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
Monday, March 19, 2012
AVL model of Goat
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2 comments:
hey i am a aerodynamics student... it was wonderful to see your work ..specially on AVL.. cud you also create a article explaining how to start working on AVL and if u have can u also upload or send me a copy of AVL editor...u would have got my email with this comment..
Hi Vishal,
Glad you like the AVL work. I'm happy to help push some more engineering data for Goat to the public.
There are several tutorials and articles about AVL if you do some Google searching. The AVL help documentation contains a lot of good info. I'll try to keep posting the run files as I use the model.
AVL is free and can be downloaded from: http://web.mit.edu/drela/Public/web/avl/
AVL Editor is not freeware, but does the exact same function as AVL with a prettier interface. It comes from Cloud Cap Technologies (http://www.cloudcaptech.com/). Who knows, if you ask nicely, perhaps they'll send you a copy.
You might try XFLR5 as an alternate freeware aerodynamics program. A lot of the RC glider guys are heavily using this software for aircraft design and optimization.
Dan
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