********************************************************************* * * * PROGRAM TSFOIL * * SOLVES * * INVISCID FLOW PAST THIN TWO DIMENSIONAL LIFTINGAIRFOIL, * * USING * * TRANSONIC SMALL DISTURBANCE THEORY * * FULLY CONSERVATIVE FINITE DIFFERENCE EQUATIONS * * SUCCESSIVE LINE OVERRELAXATION * * * * WRITTEN BY * * * * EARLL M. MURMAN AND FRANK R. BAILEY * * NASA-AMES RESEARCH CENTER * * MOFFETT FIELD, CALIFORNIA * * AND * * MARGARET L. JOHNSON * * COMPUTER SCIENCES CORPORATION * * MOUNTAIN VIEW, CALIFORNIA * * * ********************************************************************* enter name of input data file bicon.tsfoil2 TEST OF TSFOIL2, 10% thick biconvex airfoil EMACH = 0.82000 POR = 0.00000 IMIN = 1 BCTYPE = 1 AMESH = F DELTA = 0.10000 CLSET = 0.00000 IMAXI = 77 BCFOIL = 2 PHYS = T ALPHA = 0.00000 EPS = 0.20000 JMIN = 1 PSTART = 1 PSAVE = F AK = 0.00000 RIGF = 0.00000 JMAXI = 56 PRTFLO = 1 KUTTA = T GAM = 1.40000 WCIRC = 1.00000 MAXIT = 1500 IPRTER = 10 FCR = T F = 0.00000 CVERGE = 0.00001 NU = 100 SIMDEF = 3 H = 0.00000 DVERGE = 10.0 NL = 75 ICUT = 2 WE = 1.80,1.90,1.95 XIN -1.075000 -0.950000 -0.825000 -0.700000 -0.575000 -0.450000 -0.350000 -0.250000 -0.175000 -0.125000 -0.075000 -0.052500 -0.035000 -0.022500 -0.015000 -0.007500 -0.002500 0.002500 0.007500 0.012500 0.017500 0.022500 0.027500 0.032500 0.037500 0.045000 0.055000 0.065000 0.075000 0.085000 0.097500 0.115000 0.140625 0.171875 0.203125 0.234375 0.265625 0.296875 0.328125 0.359375 0.390625 0.421875 0.453125 0.484375 0.515625 0.546875 0.578125 0.609375 0.640625 0.671875 0.703125 0.734375 0.765625 0.796875 0.828125 0.859375 0.885000 0.900000 0.915000 0.930000 0.945000 0.960000 0.975000 0.990000 1.000000 1.010000 1.025000 1.050000 1.090000 1.150000 1.225000 1.300000 1.400000 1.500000 1.625000 1.750000 1.875000 YIN -5.200000 -4.400000 -3.600000 -3.000000 -2.400000 -1.950000 -1.600000 -1.350000 -1.150000 -0.950000 -0.800000 -0.650000 -0.550000 -0.450000 -0.390000 -0.340000 -0.300000 -0.270000 -0.240000 -0.210000 -0.180000 -0.150000 -0.125000 -0.100000 -0.075000 -0.050000 -0.030000 -0.010000 0.010000 0.030000 0.050000 0.075000 0.100000 0.125000 0.150000 0.180000 0.210000 0.240000 0.270000 0.300000 0.340000 0.390000 0.450000 0.550000 0.650000 0.800000 0.950000 1.150000 1.350000 1.600000 1.950000 2.400000 3.000000 3.600000 4.400000 5.200000 SCALED POR= 0.00000 AIRFOIL GEOMETRY INFORMATION PRINTOUT IN PHYSICAL VARIABLES NORMALIZED BY CHORD LENGTH MAX THICKNESS = 0.09990235 AIRFOIL VOLUME= 0.06666542 MAX CAMBER = 0.00000000 UPPER SURFACE LOWER SURFACE X Y DY/DX Y DY/DX THICKNESS CAMBER 0.00250000 0.00049875 0.19900000 -0.00049875 -0.19900000 0.00049875 0.00000000 0.00750000 0.00148875 0.19700001 -0.00148875 -0.19700001 0.00148875 0.00000000 0.01250000 0.00246875 0.19500001 -0.00246875 -0.19500001 0.00246875 0.00000000 0.01750000 0.00343875 0.19300000 -0.00343875 -0.19300000 0.00343875 0.00000000 0.02250000 0.00439875 0.19100000 -0.00439875 -0.19100000 0.00439875 0.00000000 0.02750000 0.00534875 0.18900000 -0.00534875 -0.18900000 0.00534875 0.00000000 0.03250000 0.00628875 0.18700001 -0.00628875 -0.18700001 0.00628875 0.00000000 0.03750000 0.00721875 0.18500000 -0.00721875 -0.18500000 0.00721875 0.00000000 0.04500000 0.00859500 0.18200000 -0.00859500 -0.18200000 0.00859500 0.00000000 0.05500000 0.01039500 0.17800000 -0.01039500 -0.17800000 0.01039500 0.00000000 0.06500000 0.01215500 0.17400001 -0.01215500 -0.17400001 0.01215500 0.00000000 0.07500000 0.01387500 0.17000000 -0.01387500 -0.17000000 0.01387500 0.00000000 0.08500000 0.01555500 0.16599999 -0.01555500 -0.16599999 0.01555500 0.00000000 0.09750000 0.01759875 0.16100000 -0.01759875 -0.16100000 0.01759875 0.00000000 0.11500000 0.02035500 0.15400000 -0.02035500 -0.15400000 0.02035500 0.00000000 