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Merge pull request #538 from suavecode/AVL_control_surface_updates
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Addition of control surface deflection
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@mclarke2
mclarke2 committed Nov 18, 2021
2 parents 112064f + 0ae3172 commit 6a341cc
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62 changes: 31 additions & 31 deletions regression/scripts/AVL/avl_files/body_axis_derivatives_case_01_01.txt
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Expand Up @@ -13,54 +13,54 @@

Run case: case_01_01

Alpha = -2.00000 pb/2V = -0.00000 p'b/2V = -0.00000
Alpha = -2.00000 pb/2V = 0.00000 p'b/2V = 0.00000
Beta = 0.00000 qc/2V = 0.00000
Mach = 0.050 rb/2V = -0.00000 r'b/2V = -0.00000
Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000

CXtot = 0.00049 Cltot = 0.00000 Cl'tot = 0.00000
CYtot = 0.00000 Cmtot = 0.25003
CZtot = 0.04804 Cntot = -0.00000 Cn'tot = -0.00000

CLtot = -0.04803
CDtot = 0.00119
CDvis = 0.00000 CDind = 0.0011874
CLff = -0.04796 CDff = 0.0008417 | Trefftz
CYff = 0.00000 e = 0.0922 | Plane
CXtot = 0.00047 Cltot = 0.00000 Cl'tot = 0.00000
CYtot = 0.00000 Cmtot = 0.25074
CZtot = 0.04717 Cntot = -0.00000 Cn'tot = -0.00000

CLtot = -0.04716
CDtot = 0.00118
CDvis = 0.00000 CDind = 0.00118
CLff = -0.04708 CDff = 0.00086 | Trefftz
CYff = 0.00000 e = 0.0873 | Plane

slat = 0.00000
flap = 0.00000
aileron = 0.00000
elevator = 0.00000

---------------------------------------------------------------

Geometry-axis derivatives...

axial vel. u sideslip vel. v normal vel. w
---------------- ---------------- ----------------
x force CX | CXu = -0.006897 CXv = -0.000000 CXw = -0.225567
y force CY | CYu = 0.000000 CYv = -0.813809 CYw = -0.000001
z force CZ | CZu = -0.101932 CZv = -0.000001 CZw = -5.672218
x mom. Cl | Clu = -0.000000 Clv = -0.173774 Clw = -0.000004
y mom. Cm | Cmu = 0.222183 Cmv = -0.000001 Cmw = -7.966332
z mom. Cn | Cnu = -0.000000 Cnv = 0.278184 Cnw = 0.000000
x force CX | CXu = -0.006879 CXv = -0.000000 CXw = -0.223754
y force CY | CYu = 0.000000 CYv = -0.813886 CYw = -0.000000
z force CZ | CZu = -0.103691 CZv = -0.000000 CZw = -5.672500
x mom. Cl | Clu = -0.000000 Clv = -0.173946 Clw = -0.000001
y mom. Cm | Cmu = 0.223552 Cmv = -0.000000 Cmw = -7.967508
z mom. Cn | Cnu = -0.000000 Cnv = 0.278218 Cnw = 0.000000

roll rate p pitch rate q yaw rate r
---------------- ---------------- ----------------
x force CX | CXp = -0.000000 CXq = -1.186104 CXr = 0.000000
y force CY | CYp = -0.309538 CYq = -0.000008 CYr = 0.977207
z force CZ | CZp = -0.000009 CZq = -27.108837 CZr = 0.000001
x mom. Cl | Clp = -0.484662 Clq = -0.000023 Clr = 0.165280
y mom. Cm | Cmp = -0.000016 Cmq = -77.724060 Cmr = 0.000001
z mom. Cn | Cnp = 0.138563 Cnq = 0.000004 Cnr = -0.513345
x force CX | CXp = -0.000000 CXq = -1.188567 CXr = 0.000000
y force CY | CYp = -0.309325 CYq = -0.000001 CYr = 0.977330
z force CZ | CZp = -0.000001 CZq = -27.113642 CZr = 0.000000
x mom. Cl | Clp = -0.484673 Clq = -0.000003 Clr = 0.165341
y mom. Cm | Cmp = -0.000002 Cmq = -77.741516 Cmr = 0.000000
z mom. Cn | Cnp = 0.138501 Cnq = 0.000001 Cnr = -0.513391

