Had a day off of work while on travel in Utah and made an
appointment to get a hang glider lesson with Steve P. from wingsoverwasatch.com. Only have a few photos, but am
sharing what I have. Steve made a GoPro video shot from the left wing. High value from the morning's instruction!
I'm still in one piece after a whole morning of drills, short hops, and a
few longer hops. Max altitude was in the neighborhood of 20 feet.
Turning by weight shift is such an unnatural feeling and I got caught
twisting instead of shifting early on, resulting in my only crash
landing. Glad I was wearing a helmet because my head whacked the keel
tube as one wing tip caught on the flare and pulled my nose over 
and
into the dirt. No damage other than ego. Not having direct yaw control
was weird, making crosswind flying awkward and left me wanting for a
rudder to help move the nose around. It was a lot like my last hand launch glider contest when my rudder pushrod broke and I had to tape the rudder at
neutral. Maybe that's why it was harder than I expected to hold a
straight ground track line in the occasional crosswind. At any rate,
the cross breezes gave the opportunity to try some gentle turns. If my
shoulders weren't painfully turning black and blue from holding up the
wing, I would have kept working up the hill. I launched from at least
twice as high as the dunes of Kitty Hawk if that helps build a mental
image. The South Side hill is several times taller than the dunes and a
whole lot less soft :-)
Many thanks to Steve P. at wingsoverwasatch.com.
If you're ever in the Salt Lake City area with a free morning, give
these folks a call and enjoy their excellent training and gentle
encouragement.
All in all, very awesome to get back in the air and
definitely an incentive to pull Goat back off the back-burner (at least
back as a side-project to the kitchen, hehe). Thanks for sharing with me!
The promised video:
Sunday, June 24, 2012
Wings Over Wasatch
Tuesday, April 24, 2012
More ANSYS runs
Maximum stress is still in the leading edge tube between the jury strut attachment and the wing root pin joint. There is still some bowing of the main struts upward as the jury struts are pulled on by the main spar.
Do note that these deflections are listed as approximately 0.5 inch. With such a low-fidelity FEA, I'm expecting somewhere between 3-12 inches of actual deflection during the load test, but the point is to say that the shape shown in the analysis is scaled to be visible. ANSYS has a true-load deflection option and it shows essentially no discernible deflection. Beware of length scaling. To answer a few other questions:
- I am too far into the build to change the main strut attachment location to the wing spars. So, what you see in the FEA is what I have to work with. Potentially I can add additional sleeve length or inner sleeves to the jury strut location or also change where the jury strut attaches to the main strut.
- I am not using Schrenk's approximation or an elliptical load distribution. Instead, this loading comes from the AVL model, which is likely better than Schrenk for arbitrary wing planforms. See a previous post to learn more about the AVL work.
- I have so much margin left in the wing structure near the strut attachment that I'm not going to run the asymmetric aileron loading case. The FEA in my mind was more concerned with showing how much margin to yield or ultimate load that I had, rather than doing the detail-work checking for aeroelastic effects.
I should probably add that I'm narrowing in on a decision to load-test to 4G's. According to the FEA, this should be well below where I'd anticipate any yielding and is more than I plan on pulling (no loops, heh). A cohort at work has even suggested it will be difficult to pull that many G's in the airchair, considering the high drag count and low-performance airfoil. Self-limiting is good in this case :-)
Sunday, April 15, 2012
Detail work on structures analysis
Among a bunch of great questions/suggestions/feedback, I took a look at the fore/aft distribution of forces on the two spar tubes. The previous analysis ran a simple 50/50 split of the wing total forces to get things started. What is better? AVL can help provide the answer.