0.14062500 0.02416992 0.14375000 -0.02416992 -0.14375000 0.02416992 0.00000000 0.17187500 0.02846680 0.13125001 -0.02846680 -0.13125001 0.02846680 0.00000000 0.20312500 0.03237305 0.11875000 -0.03237305 -0.11875000 0.03237305 0.00000000 0.23437500 0.03588867 0.10625000 -0.03588867 -0.10625000 0.03588867 0.00000000 0.26562500 0.03901367 0.09375000 -0.03901367 -0.09375000 0.03901367 0.00000000 0.29687500 0.04174805 0.08125000 -0.04174805 -0.08125000 0.04174805 0.00000000 0.32812500 0.04409180 0.06875000 -0.04409180 -0.06875000 0.04409180 0.00000000 0.35937500 0.04604492 0.05625000 -0.04604492 -0.05625000 0.04604492 0.00000000 0.39062500 0.04760742 0.04375000 -0.04760742 -0.04375000 0.04760742 0.00000000 0.42187500 0.04877930 0.03125000 -0.04877930 -0.03125000 0.04877930 0.00000000 0.45312500 0.04956055 0.01875000 -0.04956055 -0.01875000 0.04956055 0.00000000 0.48437500 0.04995117 0.00625000 -0.04995117 -0.00625000 0.04995117 0.00000000 0.51562500 0.04995117 -0.00625000 -0.04995117 0.00625000 0.04995117 0.00000000 0.54687500 0.04956055 -0.01875000 -0.04956055 0.01875000 0.04956055 0.00000000 0.57812500 0.04877930 -0.03125000 -0.04877930 0.03125000 0.04877930 0.00000000 0.60937500 0.04760742 -0.04375000 -0.04760742 0.04375000 0.04760742 0.00000000 0.64062500 0.04604492 -0.05625000 -0.04604492 0.05625000 0.04604492 0.00000000 0.67187500 0.04409180 -0.06875000 -0.04409180 0.06875000 0.04409180 0.00000000 0.70312500 0.04174805 -0.08125000 -0.04174805 0.08125000 0.04174805 0.00000000 0.73437500 0.03901367 -0.09375000 -0.03901367 0.09375000 0.03901367 0.00000000 0.76562500 0.03588867 -0.10625000 -0.03588867 0.10625000 0.03588867 0.00000000 0.79687500 0.03237305 -0.11875000 -0.03237305 0.11875000 0.03237305 0.00000000 0.82812500 0.02846680 -0.13125001 -0.02846680 0.13125001 0.02846680 0.00000000 0.85937500 0.02416992 -0.14375000 -0.02416992 0.14375000 0.02416992 0.00000000 0.88499999 0.02035500 -0.15400000 -0.02035500 0.15400000 0.02035500 0.00000000 0.89999998 0.01800001 -0.16000000 -0.01800001 0.16000000 0.01800001 0.00000000 0.91500002 0.01555500 -0.16600001 -0.01555500 0.16600001 0.01555500 0.00000000 0.93000001 0.01302000 -0.17200001 -0.01302000 0.17200001 0.01302000 0.00000000 0.94499999 0.01039500 -0.17800000 -0.01039500 0.17800000 0.01039500 0.00000000 0.95999998 0.00768000 -0.18400000 -0.00768000 0.18400000 0.00768000 0.00000000 0.97500002 0.00487499 -0.19000001 -0.00487499 0.19000001 0.00487499 0.00000000 0.99000001 0.00197999 -0.19600001 -0.00197999 0.19600001 0.00197999 0.00000000 1.00000000 0.00000000 -0.20000000 0.00000000 0.20000000 0.00000000 0.00000000 WE = 1.8000 EPS = 0.2000 MAXIT FOR THIS MESH = 375 ITERATION CL CM IERR JERR ERROR IRL JRL BIGRL ERCIRC 10 0.00000 0.00000 16 7 0.3718E-01 6 7 0.1027E+03 0.8941E-07 20 0.00000 0.00000 17 7 0.9134E-02 6 13 0.7046E+01 0.5960E-07 30 0.00000 0.00000 15 7 0.6418E-02 6 7 0.2157E+02 0.8941E-07 40 0.00000 0.00000 16 7 0.2039E-02 6 7 0.9436E+01 0.0000E+00 50 0.00000 0.00000 16 8 0.7795E-03 6 7 0.1125E+01 0.2980E-07 60 0.00000 0.00000 17 7 0.1600E-03 6 7 0.1059E+01 0.5960E-07 70 0.00000 0.00000 17 8 0.4918E-04 6 8 0.1884E+00 0.5960E-07 80 0.00000 0.00000 15 8 0.3002E-04 6 8 0.1681E+00 0.2980E-07 90 0.00000 0.00000 4 8 0.8524E-05 6 7 0.7154E-01 0.0000E+00 ........SOLUTION CONVERGED........ FORCE COEFFICIENTS, PRESSURE COEFFICIENT, AND MACH NUMBER (OR SIMILARITY PARAMETER) ON BODY AND DIVIDING STREAM LINE. COARSE MESH CL = 0.000000 CM = 0.000000 CP* = -0.386357 LOWER UPPER Y=0- Y=0+ I X CP M1 CP M1 1 -1.075000 0.008158 0.815771 0.008158 0.815771 B * 2 -0.575000 0.049146 0.794184 0.049146 0.794184 B * 3 -0.175000 0.126231 0.751909 0.126231 0.751909 B * 4 -0.035000 0.236328 0.687031 0.236328 0.687031 B * 5 -0.002500 0.311776 0.638779 0.311776 0.638779 B * AIRFOIL LEADING EDGE AIRFOIL LEADING EDGE 6 0.017500 0.416897 0.564716 0.416896 0.564716 B * 7 0.037500 0.301351 0.645661 0.301351 0.645661 B * 8 0.075000 0.146454 0.740418 0.146454 0.740418 B * 9 0.140625 -0.067469 0.854171 -0.067469 0.854171 B * 10 0.265625 -0.303212 0.964106 -0.303212 0.964106 B * 11 0.390625 -0.523119 1.056392 -0.523119 1.056392 * B 12 0.515625 -0.676711 1.116332 -0.676711 1.116332 * B 13 0.640625 -0.651587 1.106750 -0.651587 1.106750 * B 14 0.765625 -0.356536 0.987276 -0.356536 0.987276 B * 15 0.885000 -0.040424 0.840641 -0.040424 0.840640 B * 16 0.945000 0.134664 0.747138 0.134664 0.747138 B * 17 1.000000 0.358844 0.606736 0.358844 0.606736 B * AIRFOIL TRAILING EDGE AIRFOIL TRAILING EDGE 18 1.090000 0.270231 0.665781 0.270231 0.665781 B * 19 1.400000 0.119694 0.755585 0.119694 0.755585 B * 20 1.875000 0.032174 0.803193 0.032174 0.803193 B * Y(J) J= 1 TO 14 -5.20000 -2.40000 -1.15000 -0.550000 -0.300000 -0.180000 -0.750000E-01 0.750000E-01 0.180000 0.300000 0.550000 1.15000 2.40000 5.20000 WE = 1.9000 EPS = 0.2000 MAXIT FOR THIS MESH = 750 ITERATION CL CM IERR JERR ERROR IRL JRL BIGRL ERCIRC 10 0.00000 0.00000 27 14 0.4580E-02 11 18 0.1927E+02 0.2980E-07 20 0.00000 0.00000 27 15 0.1052E-02 10 14 0.1527E+02 0.5960E-07 30 0.00000 0.00000 27 14 0.1030E-02 10 15 0.8366E+01 0.2980E-07 40 0.00000 0.00000 8 23 0.2390E-03 10 14 0.7248E+01 0.8941E-07 50 0.00000 0.00000 27 14 0.2573E-03 10 15 0.6902E+01 0.5960E-07 60 0.00000 0.00000 7 15 0.1094E-03 10 14 0.4254E+01 0.5960E-07 70 0.00000 0.00000 27 14 0.1259E-03 10 14 0.2769E+01 0.0000E+00 80 0.00000 0.00000 9 14 0.5427E-04 10 14 0.2176E+01 0.2980E-07 90 0.00000 0.00000 7 14 0.3740E-04 10 14 0.1504E+01 0.2980E-07 100 0.00000 0.00000 8 15 0.2246E-04 10 14 0.9048E+00 0.0000E+00 110 0.00000 0.00000 7 14 0.1457E-04 10 14 0.5870E+00 0.2980E-07 120 0.00000 0.00000 8 15 0.1012E-04 10 14 0.4100E+00 0.0000E+00 130 0.00000 0.00000 6 14 0.7680E-05 10 14 0.2923E+00 0.0000E+00 ........SOLUTION CONVERGED........ FORCE COEFFICIENTS, PRESSURE COEFFICIENT, AND MACH NUMBER (OR SIMILARITY PARAMETER) ON BODY AND DIVIDING STREAM LINE. MEDIUM MESH CL = 0.000000 CM = 0.000000 CP* = -0.386357 LOWER UPPER Y=0- Y=0+ I X CP M1 CP M1 1 -1.075000 0.024142 0.807421 0.024142 0.807421 B * 2 -0.825000 0.036048 0.801145 0.036048 0.801145 B * 3 -0.575000 0.053739 0.791728 0.053739 0.791728 B * 4 -0.350000 0.088094 0.773112 0.088094 0.773112 B * 5 -0.175000 0.150165 0.738290 0.150165 0.738290 B * 6 -0.075000 0.237979 0.686012 0.237979 0.686012 B * 7 -0.035000 0.322499 0.631622 0.322499 0.631622 B * 8 -0.015000 0.385859 0.587556 0.385859 0.587556 B * 9 -0.002500 0.436468 0.549827 0.436468 0.549827 B * AIRFOIL LEADING EDGE AIRFOIL LEADING EDGE 10 0.007500 0.535447 0.467315 0.535445 0.467317 B * 11 0.017500 0.449224 0.539901 0.449223 0.539902 B * 12 0.027500 0.376426 0.594323 0.376427 0.594323 B * 13 0.037500 0.302774 0.644726 0.302775 0.644725 B * 14 0.055000 0.219155 0.697548 0.219156 0.697548 B * 15 0.075000 0.135952 0.746407 0.135952 0.746407 B * 16 0.097500 0.044217 0.796811 0.044217 0.796811 B * 17 0.140625 -0.072920 0.856872 -0.072920 0.856872 B * 18 0.203125 -0.198269 0.916797 -0.198269 0.916797 B * 19 0.265625 -0.309976 0.967076 -0.309976 0.967076 B * 20 0.328125 -0.424608 1.016088 -0.424608 1.016088 *B 21 0.390625 -0.524293 1.056863 -0.524293 1.056863 * B 22 0.453125 -0.596022 1.085256 -0.596022 1.085256 * B 23 0.515625 -0.662489 1.110918 -0.662489 1.110918 * B 24 0.578125 -0.720388 1.132799 -0.720388 1.132799 * B 25 0.640625 -0.761729 1.148166 -0.761729 1.148166 * B 26 0.703125 -0.648448 1.105546 -0.648447 1.105546 * B 27 0.765625 -0.312295 0.968092 -0.312294 0.968091 B * 28 0.828125 -0.074936 0.857870 -0.074936 0.857870 B * 29 0.885000 0.003068 0.818412 0.003068 0.818412 B * 30 0.915000 0.094939 0.769350 