slat d1 flap d2 aileron d3 elevator d4
---------------- ---------------- ---------------- ----------------
x force CX | CXd1 = 0.000066 CXd2 = -0.000418 CXd3 = -0.000000 CXd4 = -0.000173
y force CY | CYd1 = 0.000000 CYd2 = -0.000000 CYd3 = 0.000359 CYd4 = -0.000000
z force CZ | CZd1 = 0.000583 CZd2 = -0.025135 CZd3 = 0.000000 CZd4 = -0.012031
x force CX | CXd1 = 0.000056 CXd2 = -0.000432 CXd3 = 0.000000 CXd4 = -0.000175
y force CY | CYd1 = 0.000000 CYd2 = -0.000000 CYd3 = 0.000361 CYd4 = -0.000000
z force CZ | CZd1 = 0.000583 CZd2 = -0.025135 CZd3 = 0.000000 CZd4 = -0.012033
x mom. Cl | Cld1 = 0.000000 Cld2 = -0.000000 Cld3 = 0.001927 Cld4 = -0.000000
y mom. Cm | Cmd1 = 0.000088 Cmd2 = -0.023284 Cmd3 = 0.000000 Cmd4 = -0.058272
z mom. Cn | Cnd1 = -0.000000 Cnd2 = -0.000000 Cnd3 = -0.000072 Cnd4 = 0.000000


y mom. Cm | Cmd1 = 0.000089 Cmd2 = -0.023281 Cmd3 = 0.000000 Cmd4 = -0.058282
z mom. Cn | Cnd1 = -0.000000 Cnd2 = -0.000000 Cnd3 = -0.000075 Cnd4 = 0.000000
62 changes: 31 additions & 31 deletions regression/scripts/AVL/avl_files/body_axis_derivatives_case_01_02.txt
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Expand Up @@ -13,54 +13,54 @@

Run case: case_01_02

Alpha = 0.00000 pb/2V = -0.00000 p'b/2V = -0.00000
Alpha = 0.00000 pb/2V = 0.00000 p'b/2V = 0.00000
Beta = 0.00000 qc/2V = 0.00000
Mach = 0.050 rb/2V = -0.00000 r'b/2V = -0.00000
Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000

CXtot = -0.00210 Cltot = -0.00000 Cl'tot = -0.00000
CYtot = 0.00000 Cmtot = -0.02877
CZtot = -0.15023 Cntot = -0.00000 Cn'tot = -0.00000

CLtot = 0.15023
CDtot = 0.00210
CDvis = 0.00000 CDind = 0.0020971
CLff = 0.15019 CDff = 0.0014556 | Trefftz
CYff = 0.00000 e = 0.5229 | Plane
CXtot = -0.00202 Cltot = -0.00000 Cl'tot = -0.00000
CYtot = -0.00000 Cmtot = -0.02810
CZtot = -0.15111 Cntot = -0.00000 Cn'tot = -0.00000

CLtot = 0.15111
CDtot = 0.00202
CDvis = 0.00000 CDind = 0.00202
CLff = 0.15108 CDff = 0.00149 | Trefftz
CYff = 0.00000 e = 0.5181 | Plane

slat = 0.00000
flap = 0.00000
aileron = 0.00000
elevator = 0.00000

---------------------------------------------------------------

Geometry-axis derivatives...