Xle = 3.16670 Ave. Chord = 5.0000 Incidence = 1.7500 deg
Yle = -0.59947 Strip Width = 1.20000 Strip Area = 5.999976
Zle = 4.03135 Strip Dihed. = -2.9939
cl = 1.56703 cd = 0.73643 cdv = 0.67992
cn = 1.66386 ca = 0.47903 cnc = 7.78287 wake dnwsh = 0.10109
cmLE= -0.55324 cm c/4 = -0.13727
I X Y Z DX Slope dCp
241 3.19462 -0.59918 4.03134 0.17282 0.28424 5.05604
242 3.41428 -0.59918 4.03134 0.37261 0.14534 2.82108
243 3.83407 -0.59918 4.03134 0.54628 0.08312 2.09844
244 4.41670 -0.59918 4.03134 0.67141 0.03037 1.77600
245 5.11040 -0.59918 4.03134 0.73689 -0.01091 1.54034
246 5.85353 -0.59918 4.03134 0.73689 -0.04508 1.32883
247 6.58005 -0.59918 4.03134 0.67141 -0.06874 1.10233
248 7.22542 -0.59918 4.03134 0.54628 -0.08517 0.86729
249 7.73230 -0.59918 4.03134 0.37261 -0.09597 0.62162
250 8.05563 -0.59918 4.03134 0.17282 -0.10036 0.35427
The X position does not start at zero because I use the nose of the glider as the origin location. An X position of 3.1667 is the LE at the root if you were curious. Plotting the X position vs the dCp value gives the distribution of pressure along the airfoil chord.
Now some simple statics can give the force distribution between the LE and TE. Assuming each CP acts like a force on a beam, the component of load carried by the LE and TE can be found by multiplying the force by a ratio of distance to the beam length. The beam looks like this:
Note from the AVL plot that the root airfoils are working harder than the tip airfoils, with regard to a CP peak at the LE vs more evenly distributed along the chord. Using the force distribution at the root (75/25) should provide a conservative estimate of the LE spar loading out toward the tips. I may also run 100% on the LE just to see what that looks like.
Friday, April 6, 2012
Structural analysis preliminary results
Alright, I stayed 1.5hr after work and made a huge dent in the structural work. Thanks to Trent for helping me get set up on ANSYS. Here we go...
The compression ribs, cables, and jury struts were treated as LINK180 elements with the cables as tension-only elements. Note that the cables connect across the rib bays as they connect in the real aircraft.
Load was split up 50/50 between the LE and TE, which may not be exactly fair, but is a good first pass. 14 discretized loading locations represented the full half-span of load, some 938lb (conservatively including interpolation error) spread along the rib connection locations.
The LE and TE constraints are pins, fixed in x/y/z space and allowed to transmit moments only in the drag-direction (z if you're counting). Similarly, the strut connection to the fuselage carry-through tube is set up the same way. Furthermore, the strut to LE connection can only transfer forces, no moments, which reasonably represents the pinned connection out there.
Do disregard the stray line in the above screenshot; it was taken early while I was still cleaning up the element numbering.
Looking at von-Mies contour plots by element starts to quickly illuminate some unexpected results! First and foremost, the cantilever portion of the wing tip isn't bowing up. To the contrary, there is hardly much deflection out there at all, and it's ironically down. That's explained by the significant loading on the inner portion of the wing bowing UP. Yup, for a positive loading on the wing, the middle of the wing bows up.
This view also shows the minimum stress is at the leading edge attach point.
I'll be taking another look at the Euler buckling criteria to ensure the relatively smaller strut tubes aren't being pushed to the limits.
Oh also, check out the maximum stress: 8500psi. According to matweb and wikipedia, the yield stress of 6061-T6 is approximately 35,000psi. The fatigue limit is 14,000psi. Meaning, occasional pulls of 6g aren't appreciably fatiguing the structure and we're looking at a safety factor of 4 prior to yield, let alone ultimate. Take this with a grain of salt ... FEA is an approximation. But seeing such large margins of safety is really making me smile :-D
There are several uncompleted tasks / unanswered questions, namely:
- more fidelity in how the load is split between the leading and trailing edges,
- better representation of the spar to spar joint with inner sleeve, now that we know that's an important location,
- what is an appropriate sleeve arrangement for the jury-strut attach point? This ventures from analysis into design changes.
- check into buckling.
Disclaimer: This is not considered structural advice; I consider this post (and all) my personal notes and rhetorical discussion. This analysis is only for my implementation of a design I found online and does not constitute engineering advice for your project.
Cable
I was short some thick washers, so thought it would be a good time to get the cables: 3/32" Stainless Steel 7x7 control cable. I estimated 120ft was 15% more than I needed. This stuff is pretty hefty... Aircraft Spruce lists 920lb strength. The structural analysis I'm working on includes a cable member so we can get an idea of the loading in the internal bracing.
