candidate for magic airfoil section?

Discussion in 'Aircraft Design / Aerodynamics / New Technology' started by Starflight, Dec 3, 2010.

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  1. Dec 6, 2010 #21

    orion

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    Im not 100% sure but as far as I know there is really no evidence to support this "dangerous" behavior from the so called unpredictable stalls. No, the five digit series of sections are not exactly optimal but for a "turbulent" section family they do have good l/d numbers and have a low moment coefficient as compared to similar counterparts.

    True, the NACA data suggests that the sections may exhibit abrupt boundary layer separation but the wide practical application of the shape and decades of history suggests otherwise. The family has been used by Taylorcraft, Cessna (all the way up to the Citation), Douglas (DC-4 to the DC-7), Dornier, Embraer, Helio, Piel Emeraude, and on and on. And in none of those cases have there been any documented cases of deadly or for that matter, even dangerous or unpredictable behavior (at least in 30 years I haven't seen anything conclusive in that direction).
     
  2. Dec 6, 2010 #22

    topspeed100

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    What kinda planes fly at 2 000 000 Re number ?

    It does look good also slightly thinner; Airfoil Investigation Database - Showing HQ 2.5/12

    stall angle 9 !
     
    Last edited: Dec 6, 2010
  3. Dec 6, 2010 #23

    HumanPoweredDesigner

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    You would not know from the lower Re data.
    Airfoil Investigation Database - Showing GOE 324 (HANSA-BRANDENBURG)

    2,000,000 sounds about like typical utralights.

    Just remember the data is based on a perfectly sharp trailing edge. But the smaller the trailing angle, the less likely you will be able to get it that sharp and still be strong. You can use a thick foil with a smaller cord, or a thin foil with a bigger cord, and for the same spar depth both might have the same lift and drag unless one of them is much better than the other.
     
    Last edited: Dec 6, 2010
  4. Dec 6, 2010 #24

    karoliina.t.salminen

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    The latter simulations are for 4000000.

    The 2000000 simulation was for high altitude flight with high aspect ratio and thus short chord. That is because I am particularly interested in that use case (HALE).

    Not very high altitude is required for the Re 2000000:

    Velocity = 100 m/s
    Wing chord = 0.6 m
    Altitude = 25000 ft
    -> Re = 2138958 in standard atmosphere
     
    Last edited: Dec 6, 2010
  5. Dec 6, 2010 #25

    Norman

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    Here's a good atmospheric properties calculator. Reynolds number is in the bottom frame.
     
  6. Dec 23, 2010 #26

    topspeed100

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  7. Dec 23, 2010 #27

    autoreply

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    Not really. Several modern profiles easily go over 200-250. Also note that these are calculated values, not actual ones and I've seen a difference of 50% or more between those a couple of times...
    The older FX-series are measured in numerous reports so you might have a look there.
     
  8. Dec 23, 2010 #28

    Mac790

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    Jarno forget to say that most of those modern profiles are kept in safe deposit boxes, and you won't get them, I would compare our situation to a little kid who stare at a bowl with ice cream through a shop glass, dreaming about eating it, but can't buy it. Of course someone can measure it, but who has access for those gliders knowledge, tools, etc for it.

    Top, most of those high L/D sailplanes airfoils have big camber, which usually means high cm (not to mention cl for those foils), you really don't want high cm in an airplane, you could "remove" some of this camber and make it more useful but you will lose some L/D. The best coordinates which are available for a sailplane airfoil are those for HQ17, you won't find better ones period (FX79 looks also not bad, but I would like to see a proper chart for it, from what I saw it has narrow drag bucket, at high Cl, I did some mods for FX67 series (just for fun), smaller camber, thicker, bigger LE radius and got better results than for FX79, of course I don't have much faith in those cheap programs) . The problem about HQ17 is camber, you can "remove" some of it, but it will require some analysis, and a couple of phone calls to Harry :).