0.094939 0.769350 B * 31 0.945000 0.196075 0.711438 0.196075 0.711438 B * 32 0.975000 0.330816 0.626014 0.330816 0.626014 B * 33 1.000000 0.508826 0.490872 0.508826 0.490872 B * AIRFOIL TRAILING EDGE AIRFOIL TRAILING EDGE 34 1.025000 0.410687 0.569359 0.410687 0.569359 B * 35 1.090000 0.263521 0.670041 0.263521 0.670041 B * 36 1.225000 0.154474 0.735811 0.154474 0.735811 B * 37 1.400000 0.094726 0.769467 0.094726 0.769467 B * 38 1.625000 0.063137 0.786680 0.063137 0.786680 B * 39 1.875000 0.041775 0.798109 0.041775 0.798109 B * Y(J) J= 1 TO 28 -5.20000 -3.60000 -2.40000 -1.60000 -1.15000 -0.800000 -0.550000 -0.390000 -0.300000 -0.240000 -0.180000 -0.125000 -0.750000E-01-0.300000E-01 0.300000E-01 0.750000E-01 0.125000 0.180000 0.240000 0.300000 0.390000 0.550000 0.800000 1.15000 1.60000 2.40000 3.60000 5.20000 WE = 1.9500 EPS = 0.2000 MAXIT FOR THIS MESH = 1500 ITERATION CL CM IERR JERR ERROR IRL JRL BIGRL ERCIRC 10 0.00000 0.00000 57 28 0.1413E-02 21 25 0.4318E+02 0.0000E+00 20 0.00000 0.00000 53 26 0.4274E-03 17 36 0.1649E+02 0.5960E-07 30 0.00000 0.00000 57 10 0.2081E-03 18 29 0.1426E+02 0.2980E-07 40 0.00000 0.00000 57 7 0.1161E-03 17 29 0.1457E+02 0.2980E-07 50 0.00000 0.00000 66 29 0.7624E-04 17 28 0.1165E+02 0.0000E+00 60 0.00000 0.00000 53 28 0.9471E-04 17 29 0.9816E+01 0.2980E-07 70 0.00000 0.00000 53 28 0.8008E-04 17 29 0.8648E+01 0.5960E-07 80 0.00000 0.00000 72 43 0.5007E-04 17 29 0.7317E+01 0.5960E-07 90 0.00000 0.00000 12 29 0.3432E-04 17 28 0.6482E+01 0.2980E-07 100 0.00000 0.00000 14 29 0.2901E-04 17 29 0.5464E+01 0.2980E-07 110 0.00000 0.00000 16 28 0.2477E-04 17 28 0.4661E+01 0.2980E-07 120 0.00000 0.00000 14 29 0.1968E-04 17 29 0.3649E+01 0.5960E-07 130 0.00000 0.00000 14 29 0.1518E-04 17 29 0.2820E+01 0.5960E-07 140 0.00000 0.00000 16 30 0.1251E-04 17 29 0.2362E+01 0.2980E-07 150 0.00000 0.00000 16 28 0.1038E-04 17 28 0.1961E+01 0.0000E+00 160 0.00000 0.00000 16 27 0.8830E-05 17 29 0.1651E+01 0.5960E-07 ........SOLUTION CONVERGED........ PRINTOUT IN PHYSICAL VARIABLES. DEFINITION OF SIMILARITY PARAMETERS BY KRUPP BOUNDARY CONDITION FOR FREE AIR DIFFERENCE EQUATIONS ARE FULLY CONSERVATIVE. KUTTA CONDITION IS ENFORCED. MACH = 0.8200000 DELTA = 0.1000000 ALPHA = 0.0000000 K = 1.8543713 DOUBLET STRENGTH = 0.8183886 PARAMETERS USED TO TRANSFORM VARIABLES TO TRANSONIC SCALING CPFACT = 0.2500190 CDFACT = 0.0250019 CMFACT = 0.2500190 CLFACT = 0.2500190 YFACT = 2.3791752 VFACT = 5.7295780 FORCE COEFFICIENTS, PRESSURE COEFFICIENT, AND MACH NUMBER (OR SIMILARITY PARAMETER) ON BODY AND DIVIDING STREAM LINE. FINAL MESH CL = 0.000000 CM = 0.000000 CP* = -0.386357 LOWER UPPER Y=0- Y=0+ I X CP M1 CP M1 1 -1.075000 0.028214 0.805281 0.028214 0.805281 B * 2 -0.950000 0.032450 0.803047 0.032450 0.803047 B * 3 -0.825000 0.037587 0.800331 0.037587 0.800331 B * 4 -0.700000 0.044916 0.796439 0.044916 0.796439 B * 5 -0.575000 0.055552 0.790757 0.055552 0.790757 B * 6 -0.450000 0.070677 0.782605 0.070677 0.782605 B * 7 -0.350000 0.091301 0.771352 0.091301 0.771352 B * 8 -0.250000 0.120188 0.755308 0.120188 0.755308 B * 9 -0.175000 0.157107 0.734293 0.157107 0.734293 B * 10 -0.125000 0.201288 0.708324 0.201288 0.708324 B * 11 -0.075000 0.256830 0.674262 0.256830 0.674262 B * 12 -0.052500 0.313246 0.637803 0.313246 0.637803 B * 13 -0.035000 0.366172 0.601594 0.366172 0.601594 B * 14 -0.022500 0.419966 0.562407 0.419966 0.562407 B * 15 -0.015000 0.470668 0.522792 0.470668 0.522792 B * 16 -0.007500 0.522132 0.479242 0.522132 0.479242 B * 17 -0.002500 0.574632 0.430299 0.574632 0.430299 B * AIRFOIL LEADING EDGE AIRFOIL LEADING EDGE 18 0.002500 0.651773 0.346046 0.651775 0.346044 B * 19 0.007500 0.570661 0.434194 0.570662 0.434193 B * 20 0.012500 0.503762 0.495226 0.503762 0.495226 B * 21 0.017500 0.447689 0.541105 0.447688 