Function to interpolate airfoil data: Code, sample run,
& output
Matlab code file “builddatatable.m”
%Function to convert airfoil lift and drag coefficients
%>to tangential and normal coefficients
%>then the function interpolate linearly the Ctang and Cnorm data
%>in between angle of attack and in between Reynolds number
%
%INPUT
%Re = vector of airfoil data Reynolds number;
%filelist = vector of airfoil data filename character strings;
%
%OUTPUT
%uniformCtang = tangential coefficient data in table form
%>with 0.1 degree angle of attack increment between column
%>and 10e3 Reynolds number increment between row
%>!large output so suppress screen output with semicolon ';'
%uniformCnorm = normal coefficient data
function [uniformCtang, uniformCnorm] = builddatatable(Re, filelist);
alpha_INCREMENT = 0.1;
reynolds_INCREMENT = 10e3;
for i = 1:1:length(filelist(:,1))
%Load each set of lift and drag data
tabAlphaClCd = load(filelist(i,:));
%Convert each set to tangent and normal data
PRINT_data_filename = filelist(i,:) %Redundant print screen info
tabAlphaCtangCnorm = convert2CtangCnorm(tabAlphaClCd) %print screen!
%Interpolate each set of tangent and normal data to uniform alpha
uniformAlpha = [0 : alpha_INCREMENT : tabAlphaCtangCnorm(end,1)]';
CtangCnorm = interp1(tabAlphaCtangCnorm(:,1),...
tabAlphaCtangCnorm(:,2:3), uniformAlpha, 'linear');
%Separate vectors of tangent and normal data into rows in diff tab
Ctang(i,:) = CtangCnorm(:,1)';
Cnorm(i,:) = CtangCnorm(:,2)';
end
%Interpolate tangent and normal data TABLE to uniform Reynolds number
uniformRe = [0 : reynolds_INCREMENT : Re(end)];
uniformCtang = interp1(Re, Ctang, uniformRe, 'linear');
uniformCnorm = interp1(Re, Cnorm, uniformRe, 'linear');
%---------------------------------------------------------------------
function tabAlphaCtangCnorm = convert2CtangCnorm(tabAlphaClCd)
alpha = tabAlphaClCd(:,1);
Cl = tabAlphaClCd(:,2);
Cd = tabAlphaClCd(:,3);
Cnorm = -Cl.*cos(alpha/180*pi) - Cd.*sin(alpha/180*pi);
Ctang = Cl.*sin(alpha/180*pi) - Cd.*cos(alpha/180*pi);
tabAlphaCtangCnorm = [alpha Ctang Cnorm];
%%STUB figure; plot(alpha, Ctang); hold on; plot(alpha, Cnorm);
A sample run of “builddatatable.m” in Matlab
» re = [0 1200e3]'
re =
0
1200000
» fl = ['A10.NACA12Re0300.data'; 'A10.NACA12Re0300.data']
fl =
A10.NACA12Re0300.data
A10.NACA12Re0300.data
» [ct, cn] = builddatatable(re, fl); %suppress the long screen output!
PRINT_data_filename =
A10.NACA12Re0300.data
tabAlphaCtangCnorm =
0 -0.0085 0
2.3644 -0.0024 -0.2001
5.9115 0.0308 -0.4993
8.9029 0.0826 -0.7656
12.4617 0.1148 -0.8516
12.5648 0.0731 -0.8628
13.0000 0.0729 -0.8511
14.0000 0.0722 -0.8290
15.0000 0.0716 -0.8125
16.0000 0.0709 -0.8004
17.0000 0.0702 -0.7920
18.0000 0.0694 -0.7866
19.0000 0.0686 -0.7837
20.0000 0.0678 -0.7830
21.0000 0.0669 -0.7841
22.0000 0.0660 -0.7867
23.0000 0.0650 -0.7907
24.0000 0.0640 -0.7957
25.0000 0.0630 -0.8018
26.0000 0.0620 -0.8087
27.0000 0.0609 -0.8164
28.0000 0.0598 -0.8247
29.0000 0.0587 -0.8335
30.0000 0.0576 -0.8428
31.0000 0.0564 -0.8525
32.0000 0.0552 -0.8626
33.0000 0.0540 -0.8729
34.0000 0.0527 -0.8835
35.0000 0.0515 -0.8943
36.0000 0.0502 -0.9053
37.0000 0.0489 -0.9165
38.0000 0.0477 -0.9277
39.0000 0.0463 -0.9390
40.0000 0.0450 -0.9504
41.0000 0.0437 -0.9618
42.0000 0.0424 -0.9732
43.0000 0.0410 -0.9845
44.0000 0.0397 -0.9959
