Putting a tapered wing on an RV

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rv7charlie

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I think the tail on the 7 was more than just reduced inventory. The 6”s vertical has always not been not as effective as the same on the 4. The wider fuselage needed more vertical. That the 6 exploded the RV world was the magic combination. They were not going to not miss out and not have an improved side by side. The 4 and 8 are the best dream makers; the 6 and 7 are the best reality makers.
The -8 added a counterbalance to the rudder (missing on the -4 & -6); the -7 originally used the -8 rudder but subsequent spin testing on the wider span -7 left them 'uncomfortable' with spin recovery, so they started shipping the larger -9 rudder with -7 kits.

When you say 'vertical', do you mean the stabilizer? I owned -4s for about 25 years, and I'm about 2 years into owning a -6. Never thought to measure the -4 & compare it to the -6, but I haven't noticed any significant differences between the -4 & -6; not much more variation than between the two -4s I owned. But I might not be tuned up enough as a pilot to detect the issues. :)
 

TFF

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You said you can tell, just not significant. Some may complain and some won’t. I was not saying wooo. Stay away. Crappy airplane. It’s a grade of what is better. The closer the tolerance harder to grade. If I was to go out and buy a flying RV with my budget, it would have to be a 6, because it will do my job the best overall, but I would want a 4. It won’t do the same job though.
 

davefried

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Directional stability

I did a static stability analysis to determine if the vertical stabilizer remained suitable. The big driver is the tail contribution and that has a volume coefficient impacted by the change in wing planform.

VT volume = (Sv/Sw)*(lv/bw)

The original planform coefficient (0.0456) was greater that the tapered planform coefficient (0.0446) suggesting a reduction in effectiveness. The desired value of the stability derivative has a span term in the denominator of the equation so it was reduced as well. In the end the tapered wing stability relative to it’s desired stability was greater than that for the original wing so I left the fin alone. The analysis held for rudder fixed and free.

Having said that, I have noticed the occasional yaw oscillation in light chop. I have heard similar comments for the RV-6 series. Perhaps that is part of the reason why the 7 got a larger fin.
 

Toobuilder

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Somewhat related to this discussion is the fact that I'm planning to do a new vertical tail on the Rocket at some point. The reason is because mine was built with some "parts bin engineering" and the guy stuck an RV-8 tail on it. The -8 tail is quite a bit bigger than the -4 and the latter is more than adequate on the Rocket. By definition, the shorter wings, longer fuselage and raised turtledeck on the Rocket points to a tail even smaller than the -4 for similar effectiveness. The larger MMoI of the Rocket in a spin is one area that would point to a larger tail, but the -4 unit has proven adequate in that realm as well.

The bigger -8 tail costs speed on a Rocket, and we can't have that.
 

Lendo

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Dave, On the issue of Flap and Aileron span Percentages, what are they on the exposed span (Tapered wing) and is this similar or the same as the Hershey Bar wing- in terms of Percentages. Also how much is the actual end Tip distance to the Aileron.
I have flown an RV4 and found the control balance to be good, i.e. good Aileron response without being overly sensitive, ( compared to some others) I assume that all RVs are similar and by the sound of it so is your Tapered Wing.
George
 

davefried

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Dave, On the issue of Flap and Aileron span Percentages, what are they on the exposed span (Tapered wing) and is this similar or the same as the Hershey Bar wing- in terms of Percentages. Also how much is the actual end Tip distance to the Aileron.
I have flown an RV4 and found the control balance to be good, i.e. good Aileron response without being overly sensitive, ( compared to some others) I assume that all RVs are similar and by the sound of it so is your Tapered Wing.
George
On the original wing, flaps start at Y = 20 and extend to 78 in. Ailerons extend from there to 126 in, 12 in from the tip. Percentage of wing area ahead of flaps is 42.0 and 34.8 for ailerons.

On the tapered wing, flaps start at Y = 20 and extend to 90 in. Ailerons extend from there to 140 in, 10 in from the tip. Percentage of wing area ahead of flaps is 46.7 and 33.3 for ailerons.

I had the opportunity to pick the split to adjust CLmax. I will post some details in a bit.
 
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User27

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Different wing on the RV-7, 25 foot span and 121 ft^2 wing area.
Actually the same wings used on a 7 and an 8, as the fuselage is wider the 7 has a slightly longer span.
Yes, the larger rudder was to improve spin recovery handling.
The 4 & 6 have non-counterweighted rudders where as the 8 & 7 have counter weights.
The 9 has significantly longer wings so needed greater fin/rudder area
 

Toobuilder

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Some pictures of the tapered carbon fiber wings we're making. This set will go on a pylon racing RV-6...
I'm guessing that particular RV-6 is bright red?

If so, Are you addressing the 7 knots lost with the addition of the RV-8 tail to that "red" -6?
 

BoKu

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I'm guessing that particular RV-6 is bright red?

If so, Are you addressing the 7 knots lost with the addition of the RV-8 tail to that "red" -6?
Yes indeed bright red outside, probably still a few feathers on the inside.

The tail, that's above my pay grade. You'll have to ask Bob or James about that.
 

Toobuilder

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Well, it was Bob who documented the substantial loss of speed on the course with the addition of the -8 tail and got me looking at mine with an eye towards more speed. Maybe after I get mine installed and tested I'll ship the drawings over to him.
 

davefried

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Stall speeds

I used Schrenk to determine the speed at which the wing would reach the airfoil’s maximum section Cl somewhere along the span. This is just the wing only and lift may continue to increase for a bit before the stall completely develops. There is a component of lift generated by the rest of the aircraft to consider but is not readily calculated.

A good example of published test data for the aircraft comes from the CAFE Foundation’s APR for the RV-6A. I compared their stall speeds with my analysis to get an estimate of a CLmax increment from wing to aircraft.