The structural analysis is still ongoing. It's slow working for ten minutes at a time before/after work. The simple quick tests that my coworker has shown me look absolutely fantastic, so I'm hopeful for good results. My plan, I believe, is to run the structure that I'm building to:
1. Find the loading for yield/ultimate.
2. Figure out the first component to fail (which joint is most stressed).
3. Decide on any changes and rerun the analysis.
4. Get tip deflection numbers to anticipate during the load test (if they are within 15% and I would feel pretty awesome about the analysis).
5. Based on the analysis and load test, make a V-n diagram that I can hold myself to... this should also help justify the "has not been known to fly faster than 45mph" Vne.
What I'm really looking for in the structural analysis is a warm-fuzzy that the structure can carry me without breaking. Beyond that is gravy.
Monday, April 2, 2012
Encouragement video
The Yando Goat is doing some more flying this autumn (southern hemisphere) season. I thought this would be worthy of posting as encouragement for myself, as well as showing the structure actually has flown. If you're interested in the structure of the Yando Goat, join the airchair yahoogroup.
Saturday, March 24, 2012
Wing loading at +6g
Starting the math...
Using the AVL model I gave in a previous post, we'd like the wing loading for say a 6g pull-up maneuver. Say CLmax is 1.5 and load factor is 6 (using gross mass of 305 lb-mass), giving a necessary condition of 80 ft/s (54mph) and -21 deg elevator. At this test case, the trefftz plane shows:
# Chordwise = 10 # Spanwise = 10 First strip = 1
Surface area = 179.309448 Ave. chord = 4.974011
CLsurf = 1.50090 Clsurf = 0.00000
CYsurf = 0.00000 Cmsurf = -0.01353
CDsurf = 0.50339 Cnsurf = 0.00000
CDisurf = 0.07366 CDvsurf = 0.42973
Forces referred to Ssurf, Cave about hinge axis thru LE
CLsurf = 1.32253 CDsurf = 0.44357
Deflect =
Strip Forces referred to Strip Area, Chord
j Yle Chord Area c cl ai cl_norm cl cd cdv cm_c/4 cm_LE C.P.x/c
1 -17.8428 4.2515 2.6609 1.5916 0.2041 0.4198 0.3729 0.0428 0.0047 -0.0402 -0.1256 0.359
2 -15.8461 5.0000 18.1066 5.0004 0.1971 1.1179 0.9994 0.1290 0.0453 -0.1015 -0.3534 0.352
3 -11.1640 5.0000 28.2017 6.7405 0.1283 1.4925 1.3546 0.4104 0.3418 -0.1260 -0.4765 0.343
4 -6.2763 5.0000 19.2959 7.3489 0.1029 1.6046 1.4793 0.5793 0.5221 -0.1328 -0.5206 0.340
5 -2.1627 5.0000 21.3896 7.6420 0.0951 1.6471 1.5386 0.6810 0.6267 -0.1366 -0.5431 0.339
6 2.1627 5.0000 21.3897 7.6419 0.0934 1.6471 1.5385 0.6810 0.6267 -0.1366 -0.5431 0.339
7 6.2763 5.0000 19.2959 7.3489 0.1029 1.6046 1.4793 0.5793 0.5221 -0.1328 -0.5206 0.340
8 11.1640 5.0000 28.2017 6.7405 0.1283 1.4925 1.3546 0.4104 0.3418 -0.1260 -0.4765 0.343
9 15.8461 5.0000 18.1066 5.0004 0.1971 1.1179 0.9994 0.1290 0.0453 -0.1015 -0.3534 0.352
10 17.8428 4.2515 2.6609 1.5916 0.2041 0.4198 0.3729 0.0428 0.0047 -0.0402 -0.1256 0.359
The column labeled cl_norm helps us get to the z-axis force at the local section. Backing out through the standard normalization CL=F_z/(Q * Sref) where Q = 0.5 * rho * V^2 ... section 1 (a tip) has the following loading:
Dynamic pressure:
q = 0.5 * rho * V^2 = 0.5 * 0.002378 slug/ft^3 * (80.7 ft/s)^2 = 7.743 slug/ft/s^2
Section force:
F_z = cl_norm * q * S_section = 0.4198 * 7.743 slug/ft/s^2 * 2.6609 ft^2
= 8.65 slug*ft/s^2 = 8.65 lb-force
(this is where I hate English units and love SI units ... kg makes more sense than lb vs lb-mass vs lb-force)
