Archive for February, 2009

Aircraft range calculator

You can download it from here (it is in the katix.org gforge svn):
RangeCalculator.ods

Some calculated results:

Target range = 1500 nm
Fuel consumption = 31.5 liters/h (2 x Rotax 912ULS, with economy cruise power)


[kts] [h]  [l] [kg]
Speed Endurance required Fuel liters Fuel weight
100 15 472.47 335.45
110 13.64 429.52 304.96
120 12.5 393.73 279.55
130 11.54 363.44 258.04
140 10.71 337.48 239.61
150 10 314.98 223.64
160 9.38 295.29 209.66
170 8.82 277.92 197.33
180 8.33 262.48 186.36
190 7.89 248.67 176.56
200 7.5 236.24 167.73
210 7.14 224.99 159.74
220 6.82 214.76 152.48
230 6.52 205.42 145.85
240 6.25 196.86 139.77
250 6 188.99 134.18
260 5.77 181.72 129.02
270 5.56 174.99 124.24
280 5.36 168.74 119.81
290 5.17 162.92 115.67
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Aircraft range calculator

You can download it from here (it is in the katix.org gforge svn):
RangeCalculator.ods

Some calculated results:

Target range = 1500 nm
Fuel consumption = 31.5 liters/h (2 x Rotax 912ULS, with economy cruise power)


[kts] [h]  [l] [kg]
Speed Endurance required Fuel liters Fuel weight
100 15 472.47 335.45
110 13.64 429.52 304.96
120 12.5 393.73 279.55
130 11.54 363.44 258.04
140 10.71 337.48 239.61
150 10 314.98 223.64
160 9.38 295.29 209.66
170 8.82 277.92 197.33
180 8.33 262.48 186.36
190 7.89 248.67 176.56
200 7.5 236.24 167.73
210 7.14 224.99 159.74
220 6.82 214.76 152.48
230 6.52 205.42 145.85
240 6.25 196.86 139.77
250 6 188.99 134.18
260 5.77 181.72 129.02
270 5.56 174.99 124.24
280 5.36 168.74 119.81
290 5.17 162.92 115.67

KS20 airfoil simulation

KS20:

Cl – Cd(low reynolds numbers also included, plus also flapped version (+10deg and +20 deg)

L/D vs. alpha:

Cm vs. Alpha:

Cl – alpha:

Printable profile picture of KS20 (black on white background):


QFLR5_v.0001

Calculated polar for: KS20

1 1 Reynolds number fixed Mach number fixed

xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.270 Re = 5.000 e 6 Ncrit = 9.000

alpha CL CD CDp CM Top Xtr Bot Xtr Cpmin Chinge XCp
------- -------- --------- --------- -------- ------- ------- -------- --------- ---------
-2.500 0.0172 0.00639 0.00156 -0.0630 0.5734 0.1263 -0.9628 0.0000 3.9974
-2.000 0.0779 0.00594 0.00135 -0.0638 0.5687 0.2065 -0.7727 0.0000 1.0832
-1.500 0.1376 0.00529 0.00111 -0.0647 0.5613 0.3515 -0.6885 0.0000 0.7246
-1.000 0.1985 0.00466 0.00091 -0.0657 0.5560 0.4943 -0.7241 0.0000 0.5818
-0.500 0.2604 0.00434 0.00082 -0.0667 0.5474 0.5909 -0.7638 0.0000 0.5048
0.000 0.3235 0.00430 0.00084 -0.0677 0.5380 0.6224 -0.8063 0.0000 0.4565
0.500 0.3863 0.00438 0.00088 -0.0686 0.5260 0.6346 -0.8489 0.0000 0.4238
1.000 0.4491 0.00445 0.00094 -0.0696 0.5127 0.6478 -0.8954 0.0000 0.4002
1.500 0.5115 0.00461 0.00102 -0.0704 0.4945 0.6510 -0.9563 0.0000 0.3823
2.000 0.5732 0.00475 0.00111 -0.0712 0.4723 0.6587 -1.0304 0.0000 0.3682
2.500 0.6344 0.00497 0.00124 -0.0719 0.4472 0.6628 -1.1188 0.0000 0.3567
3.000 0.6941 0.00529 0.00142 -0.0723 0.4114 0.6656 -1.2143 0.0000 0.3470
3.500 0.7516 0.00577 0.00168 -0.0724 0.3616 0.6675 -1.3096 0.0000 0.3385
4.000 0.8098 0.00619 0.00194 -0.0726 0.3230 0.6691 -1.4113 0.0000 0.3313
4.500 0.8664 0.00670 0.00226 -0.0725 0.2804 0.6702 -1.5275 0.0000 0.3247
5.000 0.9229 0.00719 0.00260 -0.0724 0.2437 0.6718 -1.6463 0.0000 0.3189

