LAMDES Manual - the AOE home page
Contents
1. Lamar Design Code mods by W H Mason Lamar program sample input revised forward swept wing plan 2 0 xmref 8 0000 cref 89 5000 tdklue 1 0 Case 3 0 spnklu 0 0 sref 26640 0000 lst REFERENCE PLANFORM HAS 5 CURVES ROOT CHORD HEIGHT 8 8000 POINT X PS SWEEP DIHEDRAL REF REF ANGLE ANGLE L 76 9500 0 0000 0 00000 0 00000 2 76 9500 34 0000 31 71155 0 00000 3 57 6100 65 3000 90 00000 0 00000 4 33 6400 65 3000 6 18142 0 00000 5 30 2500 34 0000 0 00000 0 00000 6 30 2500 0 0000 2nd REFERENCE PLANFORM HAS 5 CURVES ROOT CHORD HEIGHT 0 0000 POINT X Y SWEEP DIHEDRAL REF REF ANGLE ANGLE 1 17 9000 0 0000 0 00000 0 00000 2 17 9000 34 0000 26 21138 0 00000 3 46 1000 164 0000 90 00000 0 00000 4 5 6000 164 0000 45 29687 0 00000 5 139 9000 20 0000 0 00000 0 00000 6 139 9000 0 0000 SCW 10 0 vic 20 0 xitmax 40 0 epsmax 0 00060 CONFIGURATION NO 1 delta ord shift for moment 8 0000 CURVE 1 IS SWEP 0 0000 DEGREES ON PLANFORM 1 CURVE 1 IS SWEP 0 0000 DEGREES ON PLANFORM 2 BREAK POINTS FOR THIS CONFIGURATION POINT X Y Z SWEEP DIHEDRAL ANGLE ANGLE 1 76 9500 0 0000 8 8000 0 0000 0 0000 2 76 9500 20 0000 8 8000 0 0000 0 0000 3 76 9500 34 0000 8 8000 31 7116 0 0000 4 57 6100 65 3000 8 8000 90 0000 0 0000 5 33 6400 65 3000 8 8000 6 1814 0 0000 6 30 2500 34 0000 8 8000 0 0000 0 0000 7 30 2500 0 0000 82. 5804 4030 2256 delta z 4093 4745 5414 6080 6674 7122 7381 7444 7339 7102 6761 z zle c 0000 0046 0092 0138 0182 0222 0255 0283 0306 0325 0340 D 30 Applied Computational Aerodynamics 0 2750 0 1104 9 0481 236333 0 0353 0 3000 0 1089 9 8707 5 9 8 19 0 0363 0 3250 0 1070 10 6932 3 5220 0 0371 0 3500 0 1050 11 5158 3 4534 0 0376 0 3750 0 1026 12533883 3 3766 0 0379 0 4000 0 1001 13 1609 33 4 2920 0 0379 0 4250 0 0973 13 9834 3 2003 0 0377 0 4500 0 0943 14 8060 3 1020 0 0373 0 4750 0 0911 15 6285 2 9975 0 0367 0 5000 0 0878 16 4511 2 8872 0 0359 0 5250 0 0842 17 2737 SA LS 0 0350 0 5500 0 0806 18 0962 2 6505 0 0339 0 5750 0 0767 18 9188 2 5247 0 0327 0 6000 0 0728 19 7413 2 3945 0 0313 0 6250 0 0687 20 5639 2 2601 0 0298 0 6500 0 0645 21 3864 2 1219 0 0282 0 6750 0 0602 22 2090 1 9804 0 0265 0 7000 0 0558 23 0315 1 8358 0 0247 0 7250 0 0913 23 8541 1 6886 0 0228 0 7500 0 0468 24 6766 1 5391 0 0209 0 7750 0 0422 25 4992 1 3875 0 0189 0 8000 0 0375 26 3218 1 2341 0 0168 0 8250 0 0328 27 1443 1 0792 0 0147 0 8500 0 0281 27 9669 09233 0 0125 0 8750 0 0233 28 7894 0 7671 0 0104 0 9000 0 0186 29 6120 0 6114 0 0082 0 9250 0 0139 30 4345 0 4569 0 0061 0 9500 0 0092 31 420 11 0 3038 0 0041 0 9750 0 0046 32 0796 Q LOLA 0 0020 1 0000 0 0000 32 9022 0 0000 0 0000 Note this output is repeated for each span st
