MissileDatCom is a public domain program that computes the lift and drag coefficients of a parametrically defined missile for a set of user-specified Mach numbers.
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# # Missile DATCOM File Wrapper # # @author: Phoenix Integration # @version: Training # @description: Missile DATCOM, 1 alpha at 1 Mach # RunCommands { generate inputFile run "mdc" parse outputFile } RowFieldInputFile inputFile { templateFile: MDC1.template initializationFile: MDC1.template fileToGenerate: for005.dat setDelimiters "= ," # name type row field #----------------------------------------------------- markAsBeginning "FLTCON" setGroup "FLTCON" variable: MACH double 2 2 variable: ALT double 3 2 units="ft" variable: ALPHA double 5 2 units="degrees" variable: BETA double 6 2 units="degrees" markAsBeginning "REFQ" setGroup "REFQ" variable: XCG double 1 3 units="ft" variable: LREF double 1 5 units="ft^2" variable: SREF double 1 7 units="ft^2" variable: SCALE double 1 9 variable: ROUGH double 2 2 variable: BLAYER string 2 4 markAsBeginning "AXIBOD" setGroup "AXIBODY" variable: TNOSE string 1 3 enumValues="CONE,OGIVE,POWER,HAACK,KARMAN" variable: LNOSE double 1 5 units="ft" lowerBound=1 variable: DNOSE double 1 7 units="ft" lowerBound=0.1 variable: LCENTR double 2 2 units="ft" lowerBound=1 variable: DCENTR double 2 4 units="ft" lowerBound=0.1 variable: TAFT double 3 2 units="ft" variable: LAFT double 3 4 units="ft" lowerBound=1 variable: DAFT double 3 6 units="ft" lowerBound=0.1 variable: DEXIT double 3 8 units="ft" lowerBound=0.1 markAsBeginning "FINSET1" setGroup "FINSET1" variable: SECTYP string 1 3 variable: ZUPPER double 2 2 variable: LMAXU1 double 3 2 variable: LMAXU2 double 3 3 variable: LFLATU1 double 4 2 variable: LFLATU2 double 4 3 variable: SSPAN1 double 5 2 units="ft" variable: SSPAN2 double 5 3 units="ft" variable: CHORD1 double 6 2 units="ft" variable: CHORD2 double 6 3 units="ft" variable: SWEEP double 7 2 units="degrees" variable: XLE double 8 2 units="ft" variable: STA double 9 2 variable: NPANEL double 10 2 markAsBeginning "TRIM" setGroup "TRIM" variable: set double 1 3 variable: PANL1 string 1 5 variable: PANL2 string 1 7 variable: PANL3 string 1 9 variable: PANL4 string 1 11 } RowFieldOutputFile outputFile { fileToParse: for006.dat setDelimiters "= " markAsBeginning "STATIC AERODYNAMIC COEFFICIENTS TRIMMED IN PITCH" markAsBeginning "DELTA" setGroup "OUTPUT" # name type row field options #----------------------------------------------------- variable: DELTA double 3 2 ignoreConversionErrors=true variable: CL double 3 3 ignoreConversionErrors=true variable: CD double 3 4 ignoreConversionErrors=true variable: CN double 3 5 ignoreConversionErrors=true variable: CA double 3 6 ignoreConversionErrors=true }
*------------------------------------------------------------------------ CASEID HYPMIS * Use newtonian imact theory. HYPER * Dimensions are in inches DIM IN * *---------------------------------------------------------------------------- * $FLTCON NMACH=1.0, MACH=1.0, ALT=0.0, NALPHA=2.0, ALPHA(1)=10.0,0.0, BETA=0.0, $END * *---------------------------------------------------------------------------- * $REFQ XCG=81.0,LREF=21.0,SREF =346.4,SCALE=1.0, ROUGH=2.5E-4,BLAYER=TURB, $END * *---------------------------------------------------------------------------- * * BODY $AXIBOD TNOSE=CONE,LNOSE=20.0,DNOSE=20.0, LCENTR=140.0,DCENTR=20.0, TAFT=0.0,LAFT=20.0,DAFT=20.0,DEXIT=16.0, $END * *---------------------------------------------------------------------------- * * TAILS WITH FOUR PANELS $FINSET1 SECTYP=ARC, ZUPPER=0.04, LMAXU=0.4,0.4, LFLATU=0.0,0.0, SSPAN=10.0,15.0, CHORD=20.0,10.8, SWEEP=30.0, XLE=0.0, STA=0.0, NPANEL=4.0, PHIF=45.0,135.0,225.0,315.0, $END $TRIM SET=1.0,PANL1=.TRUE.,PANL2=.TRUE.,PANL3=.TRUE.,PANL4=.TRUE., $END PART SAVE NEXT CASE
1 ***** THE USAF AUTOMATED MISSILE DATCOM * REV 5/97 *****
AERODYNAMIC METHODS FOR MISSILE CONFIGURATIONS
CONERR - INPUT ERROR CHECKING
ERROR CODES - N* DENOTES THE NUMBER OF OCCURENCES OF EACH ERROR
A - UNKNOWN VARIABLE NAME
B - MISSING EQUAL SIGN FOLLOWING VARIABLE NAME
C - NON-ARRAY VARIABLE HAS AN ARRAY ELEMENTDESIGNATION - (N)
D - NON-ARRAY VARIABLE HAS MULTIPLE VALUES ASSIGNED
E - ASSIGNED VALUES EXCEED ARRAY DIMENSION
F - SYNTAX ERROR
************************* INPUT DATA CARDS *************************
1 *
2 *23456789*123456789*123456789*123456789*123456789*123456789*123456789*123456789*
3 *
4 *------------------------------------------------------------------------
5 CASEID HYPMIS
6 * Use newtonian imact theory.