Carrying out these calculations on the other sections gives the following section forces:
| section # | Yle (ft) | F_z (lb) |
| 1 | -17.84 | 8.65 |
| 2 | -15.85 | 156.74 |
| 3 | -11.16 | 325.93 |
| 4 | -6.28 | 239.75 |
| 5 | -2.16 | 272.80 |
| 6 | 2.16 | 272.81 |
| 7 | 6.28 | 239.75 |
| 8 | 11.16 | 325.93 |
| 9 | 15.85 | 156.74 |
| 10 | 17.84 | 8.65 |
The total sum of all F_z is 2007.74 lb, which is more than 6G's * 305lb, but reflects the additional loading due to dihedral angle (some aero force is pointing in the y-direction too). This tells that the outer two panels, from 18 to 17.68ft and 17.68 to 14.02ft combined have to carry approximately 165lb at the 6G loading case. Now the strut actually joins my wing at 139in (11.58ft), so we really should change the panel locations to correspond better with what is beyond the strut attach point. Here is a top-down view of the 10 panel wing for reference.
Going back to the 50 panel wing we started with at the top of the post, the wing outside the strut carries 269.5lb, the wing between the strut and the jury attach carries 353.8lb, and the inside wing between the jury attach and the root carries 378.5lb. Do note that these forces are not centered on the panels, especially the tip weight; rather, the lift distribution governs the location. That we can figure out from the lift distribution too ... but will wait for another day.
Here is the full 50 panel wing to get a better idea of the number of strips:
I drew up the spars in Solidworks and will run them through an FEA analysis to get an idea (*idea) of the stress distribution. Notably, there is area between the sleeves taken up by several wraps of electrical tape (as noted on Sandlin's drawings). The FEA will be assuming the walls touch and do not slip, which is not a conservative assumption. I'll chat with some folks at work to get some additional input.
Please, if you are a reader and note a mistake in my math or assumptions somewhere, please please let me know. I will not discard your input and would be happy to spend time working with you to get this right. Your life-saving thoughts are most appreciated :-)
Friday, March 23, 2012
Sleeves
Decided on placement for the outer rib: 139" from the spar tube end. I also added 2" to the inner spar tube, so the inner rib is at 70" from the spar tube end. The compression ribs are spaced evenly in the remaining distance. These placements weren't scientific, rather admission to myself that other Goats built to these drawings are flying and doing fine. If there is a problem with the dimensions I've chosen, I'm counting that it'll show during the load test.
Need to start thinking about making strut connection brackets on the cabane end and also how to get the final strut lengths (incl matching washout angles between the two wings).
Monday, March 19, 2012
AVL model of Goat
As promised, the Goat AVL model. Units are in feet, seconds, and slugs. No guarantees this matches anyone's as-built Goat. I do plan on measuring my own creation after it comes together, but there is likely little to be gained with improved fidelity of this model compared to what you can learn from what is posted below.
Mass file:
# GOAT
# Mass & Inertia breakdown
#
# xyz is location of item's own CG
# Ixx.. are item's inertia about item's own CG
#
# x back
# y right
# z up
#
# x,y,z system here must have origin
# at same location as AVL input file
#
Lunit = 1.0 ft
Munit = 1.0 slug
Tunit = 1.0 s
g = 32.18
rho = 0.002378
# mass x y z Ixx Iyy Izz [ Ixy Ixz Iyz ]
9.5 4.27 0 3.0 1. 1. 1.