3. 6750 7750 8750 9750 0000000000 dz dx dz dx 0783 0034 0572 0982 1306 1597 1740 1854 1898 1845 mean camber shap x c 0000 0250 0500 0750 1000 1250 1500 1750 2000 2250 2500 00000 000000o 0073000 Y NNROOD 0000 d OY ON dm 19 20 20 21 22 22 23 24 24 294 26 26 7158 3895 0632 7368 4105 0842 1579 4316 1053 7790 4526 2 2 63 8000 4737 1474 8211 4948 1684 8421 5158 1895 8632 5369 2105 8842 5579 23106 9053 5790 2527 9263 6000 2737 9474 y b 2 z c 1036 1056 1076 1097 Tas 1128 1136 11 38 1135 1128 TELS Tuesday January 21 1997 0 0 0 0 V 0 0 0 0 0 0 3456 3857 4 4 4 4 4 4 4 4 4 166 399 563 660 689 651 548 383 160 3884 43596 3181 1 2760 L 22 97 1794 1254 0679 1 0074 0 9440 8781 8100 7400 6682 5950 5205 4452 3696 2942 2196 1458 0728 0000 3232 at control points 0252 0275 0294 0310 0323 0334 0343 0349 0353 0354 0353 0301 0346 0339 0331 0322 0310 0298 0284 0269 0253 0236 0218 0200 0181 0161 0141 0120 0100 0079 0059 0039 0020 0000 chord 32 9022 from front to rear interpolated to 41 points delta x 0 0 3004 Ex Nro 0000 8226 6451 4677 2902 1128 9353 adoro
4. 0044 0049 0054 0059 0064 0069 0074 0079 0084 0088 0092 0096 0100 0102 0105 0106 0106 0105 0103 0100 0095 0089 0083 0076 0067 0059 0049 0040 0030 0020 0010 0000
5. 8000 SECOND PLANFORM BREAK POINTS 119000 0 0000 0 0000 0 0000 0 0000 2 17 9000 34 0000 0 0000 26 2114 0 0000 3 2 4908 65 3000 0 0000 26 2114 0 0000 4 46 1000 164 0000 0 0000 90 0000 0 0000 3 5 6000 164 0000 0 0000 45 2969 0 0000 6 139 9000 20 0000 0 0000 0 0000 0 0000 7 139 9000 0 0000 0 0000 Tuesday January 21 1997 report typos and errors to W H Mason Appendix D Programs D 25 280 HORSESHOE VORTICES USED PLANFORM TOTAL SPANWISE 1 80 8 2 200 20 10 HORSESHOE VORTICES IN EACH CHORDWISE xcfw 0 00 xcft 0 65 fkon ficam 1 00 punch 0 00 crbmnt cmb 10 iflag dl relax 0 03 fioutw 1 00 cdo firbm 0 00 yrbm 0 0000 zrbm drag polar on canard conv sec there are 1 0 polars on this surface 18 0 points this polar qcl 0000 1000 2500 3000 4000 5000 5500 6000 6500 7000 7500 8000 8500 8800 9150 0000 2000 8000 HhHHOOOooo 000000000 drag polar SO 0000000 0 0 00 0 00 00 acd 0000 0000 0002 0008 0018 0032 0040 0054 0069 0088 0113 0148 0198 0240 0360 0880 2680 9880 planform 1 there are 1 0 polars on this surface 22530 points this polar qal 0000 2000 3000 4000 5000 6000 7000 8000 9000 9500 9700 9900 0000 0200 0400 HhHHOOoOooo 0000000 Tuesday January 21 1997 2900000000 000000 acd 0003 0003 0005 0008 0012 0018 0024 0032 00
6. up to four digits Number of chordwise horseshoe vortices to be used to represent the wing a maximum of 20 may be used do not set to zero nominal number of spanwise rows at which chordwise horseshoe may be located a maximum of 50 may be used The product of SCW and SSW cannot exceed 400 see VLM4997 chapter for details of vortex layout Mach number used to apply Prandtl Glauert comressibility correction factor design lift coefficient for lifting system Maximum number of iterations allowed in finding the solution for minimum pressure drag with arbitrary polars input Must be less than 50 20 is sufficient for most cases The convergence criteria for the general polar case A value of 0005 appears to be reasonable The chord fraction a at which the chord load shape changes from rooftop to a linear decrease to zero at the trailing edge on the first planform See the introduction to this section for more discussion Same as XCFW except applies to the second planform Clue for constraints 0 body moment constraint no constraints 2 root bending moment constraint 3 both moment anf root bending moment constraints The design wing CM when FKON 0 Camber computation clue 0 no cambers computed 1 wing cambers computed D 22 Applied Computational Aerodynamics 6 PUNCH 7 CRBMT 3 C 8F10 6 1 RELAX 2 FIOUTW 3 CDO clue to punch cambers out 0 no punch file created 1
7. used to model the effects of viscosity the po lars are input in a streamwise coordinate system The user is responsible for adjusting them from 2D to 3D This program uses an input file that is very similar to but not the same as the VLMpcv2 code It is based on the same geometry and coordinate system ideas Section D 6 should be consulted for a discussion of the geometry system Card Format Field Name Remarks 1 Literal DATA Title card for the data set 2 8F10 6 1 PLAN Number of lifting surfaces for the configuration use 1 or 2 2 XMREF c g shift from origin of input planform coordinate system the program originally trimmed the configuration about the input planform origin is a c g shift forward is a c g shift aft 3 CREF reference chord of the configuration used only to nondimensionalize the pitching moment coefficients 4 SREF reference area of the configuration Tuesday January 21 1997 D 20 Applied Computational Aerodynamics 5 6 7 TDKLUE CASE SPNKLU minimization clue 0 minimize induced drag only minimize induced plus pressure drag options for the drag polar 0 model polar same a CLmin CDO for each surface see note 3 below 1 model polar each surface has its own a CLmin CDO 2 one general polar for entire config 3 one general polar for each surface spanload clue 0 spanload is internally computed using the minimization 1 no minimizatio
8. 0000 0 06850 0 01053 0 07903 6 0 03155 0 90000 0 06863 0 00976 0 07839 7 0 02773 0 90000 0 06876 0 00915 0 07791 8 0 02043 0 90000 0 06885 0 00886 0 07772 9 0 01549 0 90000 0 06893 0 00868 0 07761 10 0 01218 0 90000 0 06898 0 00856 0 07754 11 0 00994 0 90000 0 06902 0 00847 0 07749 12 0 00847 0 90000 0 06905 0 00841 0 07746 13 0 00724 0 90000 0 06907 0 00836 0 07743 14 0 00616 0 90000 0 06909 0 00832 0 07741 15 0 00519 0 90000 0 06911 0 00829 0 07740 16 0 00442 0 90000 0 06913 0 00826 0 07739 17 0 00371 0 90000 0 06915 0 00823 0 07738 18 0 00310 0 90000 0 06916 0 00821 0 07737 19 0 00263 0 90000 0 06917 0 00819 0 07736 20 0 00221 0 90000 0 06918 0 00817 0 07736 21 0 00183 0 90000 0 06919 0 00816 0 07735 22 0 00154 0 90000 0 06920 0 00815 Dl 07 735 23 0 00131 0 90000 0 06921 0 00814 0 07734 24 0 00112 0 90000 0 06921 0 00813 0 07734 25 0 00095 0 90000 0 06922 0 00812 0 07734 26 0 00084 0 90000 0 06923 0 00811 0 07734 27 0 00076 0 90000 0 06923 0 00810 0 07733 28 0 00069 0 90000 0 06924 0 00810 0 07733 29 0 00064 0 90000 0 06924 0 00809 0 07733 30 0 00061 0 90000 0 06924 0 00809 0 07733 31 0 00057 0 90000 0 06925 0 00808 0 07733 induced pressure drag was minimized on this run ref chord 89 500 c average 81 2195 true area 32771 566 ref area 26640 000 b 2 164 0000 ref ar 4 0384 true ar 3 2828 Mach number 0 9000 first planform cl 0 17126 cm 0 11493 cb 0 01502 second planform cl 0 72874 cm 0 21493 cb 0 18341 lst planfor
9. 001 50000 30002 10001 90002 70002 50002 30003 10003 90003 55002 20000 00000 80000 35000 90000 90000 90000 90000 0584 0572 0559 0547 0535 0522 0510 0497 0485 0472 0458 0445 0431 0416 0401 0385 0368 0349 0330 0309 0287 0264 0240 0215 0189 0163 0137 0110 0082 0055 0027 0000 21 30 333 36 394 42 45 48 SL 54 51 61 64 67 70 Loe 76 79 82 85 88 91 94 97 100 103 106 109 TL2 115 118 L223 y b 2 0 32317 APS 22774 18232 13963 09695 03598 97500 92500 87500 82500 77500 72500 67500 62500 57500 52500 47500 42409 31317 32317 271317 22774 18232 13963 09695 03598 Tuesday January 21 1997 37317 4500 5000 5500 6000 6500 7000 7500 8000 8500 9000 9500 0000 0500 1000 1500 2000 2500 3000 3500 4000 4500 5000 5500 6000 6500 7000 7500 8000 8500 9000 9500 0000 WWW WU 754 ds Y RA DN BO Oo twist 71469 91587 36720 25835 47910 60813 49868 91663 45816 44655 38027 36750 75520 sol O73 46040 34168 13 L04 67249 88238 36595 52797 51491 49845 79378 77474 11226 22109 98970 1286 9763 8249 6742 5237 3728 2210 0676 P 7537 5919 4262 2558 0791 8940 6978 4878 2627 0221 7669 4982 2174 9253 6231 3115 9920 6662 P 0031 6690 3345 0000
10. 1 46602 0 39760 0 00079 15 9000 0457323 1 50210 0 38162 0 00074 5 9000 0 56730 1 50210 0 37767 0 00073 mean camber lines to obtain the spanload subsonic linear theory y 61 2000 y b 2 0 3732 chord 26 9474 slopes dz dx at control points from front to rear E dz dx 0 0750 0 1295 0 1750 0 0672 0 2750 0 0194 0 3750 0 0200 0 4750 0 0522 0 5750 0 0775 0 6750 0 0960 0 7750 0 1077 0 8750 0 1122 0 9750 0 1081 mean camber shape interpolated to 41 points x c Ze delta x delta z z zle c 0 0000 0 0299 0 0000 0 8067 0 0000 0 0250 0 0332 0 6737 0 8944 0 0040 0 0500 0 0365 1 3474 0 9831 0 0080 0 0750 0 0398 2 0211 1 0717 0 0121 0 1000 0 0429 2 6947 1 1558 0 0159 0 1250 0 0457 3 3684 1 2310 0 0195 0 1500 0 0480 4 0421 1 2945 0 0226 Tuesday January 21 1997 report typos and errors to W H Mason Appendix D Programs D 29 4 150 2000 2250 2500 2750 3000 3250 3500 3750 4000 4250 4500 4750 5000 5250 5500 5750 6000 6250 6500 6750 7000 ALO 7500 FOO 8000 8250 8500 8750 9000 9250 9500 9750 0000 PRO0000000000000000000000000000000o0Oo 0499 0514 0526 0534 0540 0544 0545 0544 0540 0534 0525 0515 0503 0489 0474 0456 0438 0418 0396 0374 0350 0326 0301 0275 0248 0221 0193 0165 0137 0109 0081 0054 0027 0000 y 53 0000 slopes x c 0750 1750 2750 3750 4750 5750
11. 44 0053 0057 0062 0065 0073 0082 planform 2 ROW 00 000 0000 0000 D 26 Applied Computational Aerodynamics KA G A A L A LO LM 70 IL 71 BOTL 164 000 NMA KBOT 50 induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced induced Tuesday January 21 0600 0800 1000 1250 1300 2000 0000 JM BOL KBOT drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag 1997 Nooooo o 72 cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd cd 0093 0109 0128 0240 0360 2040 1240 IM 73 65 300 0 06815 0 06818 0 06827 0 06839 0 06850 0 06863 0 06876 0 06885 0 06893 0 06898 0 06902 0 06905 0 06907 0 06909 0 06911 0 06913 0 06915 0 06916 0 06917 0 06918 0 06919 0 06920 0 06921 0 06921 0 06922 0 06923 TSPAN 164 000 SNN 1 6400 NMA KBIT pressure pressure pressure pressure pressure pressure pressure pressure pressure pressure pressure pressure pressure pressure pressure pressure pressure pressure pressure pressu
