# Thickness/Chord Question



## Zipper730 (Nov 15, 2019)

I was thinking about thickness-to-chord ratio and the point on the wing (% chord) where the thickest area is: Some WWII aircraft have the thickest area around 20-30% (the general subsonic foil), some (such as laminar flow foils) around 37-45% (Tempest: 38%; P-51: 40%; CAC 15: 45%). I'm curious how to compute the effect on thickness to chord on the position of the thickest part of the wing.

For example, if I were to move the thickest part from...

25-38%
25-40%
25-45%
How much thicker can the wing be made for similar drag? I ask this based on the fact that, if the thickest part is further back, the wing behaves (generally) like it is effectively thinner than it was before, and one can compensate by fattening it up.


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## swampyankee (Nov 15, 2019)

CFD. Xfoil may be enough.


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## Zipper730 (Nov 15, 2019)

swampyankee said:


> CFD. Xfoil may be enough.


Firstly, what's X-Foil? Secondly: There's no basic rule of thumb estimate? I'm not planning on building an aircraft, but simply making a drawing that follows some basic realism.


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## swampyankee (Nov 15, 2019)

Zipper730 said:


> Firstly, what's X-Foil? Secondly: There's no basic rule of thumb estimate? I'm not planning on building an aircraft, but simply making a drawing that follows some basic realism.


XFOIL is an airfoil analysis program. If you're drawing airfoils, the NACA four-digit and five-digit airfoils have algebraic formulas for their contour; the NACA 6 series (laminar flow) airfoils don't. On the other hand, vast numbers of airfoils have their coordinates on the UIUC Airfoil Database (UIUC Airfoil Data Site)

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## eljefe (Nov 21, 2019)

To the best of my knowledge (and recall, I am an engineer and a pilot, but not an aero engineer), there is no simple formula that will let you plug in a few numbers that represent an airfoil in terms of shape that will give you an idea of lift or drag on that particular airfoil. For many years, the study of airfoils was very empirical, with airfoil after airfoil submitted to a wind tunnel and the results carefully cataloged (mostly by NACA but also by similar organizations around the world).

Airfoil design in that era consisted largely of plucking a set of numbers from a reference that met the proposed performance of the new aircraft, and fabricating a wing that matched the description of that airfoil in the reference. If nothing was available that met the specs, or someone thought they could "invent a better mousetrap", that involved fabricating a slightly different airfoil from the ones in the books, and testing, testing, testing (again, mostly by NACA).

That didn't change until computers became more and more available, and solving complicated fluid-flow equations through numerical means became economically practical. You theoretically COULD design a new airfoil by using the Navier-Stokes equations by putting together a lot of people with slide rules in a room, but it would take so long and cost so much as to be impractical unless you had the budget to make a gold-plated aircraft fly. But with modern computers, which are the equivalent of literally dozens of Cray-1 supercomputers from the 80s on a desk, and modern supercomputers which are literally the equivalent of millions of such outdated machines, numerically "solving" a proposed airfoil has become feasible. You can do it yourself to a level better than any "napkin calculation" if you understand aerodynamics, the software package you're using, and have the software and sufficient computing power to run it on.

Hence, CFD which is "Computational Fluid Dynamics". There are several commercially avialable CFD packages, but one that will reliably predict performance on a real airfoil would probably cost more than you or any other hobbyist would be willing to spend. XFOIL is freeware developed by MIT in the 1980s that has been reported several times and is still valuable today, but like most freeware is inadequate for professional use. As was mentioned, it might be good enough for your purposes, though.


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## Zipper730 (Nov 21, 2019)

What I was thinking of was something akin to rise/run: If the wing was thicker further back, the curve from the start of the foil to the rear of the foil would be less abrupt... similar to a shallow slope versus a highly steep slope.

I was curious if you could use that to apply some kind of relationship.


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## swampyankee (Nov 21, 2019)

There are quite a few quite good, not hideously* expensive CFD systems out there. Generally, the more expensive a CFD system is, the more expertise is required to use it properly. 

XFOIL is actually pretty decent for 2-D airfoils, and if I were in the business of designing light aircraft, it's probably good enough. On the other hand, there is a reason that some people (formerly, John Roncz; possibly Harry RIblett is still active, although is diatribe against the 23012 is nonsensical). 

-----
* Obviously, different people have different ideas of how much is involved in something being "hideously expensive." At some point, it probably makes more sense to just write a check to someplace like Analytical Methods or someone like Riblett to design your airfoil. There are people...


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## pbehn (Nov 21, 2019)

Zipper730 said:


> I was thinking about thickness-to-chord ratio and the point on the wing (% chord) where the thickest area is: Some WWII aircraft have the thickest area around 20-30% (the general subsonic foil), some (such as laminar flow foils) around 37-45% (Tempest: 38%; P-51: 40%; CAC 15: 45%). I'm curious how to compute the effect on thickness to chord on the position of the thickest part of the wing.
> 
> For example, if I were to move the thickest part from...
> 
> ...