7 HYPER
8 * Dimensions are in inches
9 DIM IN
10 *
11 *----------------------------------------------------------------------------
12 *
13 $FLTCON NMACH=1.0,
14 MACH=1.0,
15 ALT=0.0,
16 NALPHA=2.0,
17 ALPHA(1)=10.0,0.0,
18 BETA=0.0,
19 $END
20 *
21 *----------------------------------------------------------------------------
22 *
23 $REFQ XCG=81.0,LREF=21.0,SREF =346.4,SCALE=1.0,
24 ROUGH=2.5E-4,BLAYER=TURB,
** SUBSTITUTING NUMERIC FOR NAME TURB
25 $END
26 *
27 *----------------------------------------------------------------------------
28 *
29 * BODY
30 $AXIBOD TNOSE=CONE,LNOSE=20.0,DNOSE=20.0,
** SUBSTITUTING NUMERIC FOR NAME CONE
31 LCENTR=140.0,DCENTR=20.0,
32 TAFT=0.0,LAFT=20.0,DAFT=20.0,DEXIT=16.0,
33 $END
34 *
35 *----------------------------------------------------------------------------
36 *
37 * TAILS WITH FOUR PANELS
38 $FINSET1 SECTYP=ARC,
** SUBSTITUTING NUMERIC FOR NAME ARC
39 ZUPPER=0.04,
40 LMAXU=0.4,0.4,
41 LFLATU=0.0,0.0,
42 SSPAN=10.0,15.0,
43 CHORD=20.0,10.8,
44 SWEEP=30.0,
45 XLE=0.0,
46 STA=0.0,
47 NPANEL=4.0,
48 PHIF=45.0,135.0,225.0,315.0,
49 $END
50 $TRIM SET=1.0,PANL1=.TRUE.,PANL2=.TRUE.,PANL3=.TRUE.,PANL4=.TRUE.,
51 $END
52 PART
53 ** BLANK CARD - IGNORED
54 SAVE
55 NEXT CASE
1 ***** THE USAF AUTOMATED MISSILE DATCOM * REV 5/97 ***** CASE 1
AERODYNAMIC METHODS FOR MISSILE CONFIGURATIONS PAGE 1
CASE INPUTS
FOLLOWING ARE THE CARDS INPUT FOR THIS CASE
*
*23456789*123456789*123456789*123456789*123456789*123456789*123456789*123456789*
*
*------------------------------------------------------------------------
CASEID HYPMIS
* Use newtonian imact theory.
HYPER
* Dimensions are in inches
DIM IN
*
*----------------------------------------------------------------------------
*
$FLTCON NMACH=1.0,
MACH=1.0,
ALT=0.0,
NALPHA=2.0,
ALPHA(1)=10.0,0.0,
BETA=0.0,
$END
*
*----------------------------------------------------------------------------
*
$REFQ XCG=81.0,LREF=21.0,SREF =346.4,SCALE=1.0,
ROUGH=2.5E-4,BLAYER=0.,
$END
*
*----------------------------------------------------------------------------
*
* BODY
$AXIBOD TNOSE=0.,LNOSE=20.0,DNOSE=20.0,
LCENTR=140.0,DCENTR=20.0,
TAFT=0.0,LAFT=20.0,DAFT=20.0,DEXIT=16.0,
$END
*
*----------------------------------------------------------------------------
*
* TAILS WITH FOUR PANELS
$FINSET1 SECTYP=2.,
ZUPPER=0.04,
LMAXU=0.4,0.4,
LFLATU=0.0,0.0,
SSPAN=10.0,15.0,
CHORD=20.0,10.8,
SWEEP=30.0,
XLE=0.0,
STA=0.0,
NPANEL=4.0,
PHIF=45.0,135.0,225.0,315.0,
$END
$TRIM SET=1.0,PANL1=.TRUE.,PANL2=.TRUE.,PANL3=.TRUE.,PANL4=.TRUE.,
$END
PART
SAVE
NEXT CASE
THE BOUNDARY LAYER IS ASSUMED TO BE TURBULENT
THE INPUT UNITS ARE IN INCHES, THE SCALE FACTOR IS 1.0000
1 ***** THE USAF AUTOMATED MISSILE DATCOM * REV 5/97 ***** CASE 1
AERODYNAMIC METHODS FOR MISSILE CONFIGURATIONS PAGE 2
HYPMIS
AXISYMMETRIC BODY DEFINITION
NOSE CENTERBODY AFT BODY TOTAL
SHAPE CONIC CYLINDER CONIC
LENGTH 20.000 140.000 20.000 180.000 IN
FINENESS RATIO 1.000 7.000 1.000 9.000
PLANFORM AREA 200.000 2800.000 400.000 3400.000 IN**2
AREA CENTROID 13.333 90.000 170.000 94.902 IN
WETTED AREA 702.481 8796.466 1256.636 10755.580 IN**2
VOLUME 2094.395 43982.320 6283.190 52359.910 IN**3
VOL. CENTROID 15.000 90.000 170.000 96.600 IN
MOLD LINE CONTOUR
LONGITUDINAL STATIONS .0000 2.0000 4.0000 6.0000 8.0000
10.0000 12.0000 14.0000 16.0000 18.0000 20.0000* 34.0000
48.0000 62.0000 76.0000 90.0000 104.0000 118.0000 132.0000
146.0000 160.0000* 162.0000 164.0000 166.0000 168.0000 170.0000
172.0000 174.0000 176.0000 178.0000 180.0000*
BODY RADII .0000 1.0000 2.0000 3.0000 4.0000
5.0000 6.0000 7.0000 8.0000 9.0000 10.0000* 10.0000
10.0000 10.0000 10.0000 10.0000 10.0000 10.0000 10.0000
10.0000 10.0000* 10.0000 10.0000 10.0000 10.0000 10.0000
10.0000 10.0000 10.0000 10.0000 10.0000*
NOTE - * INDICATES SLOPE DISCONTINUOUS POINTS
1 ***** THE USAF AUTOMATED MISSILE DATCOM * REV 5/97 ***** CASE 1
AERODYNAMIC METHODS FOR MISSILE CONFIGURATIONS PAGE 3
HYPMIS
FIN SET NUMBER 1 AIRFOIL SECTION
NACA S-2-50.0-06.6-00.0
X/C X-UPPER Y-UPPER X-LOWER Y-LOWER MEAN LINE THICKNESS
.00000 .00000 .00000 .00000 .00000 .00000 .00000
.00100 .00100 .00013 .00100 -.00013 .00000 .00026
.00200 .00200 .00026 .00200 -.00026 .00000 .00053
.00300 .00300 .00040 .00300 -.00040 .00000 .00079
.00400 .00400 .00053 .00400 -.00053 .00000 .00106
.00500 .00500 .00066 .00500 -.00066 .00000 .00132
.00600 .00600 .00079 .00600 -.00079 .00000 .00158
.00800 .00800 .00105 .00800 -.00105 .00000 .00210
.01000 .01000 .00131 .01000 -.00131 .00000 .00262
.02000 .02000 .00260 .02000 -.00260 .00000 .00520
.03000 .03000 .00386 .03000 -.00386 .00000 .00771
.04000 .04000 .00509 .04000 -.00509 .00000 .01018
.05000 .05000 .00629 .05000 -.00629 .00000 .01258
.06000 .06000 .00747 .06000 -.00747 .00000 .01494
.08000 .08000 .00975 .08000 -.00975 .00000 .01949
.10000 .10000 .01191 .10000 -.01191 .00000 .02383
.12000 .12000 .01397 .12000 -.01397 .00000 .02795
.14000 .14000 .01593 .14000 -.01593 .00000 .03186
.16000 .16000 .01778 .16000 -.01778 .00000 .03555
.18000 .18000 .01952 .18000 -.01952 .00000 .03904
.20000 .20000 .02115 .20000 -.02115 .00000 .04231
.22000 .22000 .02268 .22000 -.02268 .00000 .04536
.24000 .24000 .02411 .24000 -.02411 .00000 .04821
.26000 .26000 .02542 .26000 -.02542 .00000 .05084
.28000 .28000 .02663 .28000 -.02663 .00000 .05327
.30000 .30000 .02774 .30000 -.02774 .00000 .05548
.32000 .32000 .02874 .32000 -.02874 .00000 .05748
.34000 .34000 .02963 .34000 -.02963 .00000 .05927
.36000 .36000 .03042 .36000 -.03042 .00000 .06085
.38000 .38000 .03111 .38000 -.03111 .00000 .06221
.40000 .40000 .03169 .40000 -.03169 .00000 .06337
.42000 .42000 .03216 .42000 -.03216 .00000 .06432
.45000 .45000 .03267 .45000 -.03267 .00000 .06534
.50000 .50000 .03300 .50000 -.03300 .00000 .06600
.55000 .55000 .03267 .55000 -.03267 .00000 .06534
.60000 .60000 .03169 .60000 -.03169 .00000 .06337
.65000 .65000 .03004 .65000 -.03004 .00000 .06008
.70000 .70000 .02774 .70000 -.02774 .00000 .05548
.75000 .75000 .02478 .75000 -.02478 .00000 .04955
.80000 .80000 .02115 .80000 -.02115 .00000 .04231
.82000 .82000 .01952 .82000 -.01952 .00000 .03904
.84000 .84000 .01778 .84000 -.01778 .00000 .03555
.86000 .86000 .01593 .86000 -.01593 .00000 .03186
.88000 .88000 .01397 .88000 -.01397 .00000 .02795
.90000 .90000 .01191 .90000 -.01191 .00000 .02383
.92000 .92000 .00975 .92000 -.00975 .00000 .01949
.94000 .94000 .00747 .94000 -.00747 .00000 .01494
.96000 .96000 .00509 .96000 -.00509 .00000 .01018
.98000 .98000 .00260 .98000 -.00260 .00000 .00520
1.00000 1.00000 .00000 1.00000 .00000 .00000 .00000
1 ***** THE USAF AUTOMATED MISSILE DATCOM * REV 5/97 ***** CASE 1
AERODYNAMIC METHODS FOR MISSILE CONFIGURATIONS PAGE 4
HYPMIS
GEOMETRIC RESULTS FOR FIN SETS
FIN SET NUMBER 1
(DATA FOR ONE PANEL ONLY)
SEGMENT PLAN ASPECT TAPER L.E. T.E. M.A.C. T/C
NUMBER AREA RATIO RATIO SWEEP SWEEP CHORD RATIO
1 77.0000 .325 .540 30.000 -51.621 15.858 .066
TOTAL 77.0000 .325 .540 30.000 -51.621 15.858 .066
1 ***** THE USAF AUTOMATED MISSILE DATCOM * REV 5/97 ***** CASE 1
AERODYNAMIC METHODS FOR MISSILE CONFIGURATIONS PAGE 5
HYPMIS
FIN SET 1SECTION AERODYNAMICS
IDEAL ANGLE OF ATTACK = .0000 DEG.
ZERO LIFT ANGLE OF ATTACK = .0000 DEG.
IDEAL LIFT COEFFICIENT = .0000
ZERO LIFT PITCHING MOMENT COEFFICIENT = .0000
MACH ZERO LIFT-CURVE-SLOPE = .0913 /DEG.
LEADING EDGE RADIUS = .0000 FRACTION CHORD
MAXIMUM AIRFOIL THICKNESS = .0660 FRACTION CHORD
DELTA-Y = .7271 PERCENT CHORD
*** CREST CRITICAL MACH NUMBER EXCEEDED ***
CREST CRITICAL MACH = .8250
LOCATION = .4337 FRACTION CHORD
LIFT-CURVE-SLOPE = .1301 /DEG.
1 ***** THE USAF AUTOMATED MISSILE DATCOM * REV 5/97 ***** CASE 1
AERODYNAMIC METHODS FOR MISSILE CONFIGURATIONS PAGE 6
HYPMIS
BODY ALONE PARTIAL OUTPUT
******* FLIGHT CONDITIONS AND REFERENCE QUANTITIES *******
MACH NO = 1.00 REYNOLDS NO = 7.068E+06 /FT
ALTITUDE = .0 FT DYNAMIC PRESSURE = 1481.36 LB/FT**2
SIDESLIP = .00 DEG ROLL = .00 DEG
REF AREA = 346.400 IN**2 MOMENT CENTER = 81.000 IN
REF LENGTH = 21.00 IN LAT REF LENGTH = 21.00 IN
ALPHA CA-FRIC CA-PRES/WAVE CA-BASE CA-PROT CA-SEP CA-ALP
.00 .0646 .4734 .0596 .0596
10.00 .0627 .4591 .0599 .0667
CROSS FLOW DRAG PROPORTIONALITY FACTOR = .83919
ALPHA CN-POTEN CN-VISC CN-SEP CM-POTEN CM-VISC CM-SEP CDC
.00 .000 .000 .000 .000 .740
10.00 .370 .210 .693 -.139 .846
FIN PRESSURE DIST. CANNOT BE CALCULATED DATCOM METHOD USED FOR WAVE DRAG
1 ***** THE USAF AUTOMATED MISSILE DATCOM * REV 5/97 ***** CASE 1
AERODYNAMIC METHODS FOR MISSILE CONFIGURATIONS PAGE 7
HYPMIS
FIN SET 1 CA PARTIAL OUTPUT
******* FLIGHT CONDITIONS AND REFERENCE QUANTITIES *******
MACH NO = 1.00 REYNOLDS NO = 7.068E+06 /FT
ALTITUDE = .0 FT DYNAMIC PRESSURE = 1481.36 LB/FT**2
SIDESLIP = .00 DEG ROLL = .00 DEG
REF AREA = 346.400 IN**2 MOMENT CENTER = 81.000 IN
REF LENGTH = 21.00 IN LAT REF LENGTH = 21.00 IN
SINGLE FIN PANEL ZERO-LIFT AXIAL FORCE COMPONENTS
SKIN FRICTION .0014
SUBSONIC PRESSURE .0001
TRANSONIC WAVE .0018
SUPERSONIC WAVE .0000
LEADING EDGE .0000
TRAILING EDGE .0000
TOTAL CAO .0033
FIN AXIAL FORCE DUE TO ANGLE OF ATTACK
ALPHA CA DUE TO LIFT (SINGLE PANEL) CA-TOTAL (4 FINS)
.00 .0000 .0130
10.00 .0000 .0128
1 ***** THE USAF AUTOMATED MISSILE DATCOM * REV 5/97 ***** CASE 1
AERODYNAMIC METHODS FOR MISSILE CONFIGURATIONS PAGE 8
HYPMIS
FIN SET 1 CN, CM PARTIAL OUTPUT
******* FLIGHT CONDITIONS AND REFERENCE QUANTITIES *******
MACH NO = 1.00 REYNOLDS NO = 7.068E+06 /FT
ALTITUDE = .0 FT DYNAMIC PRESSURE = 1481.36 LB/FT**2
SIDESLIP = .00 DEG ROLL = .00 DEG
REF AREA = 346.400 IN**2 MOMENT CENTER = 81.000 IN
REF LENGTH = 21.00 IN LAT REF LENGTH = 21.00 IN
NORMAL FORCE SLOPE AT ALPHA ZERO, CNA = .00398/DEG (1 PANEL)
CENTER OF PRESSURE FOR LINEAR CN = 3.76531 (CALIBERS FROM C.G.)
CENTER OF PRESSURE FOR NON-LINEAR CN = 3.41768 (CALIBERS FROM C.G.)
ALPHA CN CN CN CM CM CM
LINEAR NON-LINEAR TOTAL LINEAR NON-LINEAR TOTAL
.00 .0000 .0000 .0000 .0000 .0000 .0000
10.00 .0792 .0332 .1123 .2980 .1134 .4114
1 ***** THE USAF AUTOMATED MISSILE DATCOM * REV 5/97 ***** CASE 1
AERODYNAMIC METHODS FOR MISSILE CONFIGURATIONS PAGE 9
HYPMIS
STATIC AERODYNAMIC COEFFICIENTS TRIMMED IN PITCH
******* FLIGHT CONDITIONS AND REFERENCE QUANTITIES *******
MACH NO = 1.00 REYNOLDS NO = 7.068E+06 /FT
ALTITUDE = .0 FT DYNAMIC PRESSURE = 1481.36 LB/FT**2
SIDESLIP = .00 DEG ROLL = .00 DEG
REF AREA = 346.400 IN**2 MOMENT CENTER = 81.000 IN
REF LENGTH = 21.00 IN LAT REF LENGTH = 21.00 IN
ALPHA DELTA CL CD CN CA
.00 .00 .000 .611 .000 .611
10.00 -18.80 .310 .739 .433 .674
PANELS FROM FIN SET 1 WERE DEFLECTED OVER THE RANGE -25.00 TO 20.00 DEG
PANEL 1 WAS VARIED
PANEL 2 WAS VARIED
PANEL 3 WAS VARIED
PANEL 4 WAS VARIED
*** END OF JOB ***
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