DATCOM_Citation_dcm——DATCOM使用方法介绍(英文版)

2021年3月1日发(作者：foreigner什么意思)

*

*

File :

*

*

Purpose : This is an input file for the Digital Datcom program.

*

*

Author : Bill Galbraith

*

Holy Cows, Inc.

*

billg (at)

*

*

Last update:

*

May 12, 2010 - Fixed vertical fin, removed ground and power effects.

*

************************

*

List of Command Card

************************

*

*

Note : Command cards MUST start in column 1. They may appear in

*

any order, with the exception of NEXT CASE.

*

* NAMELIST

The contents of each applicatble namelist is dumped for the

*

case in the input system of units. (Very useful to see

*

input parameter values).

* SA

VE

Preserve the input data for the case to be used in the following

*

cases. This would be useful if multiple or comparison cases

*

are built.

* DIM FT

Sets the system units of measure. Options are FT, IN, M, CM

* NEXT CASE Terminates the reading of input cards and begins execution of

*

the case. Case data are destroyed following execution of a case,

*

unless the SA

VE card is present.

* TRIM

Trim calculations will be performed for each subsonic Mach number

*

within the case. A vehicle may be trimmed by deflecting a control

*

device on the wing or horizontal tail or by deflecting an

*

all-moveable horizontal stabilizer.

* DAMP

Provides dynamic-derivative results in addition to the standard

*

static- derivative output

* CASEID

Provides a case identification that is printed as part of the

*

output header.

* DUMP ALL

Prints the contents of the named arrays in the foot-pound-second

*

system. See Appendix C for list of arrays and their contents.

* DUMP CASE will print all the arrays that are used during case

*

execution prior to the conventional output. (Not particularily

*

useful, as all data is in nasty arrays)

* DUMP INPT will print dump of all input data blocks used for the case.

1

*

(also not useful)

* DUMP IOM

will print all the output arrays for the case. (not useful)

* DUMP ALL

will print all program arrays, even if not used for the case.

*

(surprisingly, not useful)

* DERIV DEG Defines the output units of measure for the static and dynamic

*

stability derivatives, either RAD or DEG. The following parameters

*

are affected: CLa, Cma, Cyb, Cnb, Clb, CLq, Cmq, Clp, Cyp, Cnp,

*

Cnr, Clr, CLad, CMad. JSBSim XML output is also switched between

*

degrees and radians for alpha, beta, etc.

* PART

Provides auxiliary and partial outputs at each Mach number in the

*

case. These outputs are automatically provided for all cases at

*

transonic Mach numbers.

* BUILD

This command provides configuration build-up data. Conventional

*

static and dynamic stability data are output for a LOT of items.

* PLOT

Causes data generated by the program to be written to logical

*

unit 13, which can be retained for input to the Plot Module.

*

(Looks like it dumps the data arrays out in column format. Not

*

too useful).

DIM FT

DERIV DEG

DAMP

PART

* DUMP IOM

**********************

*

Flight Conditions *

**********************

*

WT

Vehicle Weight

*

LOOP

Program Looping Control

*

1 = vary altitude and mach together, default)

*

2 = vary Mach, at fixed altitude

*

3 = vary altitude, at fixed Mach

*

NMACH

Number of Mach numbers or velocities to be run, max of 20

*

Note: This parameter, along with NALT, may affect the

*

proper setting of the LOOP control parameter.

*

MACH

Array(20) Values of freestream Mach number

*

VINF

Array(20) Values of freestream speed (unit: l/t)

*

NALPHA

Number of angles of attack to be run, max of 20

*

ALSCHD

Array(20) Values of angles of attack, in ascending order

*

RNNUB

Array(20) Reynolds number per unit length

*

Freestream Reynolds numbers. Each array element must

2

*

correspond to the respective Mach number/freestream

*

speed input, use LOOP=1.0

*

NALT

Number of atmospheric conditions to be run, max of 20

*

input as either altitude or pressure and temperature

*

Note: This parameter, along with NMACH, may affect the

*

proper setting of the LOOP control parameter.

*

ALT

Array(20) Values of geometric altitude

*

Number of altitude and values. Note, Atmospheric conditions

*

are input either as altitude or pressure and temperature. (MAX 20)

*

PINF

Array(20) Values of freestream Static Pressure

*

TINF

Array(20) Values of freestream Temperature

*

HYPERS

=.true.

Hypersonic analysis at all Mach numbers > 1.4

*

STMACH

Upper limit of Mach numbers for subsonic analysis

*

(0.6

*

TSMACH

Lower limit of Mach number for Supersonic analysis

*

(1.01<=TSMACH<=1.4)

Default to 1.4

*

TR

Drag due to lift transition flag, for regression analysis

*

of wing-body configuration.