axial vel. u sideslip vel. v normal vel. w
---------------- ---------------- ----------------
x force CX | CXu = -0.004194 CXv = -0.000000 CXw = 0.077528
y force CY | CYu = 0.000000 CYv = -0.819873 CYw = -0.000000
z force CZ | CZu = -0.300457 CZv = -0.000001 CZw = -5.683236
x mom. Cl | Clu = -0.000000 Clv = -0.197950 Clw = -0.000004
y mom. Cm | Cmu = -0.057538 Cmv = -0.000002 Cmw = -8.014050
z mom. Cn | Cnu = -0.000000 Cnv = 0.276472 Cnw = -0.000000
x force CX | CXu = -0.004047 CXv = -0.000000 CXw = 0.081202
y force CY | CYu = -0.000000 CYv = -0.819924 CYw = -0.000000
z force CZ | CZu = -0.302224 CZv = -0.000000 CZw = -5.683418
x mom. Cl | Clu = -0.000000 Clv = -0.198123 Clw = -0.000001
y mom. Cm | Cmu = -0.056198 Cmv = -0.000000 Cmw = -8.014884
z mom. Cn | Cnu = -0.000000 Cnv = 0.276504 Cnw = -0.000000

roll rate p pitch rate q yaw rate r
---------------- ---------------- ----------------
x force CX | CXp = 0.000000 CXq = -0.196756 CXr = 0.000000
y force CY | CYp = -0.239545 CYq = -0.000007 CYr = 0.991879
z force CZ | CZp = -0.000009 CZq = -27.145086 CZr = 0.000001
x mom. Cl | Clp = -0.484302 Clq = -0.000023 Clr = 0.195371
y mom. Cm | Cmp = -0.000016 Cmq = -78.041107 Cmr = 0.000002
z mom. Cn | Cnp = 0.094390 Cnq = 0.000003 Cnr = -0.516802
x force CX | CXp = 0.000000 CXq = -0.195168 CXr = 0.000000
y force CY | CYp = -0.239303 CYq = -0.000001 CYr = 0.991999
z force CZ | CZp = -0.000001 CZq = -27.149084 CZr = 0.000000
x mom. Cl | Clp = -0.484317 Clq = -0.000003 Clr = 0.195433
y mom. Cm | Cmp = -0.000002 Cmq = -78.053085 Cmr = 0.000000
z mom. Cn | Cnp = 0.094324 Cnq = 0.000000 Cnr = -0.516846

slat d1 flap d2 aileron d3 elevator d4
---------------- ---------------- ---------------- ----------------
x force CX | CXd1 = -0.000193 CXd2 = -0.000034 CXd3 = 0.000000 CXd4 = -0.000068
y force CY | CYd1 = -0.000000 CYd2 = -0.000000 CYd3 = 0.000335 CYd4 = -0.000000
z force CZ | CZd1 = 0.000596 CZd2 = -0.025113 CZd3 = 0.000000 CZd4 = -0.012032
x force CX | CXd1 = -0.000204 CXd2 = -0.000045 CXd3 = 0.000000 CXd4 = -0.000069
y force CY | CYd1 = -0.000000 CYd2 = -0.000000 CYd3 = 0.000338 CYd4 = -0.000000
z force CZ | CZd1 = 0.000596 CZd2 = -0.025113 CZd3 = 0.000000 CZd4 = -0.012034
x mom. Cl | Cld1 = 0.000000 Cld2 = -0.000000 Cld3 = 0.001924 Cld4 = -0.000000
y mom. Cm | Cmd1 = 0.000110 Cmd2 = -0.023241 Cmd3 = 0.000000 Cmd4 = -0.058332
z mom. Cn | Cnd1 = 0.000000 Cnd2 = -0.000000 Cnd3 = -0.000038 Cnd4 = 0.000000


y mom. Cm | Cmd1 = 0.000110 Cmd2 = -0.023240 Cmd3 = 0.000000 Cmd4 = -0.058340
z mom. Cn | Cnd1 = 0.000000 Cnd2 = -0.000000 Cnd3 = -0.000041 Cnd4 = 0.000000
62 changes: 31 additions & 31 deletions regression/scripts/AVL/avl_files/body_axis_derivatives_case_01_03.txt
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Expand Up @@ -13,54 +13,54 @@