12. ation Most other stations are omitted y 5 9000 y b 2 0 0360 chord 122 0000 slopes dz dx at control points from front to rear x c dz dx 0 0750 0 0501 0 1750 0 0505 0 2750 0 0495 0 3750 0 0500 0 4750 0 0537 0 5750 0 0623 0 6750 0 0814 0 7750 0 0975 0 8750 0 1077 0 9750 kaa O REGAR O e mean camber shape interpolated to 41 points x c z c delta x delta z z zle c 0 0000 0 0697 0 0000 8 5090 0 0000 0 0250 0 0685 3 0500 8 3562 0 0005 0 0500 0 0672 6 1000 8 2034 0 0010 0 0750 0 0660 9 1500 8 0506 0 0015 0 1000 0 0647 12 2000 7 8975 0 0020 0 1250 0 0635 15 2500 7 7440 0 0024 0 1500 0 0622 18 3000 7 5900 0 0029 0 1750 0 0609 21 3500 7 4358 0 0034 0 2000 0 0597 24 4000 7 2818 0 0039 Tuesday January 21 1997 report typos and errors to W H Mason Appendix D Programs D 31 HhOoooooooooo0o000000000000000000000 2250 2500 2750 3000 3250 3500 3750 4000 4250 4500 4750 5000 5250 5500 5750 6000 6250 6500 6750 7000 7250 7500 7750 8000 8250 8500 8750 9000 9250 9500 9750 0000 twist table o ITS UE LQ A KA G 16 00 54 ON UT yi KM LA kde N o 21 22 23 24 25 26 27 28 STOP 61 53 44 37 29 22 15 5 159 tad 143 135 127 118 110 102 94 86 77 69 61 53 44 37 29 22 15 5 y 20000 00000 80000 35000 90000 90000 90000 90000 89999 70
13. cards output unit 7 Design root bending moment for FKON 2 The under relaxation factor for the general polar solution RELAX 03 to 3 1s satisfactory for most applications Output clue O full iteration history is output 1 only final results are output Basic drag coefficient that will be added to the drag computed by summing the induced drag and the profile drag contained in the input polars Arbitrary Polar Input the following cards are read only if CASE 2 2 Card Format Field Name 1 D Literal TITLE 2 D 8F10 5 1 FNCLCD 3 D 8F10 5 1 FQCL 2 FQCD Note 1 Card 3 D is read FNCLCD times Remarks The identifying title for the input drag polar for this surface The number of CL CD pairs used to define the input polar The value of streamwise lift coefficient for this pooint on the drag polar The value of streamwise drag coefficient for the given lift coefficient 2 Cards 1 D 2 D and 3 D are read for each planform if CASE 3 Spanload Input the following cards are read only if SPNKLU 1 Card Format Field Name 1 S Literal TITLE 2 9 7F10 5 1 FSPNPT 3 S 7F10 5 1 YSPNPT 2 CLSPNP Note 1 Card 3 S is read FSPNPT times Remarks This is the title card for the input spanloads Number of points on the spanload to be read in for this planform Span location in physical coordinates at which ccl ca is input y is positive here The spanload at YSPNPT 2 Cards 2 S and 3 S a