The thickest point isn't a thickness it is it is a proportion of the chord. The root of a Mustangs wing is thick (enough to put but fuel tanks in) and tips aren't. A laminar flow wing doesnt behave as if it is thinner it postpones the transition from laminar to turbulent flow until further along the chord, the reasons arguments and justifications for why it does are above my pay grade.

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## pbehn (Nov 22, 2019)

Zipper730 said:


> What I was thinking of was something akin to rise/run: If the wing was thicker further back, the curve from the start of the foil to the rear of the foil would be less abrupt... similar to a shallow slope versus a highly steep slope.
> 
> I was curious if you could use that to apply some kind of relationship.


 The relationship starts to be defined as follows, which I am sure will fill you with undiluted pleasure after reading. from wiki Fluid dynamicists define the chord Reynolds number _R_ like this: _R_ = _Vc_/_ν_, where _V_ is the flight speed, _c_ is the chord length, and _ν_ is the kinematic viscosity of the fluid in which the airfoil operates, which is 1.460×10−5 m2/s for the atmosphere at sea level.[18] In some special studies a characteristic length other than chord may be used; rare is the "span Reynolds number", which is not to be confused with spanwise stations on a wing, where chord is still used.


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## Zipper730 (Nov 22, 2019)

pbehn said:


> The thickest point isn't a thickness it is it is a proportion of the chord.


Well, what I was trying to get at was "if I move the thickness further back" what effect would that have on allowing the T/C ratio to be larger -- I might not have articulated that well enough.


> A laminar flow wing doesnt behave as if it is thinner it postpones the transition from laminar to turbulent flow until further along the chord, the reasons arguments and justifications for why it does are above my pay grade.


If I recall part of it was that the thickest part was further aft (more gradual curvature), and cusps on the rear section that produce a suction which, while it seems bad, keeps the flow attached by keeping it "sucked" on down to the wing's surface.


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## pbehn (Nov 22, 2019)

Zipper730 said:


> Well, what I was trying to get at was "if I move the thickness further back" what effect would that have on allowing the T/C ratio to be larger -- I might not have articulated that well enough.
> If I recall part of it was that the thickest part was further aft (more gradual curvature), and cusps on the rear section that produce a suction which, while it seems bad, keeps the flow attached by keeping it "sucked" on down to the wing's surface.


You cant simplify it to it being a thicker point further back, or laminar flow being sucked onto the wing. Here is a picture of the transition from laminar to turbulent flow in a candles flam/hot air stream such things were being studied before the Wright brothers took to the air, Reynolds died in 1912.


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## Zipper730 (Nov 22, 2019)

pbehn said:


> Here is a picture of the transition from laminar to turbulent flow in a candles flam/hot air stream . . . .
> 
> 
> View attachment 561604


That's a cool picture


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## pbehn (Nov 22, 2019)

Zipper730 said:


> That's a cool picture



It is a picture of something we have all seen, it takes very great minds to question what is going on and then bring it down to mathematical expressions. There are two similar effects in metallurgy, the time temperature transformation of metals as they cool from liquid to solid and the ductile to brittle transformation of steels (and others). These are extremely complex issues and the more you study them the more there is to study, they cannot be simplified. I state with absolute confidence that anything you read on the net or on here posted by people like Drgondog has been simplified for a laymans understanding and omits a massive amount of special situation detail.


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## swampyankee (Nov 23, 2019)

When a flow is accelerating along a curved surface, like on the forward part of an airfoil, the pressure decreases along the surface, which will tend to stabilize the boundary layer against separation. Where the flow is decelerating, the pressure increases along the surface, which increases the likelihood of separation. Turbulent flow tends to be more resistant to separation than laminar flow; this is why golf balls have dimples: they cause the flow to become turbulent. This means that the flow separates farther from the nose of the ball, reducing its drag. One of the pressure distribution for the rear part of airfoils is called the Stratford separation criterion (Stratford's separation criterion -- CFD-Wiki, the free CFD reference).

For some very good discussions of airfoils, in general, try CFD Online (CFD Online)

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## Zipper730 (Dec 13, 2019)

I downloaded the program, I know almost nothing about programming, so I'll see what I can even input.

What designation is there for the following aircraft's foils

Supermarine Spitfire
Root
Mid-Span
Tip

North American P-51 Mustang
Root
Mid-Span
Tip

De Havilland Mosquito
Root
Mid-Span
Tip

Hawker Tempest
Root
Mid-Span
Tip


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## swampyankee (Dec 13, 2019)

Zipper730 said:


> I downloaded the program, I know almost nothing about programming, so I'll see what I can even input.
> 
> What designation is there for the following aircraft's foils
> 
> ...



Use The Incomplete Guide to Airfoil Usage: The Incomplete Guide to Airfoil Usage to find the airfoils used and the UIUC Airfoil Database at UIUC Airfoil Data Site to get the coordinates.

Without more data -- the IGAU only lists tip and root airfoils -- it's not possible to be sure of the airfoil at mid-span. A reasonable guess would be to linearly interpolate from root to tip airfoil.


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## Zipper730 (Dec 14, 2019)

Airfoil tools seems a better site for data on airfoils... I found the Tempest foil real quick

Airfoil Tools


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