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:

 

This plot shows the loading really rolls off at about 12.5ft half-span.  There are too many wing sections to easily look at the numbers, so let's drop down to 5 panels per wing half:

The last panel isn't loaded nearly as heavily as the root areas.  Putting some numbers to this, AVL outputs the coefficients of pressure for each wing section (staying with the 10 panel wing):

  Surface # 1     wing                                   
     # 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 :-)

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.

Finished cutting the LE and TE sleeves for mounting the ribs.  Lengthened the 18" sleeves to 21" for the main strut attach location.  Goat3 has a 36" sleeve here...

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).

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
     0.584544   -0.015856
     0.574219   -0.016235
     0.563894   -0.016614
     0.553569   -0.016993
     0.543245   -0.017372
     0.532921   -0.017751
     0.522597   -0.018131
     0.512273   -0.018510
     0.501950   -0.018890
     0.491627   -0.019270
     0.481303   -0.019651
     0.470979   -0.020031
     0.460657   -0.020411
     0.450335   -0.020791
     0.440014   -0.021170
     0.429694   -0.021549
     0.419376   -0.021927
     0.409059   -0.022304
     0.398741   -0.022680
     0.388423   -0.023055
     0.378104   -0.023429
     0.367786   -0.023804
     0.357472   -0.024179
     0.347163   -0.024555
     0.336858   -0.024933
     0.326559   -0.025313
     0.316265   -0.025696
     0.305979   -0.026082
     0.295698   -0.026471
     0.285425   -0.026863
     0.275166   -0.027253
     0.264931   -0.027635
     0.254727   -0.028007
     0.244557   -0.028364
     0.234423   -0.028702
     0.224324   -0.029017
     0.214261   -0.029306
     0.204237   -0.029566
     0.194253   -0.029792
     0.184314   -0.029980
     0.174423   -0.030126
     0.164581   -0.030225
     0.154791   -0.030274
     0.145059   -0.030267
     0.135388   -0.030201
     0.125786   -0.030073
     0.116278   -0.029880
     0.106908   -0.029619
     0.097696   -0.029289
     0.088662   -0.028893
     0.079924   -0.028456
     0.071698   -0.027993
     0.064136   -0.027473
     0.057257   -0.026869
     0.051022   -0.026175
     0.045385   -0.025400
     0.040296   -0.024573
     0.035694   -0.023728
     0.031549   -0.022915
     0.027857   -0.022173
     0.024595   -0.021471
     0.021705   -0.020756
     0.019122   -0.020001
     0.016797   -0.019193
     0.014699   -0.018330
     0.012803   -0.017411
     0.011086   -0.016436
     0.009528   -0.015417
     0.008109   -0.014371
     0.006816   -0.013311
     0.005643   -0.012242
     0.004584   -0.011164
     0.003635   -0.010081
     0.002792   -0.008993
     0.002055   -0.007901
     0.001423   -0.006805
     0.000897   -0.005709
     0.000481   -0.004612
     0.000187   -0.003511
     0.000008   -0.002415
    -0.000062   -0.001325
    -0.000023   -0.000239
     0.000127    0.000852
     0.000389    0.001963
     0.000757    0.003101
     0.001229    0.004261
     0.001806    0.005442
     0.002487    0.006649
     0.003276    0.007887
     0.004174    0.009162
     0.005187    0.010481
     0.006325    0.011855
     0.007596    0.013294
     0.009013    0.014816
     0.010588    0.016440
     0.012334    0.018190
     0.014259    0.020082
     0.016382    0.022109
     0.018725    0.024254
     0.021319    0.026494
     0.024202    0.028817
     0.027414    0.031217
     0.030993    0.033686
     0.034980    0.036213
     0.039406    0.038797
     0.044290    0.041432
     0.049642    0.044104
     0.055473    0.046797
     0.061789    0.049505
     0.068576    0.052219
     0.075787    0.054922
     0.083344    0.057596
     0.091150    0.060212
     0.099134    0.062734
     0.107269    0.065145
     0.115555    0.067446
     0.123982    0.069639
     0.132534    0.071723
     0.141187    0.073700
     0.149919    0.075567
     0.158710    0.077323
     0.167540    0.078964
     0.176396    0.080486
     0.185274    0.081884
     0.194181    0.083158
     0.203133    0.084308
     0.212152    0.085341
     0.221255    0.086264
     0.230447    0.087085
     0.239729    0.087811
     0.249104    0.088449
     0.258573    0.089007
     0.268137    0.089492
     0.277794    0.089914
     0.287538    0.090281
     0.297355    0.090601
     0.307215    0.090882
     0.317083    0.091123
     0.326943    0.091319
     0.336791    0.091469
     0.346624    0.091571
     0.356440    0.091621
     0.366237    0.091619
     0.376015    0.091561
     0.385776    0.091446
     0.395524    0.091272
     0.405266    0.091036
     0.415014    0.090740
     0.424773    0.090385
     0.434544    0.089972
     0.444328    0.089504
     0.454126    0.088981
     0.463938    0.088406
     0.473763    0.087780
     0.483602    0.087105
     0.493451    0.086382
     0.503309    0.085615
     0.513170    0.084803
     0.523031    0.083947
     0.532891    0.083047
     0.542749    0.082104
     0.552605    0.081117
     0.562459    0.080087
     0.572309    0.079013
     0.582158    0.077897
     0.592005    0.076737
     0.601852    0.075534
     0.611702    0.074288
     0.621558    0.072999
     0.631422    0.071670
     0.641292    0.070301
     0.651170    0.068893
     0.661055    0.067449
     0.670948    0.065968
     0.680849    0.064453
     0.690756    0.062905
     0.700668    0.061325
     0.710583    0.059716
     0.720497    0.058077
     0.730408    0.056409
     0.740317    0.054713
     0.750223    0.052988
     0.760127    0.051234
     0.770027    0.049452
     0.779924    0.047642
     0.789818    0.045804
     0.799710    0.043939
     0.809598    0.042046
     0.819485    0.040125
     0.829370    0.038178
     0.839253    0.036204
     0.849133    0.034204
     0.859011    0.032178
     0.868885    0.030128
     0.878753    0.028053
     0.888611    0.025956
     0.898452    0.023838
     0.908267    0.021702
     0.918036    0.019551
     0.927726    0.017393
     0.937292    0.015239
     0.946686    0.013098
     0.955816    0.010994
     0.964495    0.008976
     0.972529    0.007093
     0.979790    0.005383
     0.986192    0.003871
     0.990000    0.002969