14. m CL 00001713 CDP 0 0042 CM 0 1150 CB 0 0151 2nd planform CL 0 7292 CDP 0 0038 CM 0 2149 CB 0 0000 no root bending moment constraint CL DES 0 90000 CL COMPUTED 0 9005 CM 0 0999 CD I 0 06925 E 0 9230 CDPRESS 0 00804 CDTOTAL 0 07729 Tuesday January 21 1997 D 28 Applied Computational Aerodynamics first planform Y CL C CAVE C CAVE CL CD 61 2000 0 21189 0 33178 0 63862 0 00651 53 0000 0 33566 0 40510 0 82857 0 01765 44 8000 0 41311 0 47842 0 86348 0 02166 37 3500 0 46740 0 54503 0 85757 0 02082 29 9000 0 49499 0 57498 0 86088 0 02129 22 9000 0 50260 0 57498 0 87411 0 02317 15 9000 0 50504 0 57498 0 87835 0 02377 5 9000 0 50631 0 57498 0 88056 0 02419 second planform 159 9000 0 33879 0 52480 0 64556 0 00208 151 7000 0 53136 0 57711 0 92072 0 00478 143 5000 0 64513 0 62942 1 02495 0 00752 135 3000 0 72403 0 68173 1 06206 0 00946 127 1000 0 78509 0 73404 1 06954 0 01006 118 9000 0 83563 0 78635 1 06267 0 00951 110 7000 0 87760 0 83866 1 04644 0 00855 102 5000 0 91055 0 89096 1 02198 0 00739 94 3000 0 93428 0 94327 0 99047 0 00622 86 1000 0 94681 0 99558 0 95101 0 00530 77 9000 0 94347 1 04789 0 90036 0 00443 69 5500 0 90911 1 10116 0 82559 0 00354 61 2000 0 82859 1 15442 0 71775 0 00258 53 0000 0 74419 1 20673 0 61670 0 00189 44 8000 0 67721 1 25904 0 53788 0 00145 37 3500 0 63142 1 30656 0 48327 0 00117 29 9000 0 60043 1 37894 0 43543 0 00096 22 9000 0 58289
15. n is done spanload is read in and e and pressure drag are computed Geometric Planform Data see the VLMpc section D 6 for more details Card 1 P 2 P Note 1 Format Field 8F10 6 1 2 3 4 5 6 7 8F10 6 1 2 3 4 Name AAN IT XS IT YS IT RTCDHT IT PDRGI IT PDRG2 IT PDRG3 IT XREG YREG DIH AMCD Remarks of straight lines defining this surface O not used in this code 0 not used in this code root chord height is higher CLmin q CDO X point of line segment positive is forward Y point of line segment positive is forward dihedral angle of line sweep wing move code set 1 for this program Card 2 P is read in AAN 1 times Surface description starts at forward centerline and works outboard and around returning to the aft centerline of the surface Cards 1 P and 2 P are read in as a set for each lifting surface see VLM4997 for clarification The model polar is given by C dA C 1 C Tuesday January 21 1997 2 lia E CDo report typos and errors to W H Mason Appendix D Programs D 21 Control Data corresponding to Group Two data in Lamar s nomenclature Card Format 1 C 6F5 3 2F10 6 2 C 6F10 4 Tuesday January 21 1997 Field 1 Name CONFIG SCW VIC XMCH CLDES XITMAX EPSMAX XCFW XCFT FKON CMB FICAM Remarks arbitrary configuration number or ID may include
16. re pressure pressure pressure pressure pressure pressure 20 drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag drag TSPANA DELTYB KBI1 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 cdpt 0 65 300 3 2800 01665 01441 01255 01139 01053 00976 00915 00886 00868 00856 00847 00841 00836 00832 00829 00826 00823 00821 00819 00817 00816 00815 00814 00813 00812 00811 report typos and errors to W H Mason Appendix D Programs D 27 induced drag cd 0 06923 pressure drag cdpt 0 00810 induced drag cd 0 06924 pressure drag cdpt 0 00810 induced drag cd 0 06924 pressure drag cdpt 0 00809 induced drag cd 0 06924 pressure drag cdpt 0 00809 induced drag cd 0 06925 pressure drag cdpt 0 00808 pressure drag iteration has converged k eps cl cdi cdp cdi cdp 1 28 66362 0 90000 0 06815 0 01665 0 08480 2 0 05789 0 90000 0 06818 0 01441 0 08260 3 0 05278 0 90000 0 06827 0 01255 0 08082 4 0 04274 0 90000 0 06839 0 01139 0 07978 5 0 03408 0 9