Geometry file:
#***********************************************************************************
# AVL dataset for GOAT Airchair model
# Generated by AVL Model Editor on 2 Jan 2012
#***********************************************************************************
GOAT Airchair
#Mach
0.1234
#IYsym IZsym Zsym
0 0 0.0000
#Sref Cref Bref
158.0000 5.0000 36.0000
#***********************************************************************************
# AVL Axes:
# +X downstream
# +Y out right wing
# +Z up
#***********************************************************************************
#Xref Yref Zref
4.6000 0.0000 3.0000
#CDp
0.0300
#***********************************************************************************
# Surfaces
#***********************************************************************************
#=======================================wing========================================
SURFACE
wing
#Nchord Cspace Nspan Sspace
10 1.0000 50 1.0000
SCALE
#sX sY sZ
1.0000 1.0000 1.0000
TRANSLATE
#dX dY dZ
3.1667 0.0000 4.0000
ANGLE
#Ainc
1.7500
#==================================wing section 1===================================
SECTION
#Xle Yle Zle Chord Angle
0.2500 -18.0000 0.9420 4.0000 0.0000
AFILE
#Airfoil definition
clarkysm.dat
CLAF
#CLaf = CLalpha / (2 * pi)
0.9844
CDCL
#CL1 CD1 CL2 CD2 CL3 CD3
-0.52080 0.00550 0.28900 0.00527 0.80980 0.00550
#==================================wing section 2===================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 -17.3750 0.9090 5.0000 0.0000
AFILE
#Airfoil definition
clarkysm.dat
CLAF
#CLaf = CLalpha / (2 * pi)
0.9844
CDCL
#CL1 CD1 CL2 CD2 CL3 CD3
-0.52080 0.00550 0.28900 0.00527 0.80980 0.00550
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
aileron 1.0000 0.7500 0.0000 0.0000 0.0000 1
#==================================wing section 3===================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 -8.1260 0.4250 5.0000 0.0000
AFILE
#Airfoil definition
clarkysm.dat
CLAF
#CLaf = CLalpha / (2 * pi)
0.9844
CDCL
#CL1 CD1 CL2 CD2 CL3 CD3
-0.52080 0.00550 0.28900 0.00527 0.80980 0.00550
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
aileron 1.0000 0.7500 0.0000 0.0000 0.0000 1
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
flap 1.0000 0.7500 0.0000 0.0000 0.0000 1
#==================================wing section 4===================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 0.0000 0.0000 5.0000 0.0000
AFILE
#Airfoil definition
clarkysm.dat
CLAF
#CLaf = CLalpha / (2 * pi)
0.9844
CDCL
#CL1 CD1 CL2 CD2 CL3 CD3
-0.52080 0.00550 0.28900 0.00527 0.80980 0.00550
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
flap 1.0000 0.7500 0.0000 0.0000 0.0000 1
#==================================wing section 5===================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 8.1260 0.4250 5.0000 0.0000
AFILE
#Airfoil definition
clarkysm.dat
CLAF
#CLaf = CLalpha / (2 * pi)
0.9844
CDCL
#CL1 CD1 CL2 CD2 CL3 CD3
-0.52080 0.00550 0.28900 0.00527 0.80980 0.00550
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
aileron 1.0000 0.7500 0.0000 0.0000 0.0000 -1
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
flap 1.0000 0.7500 0.0000 0.0000 0.0000 1
#==================================wing section 6===================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 17.3750 0.9090 5.0000 0.0000
AFILE
#Airfoil definition
clarkysm.dat
CLAF
#CLaf = CLalpha / (2 * pi)
0.9844
CDCL
#CL1 CD1 CL2 CD2 CL3 CD3
-0.52080 0.00550 0.28900 0.00527 0.80980 0.00550
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
aileron 1.0000 0.7500 0.0000 0.0000 0.0000 -1
#==================================wing section 7===================================
SECTION
#Xle Yle Zle Chord Angle
0.2500 18.0000 0.9420 4.0000 0.0000
AFILE
#Airfoil definition
clarkysm.dat
CLAF
#CLaf = CLalpha / (2 * pi)
0.9844
CDCL
#CL1 CD1 CL2 CD2 CL3 CD3
-0.52080 0.00550 0.28900 0.00527 0.80980 0.00550
#====================================Horizontal=====================================
SURFACE
Horizontal
#Nchord Cspace Nspan Sspace
12 2.0000 15 1.0000
SCALE
#sX sY sZ
1.0000 1.0000 1.0000
TRANSLATE
#dX dY dZ
13.3000 0.0000 4.3794
ANGLE
#Ainc
-3.5000
#===============================Horizontal section 1================================
SECTION