Run case: case_01_03

Alpha = 2.00000 pb/2V = -0.00000 p'b/2V = -0.00000
Alpha = 2.00000 pb/2V = 0.00000 p'b/2V = 0.00000
Beta = 0.00000 qc/2V = 0.00000
Mach = 0.050 rb/2V = -0.00000 r'b/2V = -0.00000
Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000

CXtot = 0.00590 Cltot = -0.00000 Cl'tot = -0.00000
CYtot = -0.00000 Cmtot = -0.30900
CZtot = -0.34840 Cntot = -0.00000 Cn'tot = -0.00000

CLtot = 0.34839
CDtot = 0.00626
CDvis = 0.00000 CDind = 0.0062649
CLff = 0.34816 CDff = 0.0052006 | Trefftz
CYff = 0.00000 e = 0.7865 | Plane
CXtot = 0.00613 Cltot = -0.00000 Cl'tot = -0.00000
CYtot = -0.00000 Cmtot = -0.30835
CZtot = -0.34929 Cntot = -0.00000 Cn'tot = -0.00000

CLtot = 0.34929
CDtot = 0.00606
CDvis = 0.00000 CDind = 0.00606
CLff = 0.34905 CDff = 0.00525 | Trefftz
CYff = 0.00000 e = 0.7834 | Plane

slat = 0.00000
flap = 0.00000
aileron = 0.00000
elevator = 0.00000

---------------------------------------------------------------

Geometry-axis derivatives...

axial vel. u sideslip vel. v normal vel. w
---------------- ---------------- ----------------
x force CX | CXu = -0.001486 CXv = -0.000000 CXw = 0.380527
y force CY | CYu = -0.000000 CYv = -0.824938 CYw = 0.000000
z force CZ | CZu = -0.498616 CZv = -0.000002 CZw = -5.687328
x mom. Cl | Clu = -0.000000 Clv = -0.221884 Clw = -0.000004
y mom. Cm | Cmu = -0.337190 Cmv = -0.000002 Cmw = -8.052002
z mom. Cn | Cnu = -0.000000 Cnv = 0.274423 Cnw = -0.000000
x force CX | CXu = -0.001211 CXv = -0.000000 CXw = 0.386059
y force CY | CYu = -0.000000 CYv = -0.824962 CYw = -0.000000
z force CZ | CZu = -0.500388 CZv = -0.000000 CZw = -5.687412
x mom. Cl | Clu = -0.000000 Clv = -0.222059 Clw = -0.000001
y mom. Cm | Cmu = -0.335879 Cmv = -0.000000 Cmw = -8.052497
z mom. Cn | Cnu = -0.000000 Cnv = 0.274454 Cnw = -0.000000

roll rate p pitch rate q yaw rate r
---------------- ---------------- ----------------
x force CX | CXp = 0.000000 CXq = 0.792832 CXr = 0.000000
y force CY | CYp = -0.169261 CYq = -0.000006 CYr = 1.005343
z force CZ | CZp = -0.000008 CZq = -27.148266 CZr = 0.000002
x mom. Cl | Clp = -0.483353 Clq = -0.000023 Clr = 0.225224
y mom. Cm | Cmp = -0.000016 Cmq = -78.263054 Cmr = 0.000003
z mom. Cn | Cnp = 0.050102 Cnq = 0.000003 Cnr = -0.519628
x force CX | CXp = 0.000000 CXq = 0.798468 CXr = 0.000000
y force CY | CYp = -0.168990 CYq = -0.000000 CYr = 1.005460
z force CZ | CZp = -0.000001 CZq = -27.151451 CZr = 0.000000
x mom. Cl | Clp = -0.483371 Clq = -0.000003 Clr = 0.225286
y mom. Cm | Cmp = -0.000002 Cmq = -78.269577 Cmr = 0.000001
z mom. Cn | Cnp = 0.050032 Cnq = 0.000000 Cnr = -0.519672