45.0000 0.0384 -1.0072
46.0000 0.0370 -1.0184
47.0000 0.0357 -1.0295
48.0000 0.0344 -1.0406
49.0000 0.0330 -1.0515
50.0000 0.0317 -1.0623
51.0000 0.0304 -1.0730
52.0000 0.0291 -1.0835
53.0000 0.0278 -1.0939
54.0000 0.0265 -1.1040
55.0000 0.0252 -1.1140
56.0000 0.0240 -1.1238
57.0000 0.0228 -1.1334
58.0000 0.0215 -1.1428
59.0000 0.0204 -1.1520
60.0000 0.0192 -1.1609
61.0000 0.0180 -1.1697
62.0000 0.0169 -1.1781
63.0000 0.0158 -1.1863
64.0000 0.0147 -1.1943
65.0000 0.0137 -1.2019
66.0000 0.0127 -1.2093
67.0000 0.0117 -1.2165
68.0000 0.0108 -1.2233
69.0000 0.0099 -1.2298
70.0000 0.0090 -1.2361
71.0000 0.0081 -1.2420
72.0000 0.0073 -1.2476
73.0000 0.0066 -1.2530
74.0000 0.0058 -1.2579
75.0000 0.0051 -1.2626
76.0000 0.0045 -1.2669
77.0000 0.0039 -1.2709
78.0000 0.0033 -1.2746
79.0000 0.0028 -1.2779
80.0000 0.0023 -1.2808
81.0000 0.0019 -1.2834
82.0000 0.0015 -1.2856
83.0000 0.0011 -1.2875
84.0000 0.0008 -1.2890
85.0000 0.0006 -1.2901
86.0000 0.0004 -1.2908
87.0000 0.0002 -1.2912
88.0000 0.0001 -1.2912
89.0000 0.0000 -1.2908
90.0000 0.0000 -1.2900
PRINT_data_filename =
A10.NACA12Re0300.data
tabAlphaCtangCnorm =
0 -0.0085 0
2.3644 -0.0024 -0.2001
5.9115 0.0308 -0.4993
8.9029 0.0826 -0.7656
12.4617 0.1148 -0.8516
12.5648 0.0731 -0.8628
13.0000 0.0729 -0.8511
14.0000 0.0722 -0.8290
15.0000 0.0716 -0.8125
16.0000 0.0709 -0.8004
17.0000 0.0702 -0.7920
18.0000 0.0694 -0.7866
19.0000 0.0686 -0.7837
20.0000 0.0678 -0.7830
21.0000 0.0669 -0.7841
22.0000 0.0660 -0.7867
23.0000 0.0650 -0.7907
24.0000 0.0640 -0.7957
25.0000 0.0630 -0.8018
26.0000 0.0620 -0.8087
27.0000 0.0609 -0.8164
28.0000 0.0598 -0.8247
29.0000 0.0587 -0.8335
30.0000 0.0576 -0.8428
31.0000 0.0564 -0.8525
32.0000 0.0552 -0.8626
33.0000 0.0540 -0.8729
34.0000 0.0527 -0.8835
35.0000 0.0515 -0.8943
36.0000 0.0502 -0.9053
37.0000 0.0489 -0.9165
38.0000 0.0477 -0.9277
39.0000 0.0463 -0.9390
40.0000 0.0450 -0.9504
41.0000 0.0437 -0.9618
42.0000 0.0424 -0.9732
43.0000 0.0410 -0.9845
44.0000 0.0397 -0.9959
45.0000 0.0384 -1.0072
46.0000 0.0370 -1.0184
47.0000 0.0357 -1.0295
48.0000 0.0344 -1.0406
49.0000 0.0330 -1.0515
50.0000 0.0317 -1.0623
51.0000 0.0304 -1.0730
52.0000 0.0291 -1.0835
53.0000 0.0278 -1.0939
54.0000 0.0265 -1.1040
55.0000 0.0252 -1.1140
56.0000 0.0240 -1.1238
57.0000 0.0228 -1.1334
58.0000 0.0215 -1.1428
59.0000 0.0204 -1.1520
60.0000 0.0192 -1.1609
61.0000 0.0180 -1.1697
62.0000 0.0169 -1.1781
63.0000 0.0158 -1.1863
64.0000 0.0147 -1.1943
65.0000 0.0137 -1.2019
66.0000 0.0127 -1.2093
67.0000 0.0117 -1.2165
68.0000 0.0108 -1.2233
69.0000 0.0099 -1.2298
70.0000 0.0090 -1.2361
71.0000 0.0081 -1.2420
72.0000 0.0073 -1.2476
73.0000 0.0066 -1.2530
74.0000 0.0058 -1.2579
75.0000 0.0051 -1.2626
76.0000 0.0045 -1.2669
77.0000 0.0039 -1.2709
78.0000 0.0033 -1.2746
79.0000 0.0028 -1.2779
80.0000 0.0023 -1.2808
81.0000 0.0019 -1.2834
82.0000 0.0015 -1.2856
83.0000 0.0011 -1.2875
84.0000 0.0008 -1.2890
85.0000 0.0006 -1.2901
86.0000 0.0004 -1.2908
87.0000 0.0002 -1.2912
88.0000 0.0001 -1.2912
89.0000 0.0000 -1.2908
90.0000 0.0000 -1.2900
» save('Ctang.A10.NACA12.table', 'ct', '-ascii');
» save('Cnorm.A10.NACA12.table', 'cn', '-ascii');
» plot([0:0.1:90]', ct(1,:)')
» plot([0:0.1:90]', cn(1,:)')
»
Plotted graph of the interpolated airfoil data at certain Re number
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