    Seb
     
  9. Dec 23, 2010 #29

    Turner

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    Orion and Starflight

    The NACA 23000 series is said to have been used by Kurt Tank. I'm unsure where the true origin of this airfoil comes from, but NACA didn't invent it. Tank is said to have used the 23009-23015 in the FW190, but when comparing it to the actual FW190 airfoil it's only a close approximation. So it's not entirely true that it is a NACA airfoil, Tank used a different way to get to the airfoil he used.

    It's interesting to note that the FW190 had a very sharp separation which resulted in sharp stalls. Especially in a fighter application with high wing loading the G stall behaviour of the FW190 series was notoriously nasty, f ex. The lighter early versions however were more nimble than their 109 counterpart of the time.

    Not sure if this really added anything to the discussion but I guess the point I want to make is that the 23009-23015 is only a close approximation of the airfoil used in the FW190. Most RC model and replica builders do not use the original airfoil by Kurt Tank, at least I've not seen one yet. Considering the change in R number the original airfoil might not even be relevant/desirable or be able to replicate original flight characteristics at all.
     
  10. Dec 23, 2010 #30

    Norman

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    Reflexing an given airfoil doesn't always result in lower L/D. True it lowers CLmax but it often also lowers CD over a wide range of AoA. The two 19% sections in the attachment are exactly the same in the forward 60%. The original (which was the sailplane section of choice in the early '70s) has best L/D 11 points higher than the reflexed section. Obviosly that superior L/D is entirely due to the higher CL because the reflexed section has lower drag. A flap should fix that, right?
     

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  11. Dec 23, 2010 #31

    Mac790

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    Yes Norman you are right reflexing (thanks for new English term) doesn't always result in lower L/D, but usually decrease in Cl is bigger than decrease in Cd (at least it were for those airfoils I played with)

    That's good remark, for some airfoils it works better for some not as good, pic1 HQ17 green original 4,51% camber, blue camber reduced to 1.98%, Cd range is about same, only a little bit lower.

    But generally it seems that some airfoils are better for modifications, than others, take a look at pic 2, original FX67(green), versus modified FX67(blue), smaller camber, thicker, bigger LE radius.

    I'm wondering about Profili2 did you compare your results with published papers? Because sometimes I noticed big differences between DesignFoil, and XFLR5, like I said previously those softwares are good for general ideas, but I wouldn't trust my life with them.

    Seb
     

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  12. Dec 24, 2010 #32

    Norman

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    I wasn't suggesting that anybody bet his life on an unproven airfoil. I know that panel codes don't deal very well with separation and therefor aren't accurate outside of the linear range. But comparing results from different wind tunnels isn't such a good idea either. All I was doing was running two airfoil sections side by side in the same software under identical settings. It was intended to show the relative behavior of each and not the absolute characteristics of either.
     
  13. Dec 25, 2010 #33

    Mac790

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    Well Norman you misunderstood me, I know that those cheap programs are in most cases useless for separations prediction, to predict them you need to use proper programs, CFD ideally (you can compare 747A415 at different AOA, sorry for no scale for it and poor quality, I have only free case reader in home, anyway first two velocity magnitude, second two turbulent viscosity) .

    But there are two problems with them, first access to those programs, for students it isn't extremely hard, there are also free CFD programs like Open Foam, which is hard to use, but thanks to my schoolmate I found another one much simpler to use (also free with some limitations), developed by if I remember correctly Sweden military, I'm not using it currently because I have limited access for other programs, but I'm going to learn it within a month/s.

    The second problem is time, there are different models for simulations, Spalart, K-epsilon, k-omega, etc, you can make your simulations more or less accurately, it takes between 15 min to even sometimes more than hour for one AOA, now imagine how much time you will need to check out 30-40 airfoils between -5 and 20 AOA, it will take literally weeks even for single Re.

    What I'm looking for, is "good" software for general selection, but it seems that even at low AOA, those programs have different results even among them. For example in DesignFoil NACA 67 looks good, but in XFLR5 it's a disaster.