Original wing model 1600 lb Vs1 = 55.6 kt @ CL(wing)max = 1.38
APR Test adjusted to 1600 lb Vs1 = 50.5 kt cas @ CL(aircraft)max = 1.67 delta CLmax = 0.29

Applying the delta CLmax to the calculations for the tapered wing gave the following stall speeds (still 1600 lb),

Tapered wing CL(wing)max = 1.46 CL(aircraft)max = 1.75 Vs1 = 51.0 kt

Continuing with the APR report, the increment to CLmax due to flap deflection was calculated.

APR Test adjusted to 1600 lb Vs0 = 44.8 kt cas @ CL(aircraft)maxf = 2.12 delta CLmaxf = 0.45

The increment to CLmaxf is a function of the flapped wing area. The extent of the flaps on the tapered wing was increased raising the flapped wing area. As a result, the delta CLmaxf increased as well.

Sf/Sw = 0.42 on the original and 0.47 on the tapered wing, increasing delta CLmaxf to 0.53

Applying the delta CLmaxf to the calculations for the tapered wing gave the following stall speeds (still 1600 lb).

Tapered wing CL(aircraft)max = 1.75 Vs1 = 51.0 kt CL(aircraft)maxf = 2.30 Vs1 = 44.6 kt
APR Summary CL(aircraft)max = 1.67 Vs1 = 50.5 kt CL(aircraft)maxf = 2.12 Vs1 = 44.8 kt
 

Fenix

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Since I am trying to learn load distribution analysis to include working with the Schrenk formula I attempted to do Schrenk diagrams on the original RV-4 wing and my tapered planform. Attached are the diagrams I attempted. I'm not sure they are correct, or in what all ways they are used.

One thing I have read is that if the Cla line does not have its highest point at the root then tip stall could be a problem. The attached diagrams do not account for any washout (I don't know if you can work that into a Schrenk diagram but I don't know how to) and you can see the Cla line is just a very small amount higher at about the 40% span point. I'm not sure if this minor increase would be a problem or not. Dave Fried (above) I believe commented that his similar tapered wing on an RV-6 had no washout and still hand good stall characteristics. I did put washout in mine, though I'm not sure it is necessary after seeing the attached diagram (and learning of Dave's experiences).

Anyway, the Schrenk diagrams are attached for any value or curiosity they may be to the group.
 

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wsimpso1

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I will keep saying it. Lift unavoidably MUST go to zero at the tips. Any lift distribution with lift at the tip is exaggerating the shear and bending moment, which you then design for. Any lift distribution with lift at the tip is also exaggerating the Cl curve as you go outward, which you then make (un-needed) changes to correct for it. Relying upon a bogus lift distribution to make fine detail changes is bogus too. Lift distribution based entirely upon planview area is putting you there. Please get beyond it. These math models are all for the dual purposes of estimating loads (for structural design) and estimating Cl distributions along the span (to anticipate and correct stall progression). They need to as accurately as possible reflect reality or you will just be putting in unneeded changes that otherwise hinder utility, speed and efficiency.

Lift distribution usually looks elliptical and is an excellent starting point. Going further, the pressure distribution along the span is what is usually really close to elliptical. If you want to be a philosopher about this, yes, the pressures are distributed over the area, so you can start with spanwise elliptical in one column of your spreadsheet, then in the next column put in the fraction of root chord you have at each station, and multiply the local elliptical by that fraction. This is modifying the lift outboard based upon plan view. This reduces total lift and lowers the curve of shear and moment vs spanwise position. Yes, you will have to adjust your calculated stall speed upwards to carry the weight of the airplane and the downforce from the tail. That will avoid finding out later that your actual stall speed is higher than you thought...

You can go further. If you washout the wing (for whatever reason), either by twist or foil choices at root and tip, you can figure out what the reduction in local Cl would be as you go outward, divide local Cl by base Cl, and multiply your local lift by that ratio. This too reduces the total lift and lowers the shear and moment curves.

Then you can do the downward deflected aileron and its adjusted Cl over the portion of the wing that is aileron influenced. The wing should remain intact when the aileron is deflected at Vne, so this is a load case you need to check for structure. How? Same way, figure out the ratio of Cldeflected divided by Clbase and multiply by that value. This increases shear and bending moment curves. You might be concerned about stall here too.

All of this drives one to say, but don't these reduce total lift? Hmm. Where your foil is clean (no control deflections) local Cl can not exceed stall Cl. And where your high lift devices are deflected, the local Cl can not exceed the stall Cl of that deflection. So when any part your local Cl exceeds Cl at stall, you need to scale back Cl everywhere. Reduce AOA everywhere. Easy way is to find the biggest delta between computed Cl (which we know is not actually available to the wing) and max CL, then divide that by your wing's lift slope (include correction for AR). That is the number of degrees you need to lower the AOA. Or if you are working only in Cl, just lower the entire Cl curve by that difference between computed and known max Cl...

I just added three columns to your spreadsheet to correct elliptical for taper, for twist, and then for aileron deflection. All make your math model more closely resemble what is actually happening, and none exaggerate the actual wing mechanical loads or tendency to tip stall.

Billski
 

davefried

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Fenix: Your wing is very close to mine. If you could post the following numbers I will run them through my Schrenk spreadsheet and post the result.

Chord at fuselage side
Spanwise location of above
Tip chord
Span
Twist

The Schrenk "approximation" is here and is easy to recreate. I have added a tweak that accounts for Reynolds number. http://www.lightaircraftassociation.co.uk/2010/Engineering/Design/schrenk approximation.pdf

Billski's post came through as I wrote this, I'll have a look.
 
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