0.541106 B * 22 0.022500 0.399936 0.577309 0.399934 0.577310 B * 23 0.027500 0.358636 0.606882 0.358633 0.606884 B * 24 0.032500 0.322398 0.631690 0.322395 0.631691 B * 25 0.037500 0.286201 0.655533 0.286201 0.655533 B * 26 0.045000 0.243752 0.682435 0.243751 0.682436 B * 27 0.055000 0.197859 0.710374 0.197860 0.710373 B * 28 0.065000 0.155276 0.735349 0.155276 0.735349 B * 29 0.075000 0.117388 0.756878 0.117388 0.756878 B * 30 0.085000 0.080915 0.777040 0.080915 0.777040 B * 31 0.097500 0.039844 0.799134 0.039844 0.799134 B * 32 0.115000 -0.011856 0.826107 -0.011856 0.826107 B * 33 0.140625 -0.074805 0.857804 -0.074805 0.857804 B * 34 0.171875 -0.141247 0.890037 -0.141247 0.890037 B * 35 0.203125 -0.203563 0.919242 -0.203563 0.919242 B * 36 0.234375 -0.260620 0.945190 -0.260620 0.945191 B * 37 0.265625 -0.313696 0.968705 -0.313696 0.968705 B * 38 0.296875 -0.364494 0.990688 -0.364494 0.990688 B* 39 0.328125 -0.417792 1.013240 -0.417792 1.013240 *B 40 0.359375 -0.467201 1.033707 -0.467201 1.033707 * B 41 0.390625 -0.508559 1.050532 -0.508559 1.050532 * B 42 0.421875 -0.547798 1.066250 -0.547798 1.066250 * B 43 0.453125 -0.585123 1.080989 -0.585123 1.080989 * B 44 0.484375 -0.620731 1.094866 -0.620731 1.094866 * B 45 0.515625 -0.654734 1.107954 -0.654734 1.107954 * B 46 0.546875 -0.687163 1.120294 -0.687163 1.120294 * B 47 0.578125 -0.717962 1.131890 -0.717962 1.131890 * B 48 0.609375 -0.746972 1.142705 -0.746973 1.142705 * B 49 0.640625 -0.773841 1.152630 -0.773841 1.152630 * B 50 0.671875 -0.797609 1.161340 -0.797610 1.161340 * B 51 0.703125 -0.813093 1.166978 -0.813093 1.166979 * B 52 0.734375 -0.650846 1.106465 -0.650849 1.106467 * B 53 0.765625 -0.268434 0.948689 -0.268437 0.948690 B * 54 0.796875 -0.061083 0.850996 -0.061083 0.850996 B * 55 0.828125 -0.057190 0.849054 -0.057189 0.849054 B * 56 0.859375 -0.022353 0.831477 -0.022353 0.831477 B * 57 0.885000 0.022278 0.808400 0.022277 0.808400 B * 58 0.900000 0.062794 0.786864 0.062795 0.786864 B * 59 0.915000 0.102895 0.764953 0.102895 0.764953 B * 60 0.930000 0.149701 0.738557 0.149701 0.738557 B * 61 0.945000 0.205499 0.705800 0.205499 0.705799 B * 62 0.960000 0.274764 0.662889 0.274764 0.662889 B * 63 0.975000 0.366591 0.601299 0.366591 0.601299 B * 64 0.990000 0.487776 0.508728 0.487776 0.508729 B * 65 1.000000 0.633890 0.367303 0.633891 0.367302 B * AIRFOIL TRAILING EDGE AIRFOIL TRAILING EDGE 66 1.010000 0.561496 0.443052 0.561496 0.443052 B * 67 1.025000 0.450600 0.538819 0.450600 0.538819 B * 68 1.050000 0.348559 0.613881 0.348559 0.613881 B * 69 1.090000 0.262234 0.670855 0.262234 0.670855 B * 70 1.150000 0.194924 0.712123 0.194924 0.712123 B * 71 1.225000 0.147731 0.739687 0.147731 0.739687 B * 72 1.300000 0.114835 0.758307 0.114835 0.758307 B * 73 1.400000 0.090279 0.771913 0.090279 0.771913 B * 74 1.500000 0.072189 0.781786 0.072189 0.781786 B * 75 1.625000 0.058856 0.788983 0.058856 0.788983 B * 76 1.750000 0.049771 0.793850 0.049771 0.793850 B * 77 1.875000 0.042679 0.797629 0.042679 0.797629 B * Y(J) J= 1 TO 56 -5.20000 -4.40000 -3.60000 -3.00000 -2.40000 -1.95000 -1.60000 -1.35000 -1.15000 -0.950000 -0.800000 -0.650000 -0.550000 -0.450000 -0.390000 -0.340000 -0.300000 -0.270000 -0.240000 -0.210000 -0.180000 -0.150000 -0.125000 -0.100000 -0.750000E-01-0.500000E-01-0.300000E-01-0.100000E-01 0.100000E-01 0.300000E-01 0.500000E-01 0.750000E-01 0.100000 0.125000 0.150000 0.180000 0.210000 0.240000 0.270000 0.300000 0.340000 0.390000 0.450000 0.550000 0.650000 0.800000 0.950000 1.15000 1.35000 1.60000 1.95000 2.40000 3.00000 3.60000 4.40000 5.20000 FLOW AT EACH GRID POINT. P PARABOLIC H HYPERBOLIC S SHOCK - ELLIPTIC 56 --------------------------------------------------------------------------- 55 --------------------------------------------------------------------------- 