    The general idea, is to choose around 30-60 airfoils, and first test them in some cheap software XFLR5, Xfoil, DesignFoil, etc, next choose 5-8, and test them in 2D CFD, next choose, another 2-3 and test them in 3D CFD again. For example I was pretty excited about Harry's 42E series, but I was very disappointed after I "tested' in in XFLR5.

    edit btw, Norman could you put those coordinates in your program, I got really good results in XFLR5 for 4.2x10^6 Re, but not as good in DesignFoil.

    Seb
     

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    Last edited: Dec 25, 2010
  14. Dec 26, 2010 #34

    HumanPoweredDesigner

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    http://www.worldofkrauss.com/foils/show_compare/?id[]=1103&id[]=483&chord=6.5

    That is the difference that smoothing the coordinates can make. Most of the airfoils in the data base are jagged and have polars accordingly.

    What I've learned from the polars posted in this thread is that the Cl climb of GOE 679 is duplicated by plenty of software.

    I had another look at the data base, and what I noticed by looking at many naca families is that the thicker foils usually have a higher Cl max, but it only helps them on the top end of the Cl range, I suppose like leading edge slats, while drag is increased through the whole range, except maybe the postponed seperation. If you look at family after family, Cl max goes up with the thickness probably because of increased frontal radius, and L/D max goes down. The higher cambered sections have better L/D because of higher Cl. But for a wing that can change camber via flaps, the thinner sections seem to be aerodynamically better, and the thicker ones are a structural compromise for cruise and aerodynamically justified mainly by their stall speed.
     
    Last edited: Dec 27, 2010
  15. Dec 27, 2010 #35

    Starflight

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    Thanks HPD for the complimentary assessment of my chosen foil *grins with pride*. My "faux' degree in aerodynamics has the qualifier S.T.E.
    attached (shade tree engineer...LOL). My curiosity about thick foil performance was triggered by watching older bush flying Fokker aircraft
    carrying heavy loads into short strips using wings with large leading edge radius. Have any of you looked at, what I would term, a medium
    camber group originated from a NACA 3xxx four digit airfoil collection? I found an interesting online configuration drawing tool contained in
    a rapid-learn aerodynamics site that allowed me to fabricate profiles using different camber position and thickness; four digit only :/
    Airfoil Geometry located with a short scroll down page. Some of these look like they
    would have great potential for G/A (more get-up than 2xxx and less Cm than 4xxx).
     
  16. Dec 27, 2010 #36

    Norman

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    The text file you sent didn't have enough coordinates so I smoothed it and the trailing edge was a mess. I straitened the TE out a bit but not much and left it open. I just left the settings where they were last time I used the program since you didn't specified anything.

    Mach=0.075
    Ncrit=11
     

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  17. Dec 27, 2010 #37

    Mac790

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    Thanks Norman, sorry for troubles, it seems that result which you got for higher Re are very similar to those I got with XFLR5, at low Re it looks really strange but I just checked out some Harry's foils and there was no big difference.

    If you take a look at those attachment you will notice difference between XFLR5 and DesignFoil, I'm wondering which one is closer to the true results. I've read somewhere that DesignFoil is better for drag bucket estimation, it generally also has lower Cl for particular foils than XFLR5, and even Orion wrote that Xfoil isn't very good at predicting Cl max. https://www.homebuiltairplanes.com/forums/aircraft-design-aerodynamics-new-technology/4154-maximum-lift-values.html

    Seb
     

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  18. Dec 27, 2010 #38

    topspeed100

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    Here is more of the HQ 17; http://www.scielo.org.ar/pdf/laar/v38n3/v38n3a01.pdf

    Seem to me that thickest point is 45-50% and they look very laminar these high performance foils.
     
  19. Dec 27, 2010 #39

    Mac790

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  20. Dec 27, 2010 #40

    topspeed100

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    Last edited: Dec 27, 2010

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