of attack closer to zero. This is a nice effect, because the airfoil tends to try and recover from the stall automatically. Coefficient of moment at low-alpha change point "CM 2" For the NACA 2412, this number is about -0.05, which is a light pitch-down. A wing with a higher camber will have a value of around -0.10, perhaps even 0.13. A symmetrical airfoil will have no pitch tendency at all here, so 0.0 should be entered for that type of airfoil. Coefficient of moment at high-alpha change point "CM 3" For the NACA 2412, this number is about 0.025, which is a very light pitch-down. A wing with a higher camber will have a value of around 0.10, perhaps even 0.13. A symmetrical airfoil will have no pitch tendency at all here, so 0.0 should be entered for that type of airfoil. Coefficient of moment at 20 degrees "CM 4" This is the coefficient of moment well into the stall. For the NACA 2412, it is about -0.10. This is a moderate pitch-down, which is desirable because this pitch-down will help recover from the stall. Finishing Up Change all of the parameters we just discussed around a bit, and select "Save As" from the "File" menu. Now type in an airfoil name and hit return. Congratulations! You have just generated your own airfoil! Drop it in the "Resources:Airfoils" folder in your X-System folder (to be usable by ALL planes) or a folder that you make called "Airfoils" in the same folder as your airplane designs to be used only by that airplane.