54 --------------------------------------------------------------------------- 53 --------------------------------------------------------------------------- 52 --------------------------------------------------------------------------- 51 --------------------------------------------------------------------------- 50 --------------------------------------------------------------------------- 49 --------------------------------------------------------------------------- 48 --------------------------------------------------------------------------- 47 --------------------------------------------------------------------------- 46 --------------------------------------------------------------------------- 45 --------------------------------------------------------------------------- 44 --------------------------------------------------------------------------- 43 ----------------------------------------------PS--------------------------- 42 --------------------------------------------PHHHHS------------------------- 41 -------------------------------------------PHHHHHS------------------------- 40 ------------------------------------------PHHHHHHHS------------------------ 39 -----------------------------------------PHHHHHHHHS------------------------ 38 -----------------------------------------PHHHHHHHHS------------------------ 37 ----------------------------------------PHHHHHHHHHS------------------------ 36 ----------------------------------------PHHHHHHHHHS------------------------ 35 ---------------------------------------PHHHHHHHHHHHS----------------------- 34 ---------------------------------------PHHHHHHHHHHHS----------------------- 33 --------------------------------------PHHHHHHHHHHHHS----------------------- 32 --------------------------------------PHHHHHHHHHHHHS----------------------- 31 --------------------------------------PHHHHHHHHHHHHS----------------------- 30 -------------------------------------PHHHHHHHHHHHHHS----------------------- 29 -------------------------------------PHHHHHHHHHHHHHS----------------------- 28 -------------------------------------PHHHHHHHHHHHHHS----------------------- 27 -------------------------------------PHHHHHHHHHHHHHS----------------------- 26 --------------------------------------PHHHHHHHHHHHHS----------------------- 25 --------------------------------------PHHHHHHHHHHHHS----------------------- 24 --------------------------------------PHHHHHHHHHHHHS----------------------- 23 ---------------------------------------PHHHHHHHHHHHS----------------------- 22 ---------------------------------------PHHHHHHHHHHHS----------------------- 21 ----------------------------------------PHHHHHHHHHS------------------------ 20 ----------------------------------------PHHHHHHHHHS------------------------ 19 -----------------------------------------PHHHHHHHHS------------------------ 18 -----------------------------------------PHHHHHHHHS------------------------ 17 ------------------------------------------PHHHHHHHS------------------------ 16 -------------------------------------------PHHHHHS------------------------- 15 --------------------------------------------PHHHHS------------------------- 14 ----------------------------------------------PS--------------------------- 13 --------------------------------------------------------------------------- 12 --------------------------------------------------------------------------- 11 --------------------------------------------------------------------------- 10 --------------------------------------------------------------------------- 9 --------------------------------------------------------------------------- 8 --------------------------------------------------------------------------- 7 --------------------------------------------------------------------------- 