Some wing progress

I needed a good weekend off from work ... busy week.  But it turned into some quiet and productive time spent on Goat.  It also helped that the weather in DC was sunny and around 75F for highs so the garage was comfortable.  Nice to be back outside.

Started with the wing at this point on Saturday morning.  Clamped a straight-edge to the wing root brackets on the right and then used a (pretty much) flat floor and a quick-square to make perpendicular marks for the wing tube retention rivets.

With the tubes in this position, it's also possible to make tangent lines atop the spar tubes using a long straight-edge.  Geometry is starting to talk again as it comes together.  Those first few marks were the hardest to sort out.

 I did find one tiny "error" in the drawings.  I'm putting error in quotes because it isn't huge, but did make me pause long enough to get it right... the compression rib between the spar tubes have the hole closest to the spar at different distances.  Sparing you the math, the LE hole is 0.25" from the end of the compression rib and 0.375" from the other end of the compression rib.  All the other drawings have holes at 0.375", so I'd recommend lengthening the LE brackets by 1/8".

 First time with a big mock-up of the wing half and struts.  Okay, so I've officially decided to go with the strutted version of Goat rather than the cable-braced version.  Not only do I think it looks cooler with struts, I've read enough suggestions from Sandlin and others the strutted version is likely less drag (provided the struts are faired of course).  I'm cool with less drag.

The choice to go strutted has some other structural implications.  My cabanes were set up with the cable attachments, so it'll take some reworking for strut attach brackets.

It's also starting to sink in how large this wing is going to be.  I took some pictures of the full 36' span and it seems gigantic!  It's so big, the pictures of the full span were either washed out or really dark.  Guess you'll have to wait for me to get the whole thing assembled for a full-span shot ;-)

Trailing edge wing root area.  Pretty neat seeing all the bolts and load paths here.  Busy. 

 Leading edge wing attach area.  Lots going on up here.  Can lose the upper bracket since struts don't need a king-post.  Every ounce counts.






Edit: Sandlin asked a couple questions about the wing load test and the AVL analysis.  First, I heard back from the load test poster who said it sadly wasn't a Goat wing.  I am definitely planning a load test of my own though, so we'll have some real numbers.  Second, the AVL analysis is primarily for stability and control analysis.  I haven't touched the model since, but will post it here for others to play with.