#Xle Yle Zle Chord Angle
0.5000 -4.0000 0.0000 2.0000 0.0000
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
elevator 1.0000 0.5000 0.0000 0.0000 0.0000 1
#===============================Horizontal section 2================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 -3.0000 0.0000 3.0000 0.0000
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
elevator 1.0000 0.5000 0.0000 0.0000 0.0000 1
#===============================Horizontal section 3================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 0.0000 0.0000 3.0000 0.0000
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
elevator 1.0000 0.5000 0.0000 0.0000 0.0000 1
#===============================Horizontal section 4================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 3.0000 0.0000 3.0000 0.0000
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
elevator 1.0000 0.5000 0.0000 0.0000 0.0000 1
#===============================Horizontal section 5================================
SECTION
#Xle Yle Zle Chord Angle
0.5000 4.0000 0.0000 2.0000 0.0000
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
elevator 1.0000 0.5000 0.0000 0.0000 0.0000 1
#=====================================vertical======================================
SURFACE
vertical
#Nchord Cspace Nspan Sspace
12 2.0000 32 0.0000
SCALE
#sX sY sZ
1.0000 1.0000 1.0000
TRANSLATE
#dX dY dZ
14.8300 0.0000 2.0257
ANGLE
#Ainc
0.0000
#================================vertical section 1=================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 0.0000 0.0000 0.7500 0.0000
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
rudder 1.0000 0.0100 0.0000 0.0000 0.0000 1
#================================vertical section 2=================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 0.0000 0.7500 1.6000 0.0000
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
rudder 1.0000 0.0100 0.0000 0.0000 0.0000 1
#================================vertical section 3=================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 0.0000 2.2080 2.1670 0.0000
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
rudder 1.0000 0.0100 0.0000 0.0000 0.0000 1
#================================vertical section 4=================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 0.0000 2.9580 2.1670 0.0000
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
rudder 1.0000 0.0100 0.0000 0.0000 0.0000 1
#================================vertical section 5=================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 0.0000 3.7080 1.0000 0.0000
CONTROL
#label gain Xhinge Xhvec Yhvec Zhvec SgnDup
rudder 1.0000 0.0100 0.0000 0.0000 0.0000 1
#=======================================keel========================================
SURFACE
keel
#Nchord Cspace Nspan Sspace
12 2.0000 20 0.0000
SCALE
#sX sY sZ
1.0000 1.0000 1.0000
TRANSLATE
#dX dY dZ
6.9167 0.0000 4.0000
ANGLE
#Ainc
0.0000
#==================================keel section 1===================================
SECTION
#Xle Yle Zle Chord Angle
7.8167 0.0000 -1.9740 0.1000 0.0000
#==================================keel section 2===================================
SECTION
#Xle Yle Zle Chord Angle
0.0000 0.0000 0.0000 7.9167 0.0000
#==================================keel section 3===================================
SECTION
#Xle Yle Zle Chord Angle
7.8167 0.0000 0.4840 0.1000 0.0000
Clark-Y airfoil (seemed a close enough approximation):
ClarkY_smoothed
1.000000 -0.000599
0.996087 -0.000743
0.989493 -0.000985
0.981993 -0.001261
0.973670 -0.001566
0.964656 -0.001897
0.955124 -0.002247
0.945264 -0.002609
0.935221 -0.002978
0.925064 -0.003351
0.914833 -0.003727
0.904561 -0.004104
0.894266 -0.004482
0.883958 -0.004861
0.873642 -0.005240
0.863322 -0.005619
0.853000 -0.005998
0.842676 -0.006377
0.832351 -0.006756
0.822026 -0.007135
0.811701 -0.007515
0.801375 -0.007894
0.791050 -0.008273
0.780724 -0.008652
0.770399 -0.009031
0.760073 -0.009411
0.749747 -0.009790
0.739422 -0.010169
0.729096 -0.010548
0.718771 -0.010927
0.708446 -0.011307
0.698120 -0.011686
0.687795 -0.012065
0.677470 -0.012444
0.667145 -0.012824
0.656820 -0.013203
0.646495 -0.013582
0.636169 -0.013961
0.625844 -0.014341
0.615519 -0.014720
0.605194 -0.015099
0.594869 -0.015478
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