6 --------------------------------------------------------------------------- 5 --------------------------------------------------------------------------- 4 --------------------------------------------------------------------------- 3 --------------------------------------------------------------------------- 2 --------------------------------------------------------------------------- 1 --------------------------------------------------------------------------- MACH NO. MAP. ROUNDED TO NEAREST .1 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888899999999999999999888888888888888888888888 8888888888888888888888888888888889999999999999999999998888888888888888888888 8888888888888888888888888888888889999999999999999999999888888888888888888888 8888888888888888888888888888888899999999999999999999999888888888888888888888 8888888888888888888888888888888899999999999999999999999888888888888888888888 888888888888888888888888888888889999999999******9999998888888888888888888888 8888888888888888888888888888888899999999*********999998888888888888888888888 8888888888888888888888888888888889999999**********99998888888888888888888888 888888888888888888888888888888888999999***********99988888888888888888888888 88888888888888888888888888888888899999********11**99988888888888888888888888 88888888888888888888888888888888899999*******1111*99988888888888888888888888 88888888888888888888888888888888999999*******1111**9988888888888887788888888 8888888877777888888888888888888899999*******111111*9988888888887777778888888 8888888877777777777788888888888899999******1111111*9988888888777777778888888 8888888777777777777777778888888899999******1111111*9988888877777777777888888 888888877777777777777777778888889999******11111111*9888888877777777777888888 888888877777777777777777777888889999******11111111*9888888777777777777888888 888888877777777777777777777888889999*****11111111119888888777777777777888888 88888887777666666666777777788888999*****111111111219888888777766666777888888 88888887776666666666667777788888999*****111111112219888888777666666777888888 +88888887776665544556666777788889999*****111111112219888888777654456777888888 -88888887776665544556666777788889999*****111111112219888888777654456777888888 88888887776666666666667777788888999*****111111112219888888777666666777888888 88888887777666666666777777788888999*****111111111219888888777766666777888888 888888877777777777777777777888889999*****11111111119888888777777777777888888 888888877777777777777777777888889999******11111111*9888888777777777777888888 888888877777777777777777778888889999******11111111*9888888877777777777888888 8888888777777777777777778888888899999******1111111*9988888877777777777888888 8888888877777777777788888888888899999******1111111*9988888888777777778888888 8888888877777888888888888888888899999*******111111*9988888888887777778888888 88888888888888888888888888888888999999*******1111**9988888888888887788888888 88888888888888888888888888888888899999*******1111*99988888888888888888888888 88888888888888888888888888888888899999********11**99988888888888888888888888 888888888888888888888888888888888999999***********99988888888888888888888888 8888888888888888888888888888888889999999**********99998888888888888888888888 8888888888888888888888888888888899999999*********999998888888888888888888888 888888888888888888888888888888889999999999******9999998888888888888888888888 8888888888888888888888888888888899999999999999999999999888888888888888888888 8888888888888888888888888888888899999999999999999999999888888888888888888888 8888888888888888888888888888888889999999999999999999999888888888888888888888 8888888888888888888888888888888889999999999999999999998888888888888888888888 8888888888888888888888888888888888899999999999999999888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 