slat d1 flap d2 aileron d3 elevator d4
---------------- ---------------- ---------------- ----------------
x force CX | CXd1 = -0.000453 CXd2 = 0.000348 CXd3 = 0.000000 CXd4 = 0.000038
y force CY | CYd1 = -0.000000 CYd2 = 0.000000 CYd3 = 0.000311 CYd4 = -0.000000
z force CZ | CZd1 = 0.000608 CZd2 = -0.025029 CZd3 = 0.000000 CZd4 = -0.012004
x force CX | CXd1 = -0.000464 CXd2 = 0.000340 CXd3 = 0.000000 CXd4 = 0.000037
y force CY | CYd1 = -0.000000 CYd2 = 0.000000 CYd3 = 0.000314 CYd4 = -0.000000
z force CZ | CZd1 = 0.000607 CZd2 = -0.025030 CZd3 = 0.000000 CZd4 = -0.012005
x mom. Cl | Cld1 = 0.000000 Cld2 = -0.000000 Cld3 = 0.001916 Cld4 = -0.000000
y mom. Cm | Cmd1 = 0.000131 Cmd2 = -0.023143 Cmd3 = 0.000000 Cmd4 = -0.058250
z mom. Cn | Cnd1 = 0.000000 Cnd2 = -0.000000 Cnd3 = -0.000003 Cnd4 = 0.000000


y mom. Cm | Cmd1 = 0.000131 Cmd2 = -0.023142 Cmd3 = 0.000000 Cmd4 = -0.058256
z mom. Cn | Cnd1 = 0.000000 Cnd2 = -0.000000 Cnd3 = -0.000007 Cnd4 = 0.000000
62 changes: 31 additions & 31 deletions regression/scripts/AVL/avl_files/body_axis_derivatives_case_01_04.txt
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Original file line number Diff line number Diff line change
Expand Up @@ -13,54 +13,54 @@

Run case: case_01_04

Alpha = 5.00000 pb/2V = -0.00000 p'b/2V = -0.00000
Alpha = 5.00000 pb/2V = 0.00000 p'b/2V = 0.00000
Beta = 0.00000 qc/2V = 0.00000
Mach = 0.050 rb/2V = -0.00000 r'b/2V = -0.00000
Mach = 0.050 rb/2V = 0.00000 r'b/2V = 0.00000

CXtot = 0.03763 Cltot = -0.00000 Cl'tot = -0.00000
CYtot = 0.00000 Cmtot = -0.72902
CZtot = -0.64335 Cntot = -0.00000 Cn'tot = -0.00000

CLtot = 0.64418
CDtot = 0.01858
CDvis = 0.00000 CDind = 0.0185847
CLff = 0.64424 CDff = 0.0166451 | Trefftz
CYff = -0.00000 e = 0.8414 | Plane
CXtot = 0.03822 Cltot = -0.00000 Cl'tot = -0.00000
CYtot = 0.00000 Cmtot = -0.72840
CZtot = -0.64423 Cntot = -0.00000 Cn'tot = -0.00000

CLtot = 0.64511
CDtot = 0.01807
CDvis = 0.00000 CDind = 0.01807
CLff = 0.64513 CDff = 0.01672 | Trefftz
CYff = -0.00000 e = 0.8401 | Plane

slat = 0.00000
flap = 0.00000
aileron = 0.00000
elevator = 0.00000

---------------------------------------------------------------

Geometry-axis derivatives...