KS20.dat Airfoil file for QFLR5, XFLR5 or Xfoil

KS20 airfoil simulation

KS20:

Cl – Cd(low reynolds numbers also included, plus also flapped version (+10deg and +20 deg)

L/D vs. alpha:

Cm vs. Alpha:

Cl – alpha:

Printable profile picture of KS20 (black on white background):


QFLR5_v.0001

Calculated polar for: KS20

1 1 Reynolds number fixed Mach number fixed

xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.270 Re = 5.000 e 6 Ncrit = 9.000

alpha CL CD CDp CM Top Xtr Bot Xtr Cpmin Chinge XCp
------- -------- --------- --------- -------- ------- ------- -------- --------- ---------
-2.500 0.0172 0.00639 0.00156 -0.0630 0.5734 0.1263 -0.9628 0.0000 3.9974
-2.000 0.0779 0.00594 0.00135 -0.0638 0.5687 0.2065 -0.7727 0.0000 1.0832
-1.500 0.1376 0.00529 0.00111 -0.0647 0.5613 0.3515 -0.6885 0.0000 0.7246
-1.000 0.1985 0.00466 0.00091 -0.0657 0.5560 0.4943 -0.7241 0.0000 0.5818
-0.500 0.2604 0.00434 0.00082 -0.0667 0.5474 0.5909 -0.7638 0.0000 0.5048
0.000 0.3235 0.00430 0.00084 -0.0677 0.5380 0.6224 -0.8063 0.0000 0.4565
0.500 0.3863 0.00438 0.00088 -0.0686 0.5260 0.6346 -0.8489 0.0000 0.4238
1.000 0.4491 0.00445 0.00094 -0.0696 0.5127 0.6478 -0.8954 0.0000 0.4002
1.500 0.5115 0.00461 0.00102 -0.0704 0.4945 0.6510 -0.9563 0.0000 0.3823
2.000 0.5732 0.00475 0.00111 -0.0712 0.4723 0.6587 -1.0304 0.0000 0.3682
2.500 0.6344 0.00497 0.00124 -0.0719 0.4472 0.6628 -1.1188 0.0000 0.3567
3.000 0.6941 0.00529 0.00142 -0.0723 0.4114 0.6656 -1.2143 0.0000 0.3470
3.500 0.7516 0.00577 0.00168 -0.0724 0.3616 0.6675 -1.3096 0.0000 0.3385
4.000 0.8098 0.00619 0.00194 -0.0726 0.3230 0.6691 -1.4113 0.0000 0.3313
4.500 0.8664 0.00670 0.00226 -0.0725 0.2804 0.6702 -1.5275 0.0000 0.3247
5.000 0.9229 0.00719 0.00260 -0.0724 0.2437 0.6718 -1.6463 0.0000 0.3189

KS20.dat Airfoil file for QFLR5, XFLR5 or Xfoil

QFLR5 Experiment: KSLaminar1 airfoil

KSLaminar1.dat
Simulated polar

KSLaminar2.dat

QFLR5 Experiment: KSLaminar1 airfoil

KSLaminar1.dat
Simulated polar

KSLaminar2.dat

Simulations: Althaus AH 94-145 vs. AH 95-160

AH-94-145:

AH-95-160:

AH-94-145-vs-95-160:

AH 94-145 simulated ailerons, 70% chord: neutral, -10, +10 deg, mach 0.27 Re 4.5M cruise: