Decalage angle

Discussion in 'Aircraft Design / Aerodynamics / New Technology' started by Eugene, May 29, 2017.

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  1. May 31, 2017 #81

    BBerson

    BBerson

    BBerson

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    The original designer is usually correct.
    Not many pushers at Reno air races. (or anywhere else)

    Don't buy a "Breezy" for transportation.
     
  2. May 31, 2017 #82

    Eugene

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    " Decalage does not affect stability."

    Wow! I am for sure confused now! I need to go back to school and open some books!

    I'll be back later.....

    fullsizeoutput_c7d.jpg

    "Decalage is the difference in incidence between the wing and horizontal stabilizer (the wing should always have a larger incidence angle than the stabilizer). A large decalage angle generally makes the airplane more self-stabilizing in pitch, but it also causes larger trim changes when airspeed is increased or reduced."

    LONGITUDINAL STABILITY
    Longitudinal stability is pitch stability, or stability around the lateral axis of the airplane.
    To obtain longitudinal stability, airplanes are designed to be nose heavy when correctly loaded. The center of gravity is ahead of the center of pressure. This design feature is incorporated so that, in the event of engine failure, the airplane will assume a normal glide. It is because of this nose heavy characteristic that the airplane requires a tailplane. Its function is to resist this diving tendency. The tailplane is set at an angle of incidence that produces a negative lift and thereby, in effect, holds the tail down. In level, trimmed flight, the nose heavy tendency and the negative lift of the tailplane exactly balance each other.
    Two principal factors influence longitudinal stability: (1) size and position of the horizontal stabilizer, and (2) position of the center of gravity.
    The Horizontal Stabilizer
    The tail plane, or stabilizer, is placed on the tail end of a lever arm (the fuselage) to provide longitudinal stability. It may be quite small. However, being situated at the end of the lever arm, it has great leverage. When the angle of attack on the wings is increased by a disturbance, the center of pressure moves forward, tending to turn the nose of the airplane up and the tail down. The tailplane, moving down, meets the air at a greater angle of attack, obtains more lift and tends to restore the balance.
    On most airplanes, the stabilizer appears to be set at an angle of incidence that would produce an upward lift. It must, however, be remembered that the tailplane is in a position to be in the downwash from the wings. The air that strikes the stabilizer has already passed over the wings and been deflected slightly downward. The angle of the downwash is about half the angle of attack of the main airfoils. The proper angle of incidence of the stabilizer therefore is very important in order for it to be effective in its function.
     
    Last edited: Jun 1, 2017
  3. Jun 1, 2017 #83

    BBerson

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    As I see it, the forward fuselage is likely creating some unneeded positive lift from the wing up wash and requiring more tail down force. This could help explain the noted excess horizontal tail negative angle
    As I mentioned before, angling the forward fuselage more negative to align with the wing up wash (like gliders) might reduce drag.
    The aft fuselage is getting some negative lift but is a small area tube.
    Pushers have much more forward area than most tractor designs.
     
  4. Jun 1, 2017 #84

    fly2kads

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    I may be just a self-taught, armchair aerodynamicist, but I have some problems with the material you just quoted. I can see why you might be confused!

    No, an aircraft does not need a horizontal tail because of a "nose heavy effect." Conventional aircraft require a horizontal tail because the combination of wing, fuselage, and propulsion produces a nose-down pitching moment. Most airfoils have a chord-wise distribution of lift that is centered aft of the aerodynamic center. This results in a net force that, if unopposed, will rotate the wing forward. The fuselage will also generate a similar, but smaller, force. The thrust produced by the propeller can be either stabilizing or destabilizing, depending on location.

    In order to maintain trimmed, level flight, that total pitching moment must be balanced by some opposing force to achieve equilibrium. If your tail is aft, then that force must act downward. The size of the tail, the length of the tailboom, and the airfoil determine the maximum balancing force that can be achieved. The angle of attack of the tail and the deflection of the elevator determine the amount of lift generated at a given condition. As others have stated here, the angle of incidence of the tail is normally set such that the force required to counterbalance the pitching moment is achieved through angle of attack alone, leaving the elevator in trail for lowest drag. For most planes, this baseline trim position is set for cruise airspeeds and typical weight, since that is where you will be spending most of your time. As has been mentioned, the wing downwash will impact the angle of the oncoming wind seen by an aft tail. An aft tail may also be operating in "dirty" air that reduces the dynamic pressure seen by the tail. These last two factors may require an increase in tail incidence in order to achieve equilibrium.

    To reiterate what others have said, the angle of incidence is often set so the fuselage is in its lowest drag position at cruise. Other considerations include in-flight visibility and, especially for tri-gear aircraft, developing enough lift to rotate the aircraft during takeoff.

    The decalage is the difference in angles of incidence between the wing and tail.

    Now back up to the first paragraph...

    It is normal to have the wing at a higher angle than the tail, but not for the reason given. The wing and tail are producing forces in opposite directions, so it is natural for them to be angled differently. It is very important to prevent the tail from stalling, but that is very different than simply saying it has to be at a lower angle of attack than the wing. As to decalage making the aircraft "more self-stabilizing," I don't see it. That statement implies that decalage increases the lift curve slope of the tail, but that's not the case. The angle of incidence of the tail (and hence, the decalage) simply fixes the point along that lift curve where the tail is operating in the trimmed position.
     
    Last edited: Jun 1, 2017
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  5. Jun 1, 2017 #85

    BBerson

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    That was very well stated.
    I would also agree that sentence from the anonymous author was new to me (and nonsense, I think).

    He said:
    "This design feature is incorporated so that, In the event of engine failure, the airplane will assume a normal glide. It is because of this nose heavy characteristic that the airplane requires a tailplane."
     
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  6. Jun 1, 2017 #86

    fly2kads

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    Lest all of that semi-technical discussion become more unintelligible "French," here is a simple diagram:
    Lift_Curves_1.jpg
    This represents the lift curves of a hypothetical wing and stabilizer. These diagrams are used to show how much the lift changes as you change angle of attack.

    The wing is represented by the upper curve. Cambered wings will have zero lift at a small negative angle of attack. The curve progresses linearly up to close to the point of stall, where it drops off in some fashion. Let's say that this wing will need to support a weight of 1000 lbs. at an airspeed of 100 knots. Using the size of the wing, the aspect ratio, etc. we can use the standard lift equation to find the lift coefficient needed. We can then use this graph to find the AoA that will produce that lift coefficient. This is Point A on the graph. Knowing that, we can then use the factors described above to set the angle of incidence to get the fuselage at the angle we want. That's our wing incidence.

    The horizontal tail is represented by the lower curve. Most light aircraft use either a flat plate or a symmetrical airfoil, so the curve will pass through the graph's origin. The steps to find the required lift coefficient are more involved for a horizontal tail, for the reasons described above. Once you figure that out, though, you use the curve in the same manner. Let's say you figure out that the tail must produce 150 lbs. of down force. You use your much larger set of variables to find the lift coefficient that will produce that 150 lbs. This becomes Point B on the graph. Because the tail is operating under the influence of the wing's wake, you need to figure out the required angle of attack based on the local flow, not the fuselage reference.

    As you can see (hopefully), the angles of the wing and tail are set independently based on their own requirements. The resulting difference in angles, the decalage, is an arbitrary value. It is what it is, and is not a design variable on its own. (The decalage between the two wings of a biplane is a design variable, but that's a whole 'nother kettle of fish.)

    ______
    It looks like you just have a pretty draggy airframe, overall. The uncowled engine will be a big drag producer, of course. The rapid convergence of the fuselage behind the cockpit may produce separated flow. This is likely compounded by the flow underneath the wing...I wouldn't be surprised to see some significant flow separation at the wing/fuselage intersection. My hunch is that without some significant clean-up, you've found the limits of this particular design.
     

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  7. Jun 1, 2017 #87

    Eugene

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    Excellent!!! Absolutely excellent!!! Thank you!

    This is what I call - plain English. That means, that everything else I was reading before was in French?

    Now, I will try to summarize what I have learned from you guys. And I will try to do it in plain English. Please let me know if I will start talking French, because it is important for me to understand and get this right.

    My airplane if would be originally designed for 100hp - Will be a little longer, horizontal stabilizer Will be little larger, or will be higher. As a result it's incidents Will be much more normal. And if designer would have chosen tractor configuration - that angle Will be even smaller. So, I should stop worrying about it.

    Going back to angle of attack. If wing loading and power loading were they need to be - I am left with drag or unbelievable propeller inefficiency. So is rest of Skyboys.

    I tried very hard to find out position of my elevator and trim tab in cruise. As far as I can tell, all three surfaces almost in line. So, there is no conflict.

    Doing my research, I found that on production airplanes after WW2 every time They change model for bigger engine they usually change Incidence on main wing. Shrinking decalage. I was hoping, since my aircraft was designed for 50hp to do the same thing, since I have 100hp.
    Looks like I was wrong again!

    Thank you. This was very educational!


    fullsizeoutput_73b.jpg
     
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  8. Jun 1, 2017 #88

    Dana

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    What's the weight and wing area? With that it's easy to calculate the AOA at any speed.

    It takes 8X the HP to go twice as fast. Put another way, doubling the HP will only get you a 25% speed increase. But doubling the HP will more than double the climb performance.

    Dana
     
  9. Jun 1, 2017 #89

    Eugene

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    We do know wing area = 138 sq. ft.
    We do know weight = 950 - 1150 lb.
    We do know speed at 5000 rpm = 82 mph at 4000 ft
    We do know AOA = about 6 to the cord line

    What we don't know is DRAG. Airplane looks relatively clean from outside.


    IMG_4106.jpg IMG_4107.jpg IMG_4108.jpg IMG_4109.jpg IMG_4110.jpg
     
  10. Jun 1, 2017 #90

    Dana

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    Something doesn't make sense. Those numbers work out to a lift coefficient of 0.54, which is at about 1° AOA for the 4412.

    Maximum L/D for the 4412 is around 5° AOA which works out to 64 mph, so the original designer may have set the wing incidence to that value for best performance with minimum power.

    Are you sure your airspeed indicator is accurate?

    Dana
     
  11. Jun 1, 2017 #91

    ragflyer

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  12. Jun 1, 2017 #92

    BBerson

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    I think there is some exaggeration going on here, because the claimed 45° glide angle is likely exaggerated as well.
    A more accurate method of measuring the wing angle in flight is needed.
    It must be in level flight, not climbing when the angle is measured.

    You could buy an Aircam with two engines and still not go any faster :gig:
     
  13. Jun 1, 2017 #93

    Eugene

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    Yes between airspeed indicator, Garmin 492 , For flight and iPhone we can be sure that speed is more or less correct . 25 years ago original designer had in mind 50 hp engine, P3 airfoil, and at least 200 pounds less weight.

    When they decided to install 100 hp engine and 4412 wing geometry of that airplane left unchanged .

    To American market with big engines they decided to make wing shorter. As far as I can tell, I do have two wings with 5 x 14 each .

    At low power settings airplane feels to relaxed and high angle of attack feels normal. At maximum power it feels like you're stuck in the middle between two surfaces that are fighting each other . Of course that is my imagination . I talked to a few guys with the same airplane and they do like my analogy of pushing plywood against the wind . Trim is very sensitive at that time An airplane is very sensitive to updraft . Makes it next to impossible to do perfectly level flight . Best I can tail with digital level that bottom of the wing is in between four and 6° . If you look at any picture online and try to hold a level to the bottom of the wing you will see about that number.
     
    Last edited: Jun 1, 2017
  14. Jun 1, 2017 #94

    ragflyer

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    Assuming 163 sq.ft. area (per wikipedia aircraft link) you get 0.37 lift coefficient and 4.3 degree AOA at 80 MPH. If you assume a 138 sq.ft. wing then the AOA is 5 deg and lift coefficient 0.42 assuming AUW of 1000lbs.
     
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  15. Jun 1, 2017 #95

    Eugene

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    So, if I understand you correctly - 5° for 138ft.² = correct angle of attack ? At 80 miles an hour

    Sorry, my brain is not working anymore ... IMG_5475.jpg
     
  16. Jun 1, 2017 #96

    ragflyer

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    Keep in mind the zero lift angle for the NACA 4412 is -4 deg. In other words it would be 5-4 = 1 degree with respect to the chord line.
     
  17. Jun 1, 2017 #97

    BJC

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    At 5 degrees alpha, 2-D Cl for a NACA 4412 is in the range of 0.8 to 1.0.

    Eugene is reporting the estimated angle between the flattish bottom of the wing and the horizon, so the actual alpha is greater that the angle that he sees. I didn't calculate any numbers, but even a Cl of 0.4 sounds low. What Reynolds number did you use, how did you calculate total wing lift, and how did you reduce 2-D lift for 3-D real world?

    It sounds to me like the wing is working really hard.


    BJC
     
  18. Jun 1, 2017 #98

    ragflyer

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    btw do you know for sure it is a 4415 airfoil? The 4415 does not exactly have a flat bottom.
     
  19. Jun 1, 2017 #99

    Eugene

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    It is for sure 4412 and you are correct it doesn't have flat bottom . But it is almost flat visually at least
     
  20. Jun 1, 2017 #100

    ragflyer

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    Yes took 3D effect into account. No the wing is not working very hard. Here is how you calculate:

    cl = (grossWt * 391)/(wingArea*airSpeed^2)

    this gives you 0.42 at 82 mph, 1000 lbs and 138 sq.ft area. The Cl tells you how hard the wing is working. For the NACA 4412 this (0.42) is in the middle of its low drag range; hence not working too hard.

    To calculate the AOA you first adjust the lift slope for 3D by liftSlope = 0.11*AR/(AR+2)

    I assumed AR is 7 based on the original skybox stats 168 sq Ft area and 34' span. This gives a lift slope of 0.086

    A CL of 0.42 then will requires 0.42/0.086 =4.9 deg of angle from the zero lift line of airfoil.

    The zero lift line for the NACA 4412 is at -4 deg ( you can see that from the 2D plot; it does not change for 3D).

    Hence AOA with respect to chord line is 0.9 deg
     

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