axial vel. u sideslip vel. v normal vel. w
---------------- ---------------- ----------------
x force CX | CXu = 0.002579 CXv = -0.000000 CXw = 0.834044
y force CY | CYu = -0.000000 CYv = -0.830650 CYw = 0.000001
z force CZ | CZu = -0.794641 CZv = -0.000002 CZw = -5.680476
x mom. Cl | Clu = -0.000001 Clv = -0.257270 Clw = -0.000004
y mom. Cm | Cmu = -0.755790 Cmv = -0.000003 Cmw = -8.090524
z mom. Cn | Cnu = -0.000000 Cnv = 0.270724 Cnw = -0.000001
x force CX | CXu = 0.003045 CXv = 0.000000 CXw = 0.842346
y force CY | CYu = -0.000000 CYv = -0.830633 CYw = 0.000000
z force CZ | CZu = -0.796416 CZv = -0.000000 CZw = -5.680411
x mom. Cl | Clu = -0.000000 Clv = -0.257446 Clw = -0.000001
y mom. Cm | Cmu = -0.754527 Cmv = -0.000001 Cmw = -8.090510
z mom. Cn | Cnu = -0.000000 Cnv = 0.270751 Cnw = -0.000000

roll rate p pitch rate q yaw rate r
---------------- ---------------- ----------------
x force CX | CXp = 0.000001 CXq = 2.275026 CXr = 0.000000
y force CY | CYp = -0.063474 CYq = -0.000006 CYr = 1.023238
z force CZ | CZp = -0.000008 CZq = -27.091022 CZr = 0.000002
x mom. Cl | Clp = -0.480824 Clq = -0.000023 Clr = 0.269478
y mom. Cm | Cmp = -0.000016 Cmq = -78.417145 Cmr = 0.000005
z mom. Cn | Cnp = -0.016426 Cnq = 0.000002 Cnr = -0.522681
x force CX | CXp = 0.000000 CXq = 2.286721 CXr = 0.000000
y force CY | CYp = -0.063161 CYq = -0.000001 CYr = 1.023350
z force CZ | CZp = -0.000001 CZq = -27.092983 CZr = 0.000000
x mom. Cl | Clp = -0.480848 Clq = -0.000003 Clr = 0.269540
y mom. Cm | Cmp = -0.000002 Cmq = -78.415459 Cmr = 0.000001
z mom. Cn | Cnp = -0.016504 Cnq = 0.000000 Cnr = -0.522722

slat d1 flap d2 aileron d3 elevator d4
---------------- ---------------- ---------------- ----------------
x force CX | CXd1 = -0.000843 CXd2 = 0.000914 CXd3 = 0.000000 CXd4 = 0.000195
y force CY | CYd1 = -0.000000 CYd2 = 0.000000 CYd3 = 0.000273 CYd4 = -0.000000
z force CZ | CZd1 = 0.000623 CZd2 = -0.024790 CZd3 = 0.000000 CZd4 = -0.011907
x force CX | CXd1 = -0.000855 CXd2 = 0.000911 CXd3 = 0.000000 CXd4 = 0.000194
y force CY | CYd1 = -0.000000 CYd2 = 0.000000 CYd3 = 0.000276 CYd4 = -0.000000
z force CZ | CZd1 = 0.000621 CZd2 = -0.024791 CZd3 = 0.000000 CZd4 = -0.011908
x mom. Cl | Cld1 = 0.000000 Cld2 = -0.000000 Cld3 = 0.001895 Cld4 = -0.000000
y mom. Cm | Cmd1 = 0.000162 Cmd2 = -0.022890 Cmd3 = 0.000000 Cmd4 = -0.057861
z mom. Cn | Cnd1 = 0.000000 Cnd2 = -0.000000 Cnd3 = 0.000048 Cnd4 = 0.000000


y mom. Cm | Cmd1 = 0.000162 Cmd2 = -0.022890 Cmd3 = 0.000000 Cmd4 = -0.057863
z mom. Cn | Cnd1 = 0.000000 Cnd2 = -0.000000 Cnd3 = 0.000044 Cnd4 = 0.000000
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