17. re read for each planform as a set Tuesday January 21 1997 report typos and errors to W H Mason Appendix D Programs D 23 Sample Input note it is important to put data in proper columns Lamar program sample input revised forward swept wing 2 000 8 000 89 50 26640 1 0 3 0 5 000 0 0 0 0 8 8 07 0 0 0 68 95 0 0 0 0 0 68 95 34 0 49 61 65 30 0 0 1 0 25 64 65 30 0 0 22 25 34 00 22 425 0 00 5 9 0 0 0 0 0 0 0 0 0 0 2 5390 0 0 0 0 0 25 90 34 0 38 10 164 0 0 0 1 0 2 40 164 0 0 0 20 147 90 2 0 0 147 90 0 0 1 0 10 0 20 0 9 0 90 40 0 0 0006 0 0 0 65 0 0 dO 1 0 0 030 1 0 0 0 0 0 0 0 0 0 drag polar on canard conv sec 18 0 0 00 0 0000 0 10 0 0000 0 25 0 0002 0 30 0 00078 0 40 0 00175 0 50 0 00315 0 55 0 0040 0 60 0 00535 0 65 0 00685 0 70 0 00880 ONES 0 01125 0 80 0 01485 0 85 0 01975 0 88 0 02400 0 915 0 03600 1 00 0 0880 1 20 0 2680 1 80 0 9880 drag polar 22 0 0 000 0 0003 0 200 0 0003 0 300 0 0005 0 400 0 0008 0 500 0 00125 0 600 0 00178 0 700 0 00244 0 800 0 00324 0 900 0 00442 0 950 0 00528 0 970 0 00570 0 990 0 00621 1 000 0 00650 1 020 0 00730 1 040 0 00820 1 060 0 00930 1 080 0 01090 1 100 0 01280 1 125 0 02400 1 130 0 03600 1 200 0 20400 2 000 2 12400 Tuesday January 21 1997 0 0 D 24 Applied Computational Aerodynamics Sample Output enter name of input file lamdes inp
18. report typos and errors to W H Mason Appendix D Programs D 19 D 4 LAMDES User s Manual This is the Lamar design program LamDes2 f It can be used as a non planar LIDRAG to get span e for multiple lifting surface cases when user supplies spanload It has also been called the Lamar Mason optimization code It finds the spanload to minimize the sum of the induced and pressure drag including canards or winglets It also provides the associated camber distribution for subsonic flow Since two surfaces are included it can find the minimum trimmed drag while satisfying a pitching moment constraint The program will prompt you for the input file name A sample input file called lamdes inp is on the disk and the output obtained from this case is included here References J E Lamar A Vortex Latice Method for the Mean Camber Shapes of Trimmed Non Coplanar Planforms with Minimum Vortex Drag NASA TN D 8090 June 1976 W H Mason Wing Canard Aerodynamics at Transonic Speeds Fundamental Considerations on Minimum Drag Spanloads AIAA Paper No 82 0097 January 1982 Input Instructions The program assumes the load distribution is constant chordwise until a designated chordwise lo cation XCFW on the first surface and XCFT on the second surface The loading then decreases linearly to the trailing edge This corresponds to a 6 amp 6A series camber distribution the value for the 6A series is usually 0 8 If airfoil polars are
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