8888888888888888888888888888888888888888888888888888888888888888888888888888 L T PRINTER PLOT OF CP ON BODY AND DIVIDING STREAMLINE U FOR CP(UPPER) L FOR CP(LOWER) B FOR CP(UPPER)=CP(LOWER) --- FOR CP SONIC BB B B B B B B B B B B B B - - - - - - - - - - ------------ --- ---- ---- --- ---- ----------- - - - - - - - - - B B B B B B B BB B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B B BB B B B B B BB B B SONIC LINE COORDINATES Y XSONIC 0.45000 0.57972 0.62315 0.39000 0.51815 0.67286 0.34000 0.48472 0.69062 0.30000 0.46154 0.70452 0.27000 0.44490 0.71227 0.24000 0.42886 0.71908 0.21000 0.41329 0.72524 0.18000 0.39795 0.73111 0.15000 0.38283 0.73702 0.12500 0.37043 0.74137 0.10000 0.35765 0.74522 0.07500 0.34588 0.74859 0.05000 0.33376 0.75150 0.03000 0.32403 0.75348 0.01000 0.31445 0.75517 BODY LOCATION -0.01000 0.31445 0.75517 -0.03000 0.32403 0.75348 -0.05000 0.33376 0.75150 -0.07500 0.34588 0.74859 -0.10000 0.35765 0.74522 -0.12500 0.37043 0.74137 -0.15000 0.38283 0.73702 -0.18000 0.39795 0.73111 -0.21000 0.41329 0.72524 -0.24000 0.42886 0.71908 -0.27000 0.44490 0.71227 -0.30000 0.46154 0.70452 -0.34000 0.48472 0.69062 -0.39000 0.51815 0.67286 -0.45000 0.57973 0.62315 SONIC LINE PLOT Y VS X * FOR SONIC POINTS + + + + + + + + + + + ....................................................................................................... 1.5 + . . + . . . . . . . . . . . . . . . . . . 1.0 + . . + . . . . . . . . . . . . . . . . . . 0.5 + . . + . * * . . * * . . * * . . * * . . ** ** . . ** * . . * * . . ** * . . ** * . 0.0 + . ------------*--BODY SLIT-----*--------- . + . * * . . * * . . * ** . . ** * . . ** ** . . * * . . * * . . * * . . * * . -0.5 + . . + . . . . . . . . . . . . . . . . . . -1.0 + . . + ....................................................................................................... + + + + + + + + + + + -0.75 -0.50 -0.25 0.00 0.25 0.50 0.75 1.00 1.25 1.50 1.75 ( 60 PTS) INVISCID WAKE PROFILES FOR INDIVIDUAL SHOCK WAVES WITHIN MOMENTUM CONTOUR SHOCK 1 WAVE DRAG FOR THIS SHOCK= 0.002283 Y CD(Y) PO/POINF 0.00000000 0.01429778 0.99047601 0.01000000 0.01406782 0.99062914 0.03000000 0.01361532 0.99093056 0.05000000 0.01282441 0.99145740 0.07500001 0.01150451 0.99233663 0.10000000 0.00997095 0.99335819 0.12500000 0.00835267 0.99443614 0.15000002 0.00675586 0.99549979 0.18000001 0.00440938 0.99706286 0.20999999 0.00369978 0.99753553 0.23999999 0.00294424 0.99803877 0.27000001 0.00223582 0.99851066 0.30000004 0.00162573 0.99891704 0.34000003 0.00064108 0.99957299 0.38999999 0.00042881 0.99971437 0.45000002 0.00000676 0.99999547 SHOCK 2 WAVE DRAG FOR THIS SHOCK= 0.002283 Y CD(Y) PO/POINF 0.00000000 0.01429776 0.99047601 -0.01000000 0.01406778 0.99062920 -0.03000000 0.01361526 0.99093062 -0.05000000 0.01282437 0.99145746 -0.07500001 0.01150445 0.99233669 -0.10000000 0.00997089 0.99335819 -0.12500000 0.00835263 0.99443614 -0.15000002 0.00675581 0.99549985 -0.18000001 0.00440939 0.99706280 -0.20999999 0.00369979 0.99753553 -0.23999999 0.00294425 0.99803877 -0.27000001 0.00223582 0.99851066 -0.30000004 0.00162572 0.99891710 -0.34000003 0.00064109 0.99957293 -0.38999999 0.00042881 0.99971437 -0.45000002 0.00000676 0.99999547 CALCULATION OF DRAG COEFFICIENT BY MOMENTUM INTEGRAL METHOD BOUNDARIES OF CONTOUR USED CONTRIBUTION TO CD UPSTREAM X = -0.175000 CDUP = 0.000038 DOWNSTREAM X = 1.225000 CDDOWN = -0.000169 TOP Y = 4.400000 CDTOP = 0.000000 BOTTOM Y = -4.400000 CDBOT = 0.000000 TOTAL CONTRIBUTIONS AROUND CONTOUR = -0.000131 THERE ARE 2 SHOCKS INSIDE CONTOUR. TOTAL CDWAVE = 0.004566 NOTE - ALL SHOCKS CONTAINED WITHIN CONTOUR CDWAVE EQUALS TOTAL WAVE DRAG DRAG CALCULATED FROM MOMENTUM INTEGRAL CD = 0.004435 enter name of input data file *** Run time error - File ' ' not found