*

= 0.0 for no transition (default)

*

= 1.0 for transition strips or full scale flight

*

GAMMA

Flight path angle

$$FLTCON WT=7000.0, LOOP=2.0,

NMACH=1.0, MACH(1)=0.4,

NALT=1.0, ALT(1)=0.0,

NALPHA=20.0,

ALSCHD(1)= -16.0, -8.0, -6.0, -4.0, -2.0, 0.0, 2.0, 4.0, 8.0, 9.0,

10.0, 12.0, 14.0, 16.0, 18.0, 19.0, 20.0, 21.0, 22.0, 24.0,

STMACH=0.6, TSMACH=1.4, TR=1.0$$

*************************

*

Reference Parameters *

pg 29

*************************

*

SREF

Reference area value of theoretical wing area used by program

*

if not input

*

CBARR

Longitudinal reference length value of theoritcal wing

*

Mean Aerodynamic Chord used by program if not input

*

BLREF

Lateral reference length value of wing span used by program

*

ROUGFC

Surface roughness factor, equivalent sand roughness, default

*

to 0.16e-3 inches (Natural sheet metal)

*

0.02/0.08E-3 - Polished metal or wood

*

0.16E-3

- Natural sheet metal

3

*

0.25E-3

- Smooth matte paint, carefully applied

*

0.40E-3

- Standard camouflage paint, average application

$$OPTINS SREF=320.8, CBARR=6.75, BLREF=51.7, ROUGFC=0.25E-3$$

**************************************

* Group II

Synthesis Parameters

*

pg 33

**************************************

*

*

page 33

*

*

XCG

Longitudinal location of cg (moment ref. center)

*

ZCG

Vertical location of CG relative to reference plane

*

XW

Longitudinal location of theoretical wing apex (where

*

leading edge would intersect long axis)

*

ZW

Vertical location of theoretical wing apex relative to

*

reference plane

*

ALIW

Wing root chord incident angle measured from reference plane

*

XH

Longitudinal location of theoretical horizontal tail apex.

*

If HINAX is input, XH and ZH are evaluated at zero incidence.

*

ZH

Vertical location of theoretical horizontal tail apex

*

relative to reference plane. If HINAX is input, XH and ZH

*

are evaluated at zero incidence.

*

ALIH

Horizontal tail root chord incidence angle measured from

*

reference plane

*

XV

Longitudinal location of theoretical vertical tail apex

*

XVF

Longitudinal location of theoretical ventral fin apex

*

ZV

Vertical location of theoretical vertical tail apex

*

This kinda makes sense only for twin tails that are canted

*

ZVF

Vertical location of theoretical ventral fin apex

*

This kinda makes sense only for twin tails that are canted

*

SCALE

Vehicle scale factor (multiplier to input dimensions)

*

VERTUP

Vertical panel above reference plane (default=true)

*

HINAX

Longitudinal location of horizontal tail hinge axis.

*

Required only for all-moveable horizontal tail trim option.

$$SYNTHS XCG=21.9, ZCG=3.125,

XW=19.1,

ZW=3.125,

ALIW=2.5,

XH=39.2,

ZH=7.75,

ALIH=0.0,

XV=36.0,

ZV=6.0,

XVF=28.0, ZVF=7.4,

SCALE=1.0, VERTUP=.TRUE.$$

4

**********************************

*

Body Configuration Parameters *

pg 36

**********************************

*

Here is an error message output by DIGDAT concerning body geometry:

*

IN

NAMELIST

BODY,

ONL

Y

THE

FOLLOWING

COMBINATIONS

OF

V

ARIABLES

CAN BE USED

*

FOR A CIRCULAR BODY, SPECIFY X AND R OR X AND S

*

FOR AN ELLIPTICAL BODY

, SPECIFY X AND R OR X AND S, AND THE V

ARIABLE

ELLIP

*

FOR OTHER BODY SHAPES X, R, S, AND P MUST ALL BE SPECIFIED

*

*

NX

Number of longitudinal body stations at which data is

*

specified, max of 20

*

X

Array(20) Longitudinal distance measured from arbitray location

*

S

Array(20) Cross sectional area at station. See note above.

*

P

Array(20) Periphery at station Xi. See note above.

*

R

Array(20) Planform half width at station Xi. See note above.

*

ZU

Array(20) Z-coordinate at upper body surface at station Xi

*

(positive when above centerline)

*

[Only required for subsonic asymmetric bodies]

*

ZL

Array(20) Z-coordinate at lower body surface at station Xi

*

(negative when below centerline)

*

[Only required for subsonic asymmetric bodies]

*

BNOSE

Nosecone type

1.0 = conical (rounded), 2.0 = ogive (sharp point)

*

[Not required in subsonic speed regime]

*

BTAIL

Tailcone type

1.0 = conical, 2.0 = ogive, omit for lbt = 0

*

[Not required in subsonic speed regime]

*

BLN

Length of body nose

*

Not required in subsonic speed regime

*

BLA

Length of cylindrical afterbody segment, =0.0 for nose alone

*

or nose-tail configuration

*

Not required in subsonic speed regime

*

DS

Nose bluntness diameter, zero for sharp nosebodies

*

[Hypersonic speed regime only]

*

ITYPE

1.0 = straight wing, no area rule

*

2.0 = swept wing, no area rule (default)

*

3.0 = swept wing, area rule

*

METHOD

1.0 = Use existing methods (default)

*

2.0 = Use Jorgensen method

$$BODY NX=8.0,

X(1)=0.0,1.0,2.7,6.0,8.8,28 .5,39.4,44.8,

5

R(1)=0.0,1.25,2 .1,2.7,2.76,2.7,1.25,0.0,

ZU(1)=3.5,4.3,4.8,5.5,7.4,7.4,6.5,5.7,

ZL(1)=3.5,2.5,2.25,2.1,2.0,2.2,4.3,5.7,

BNOSE=1.0, BLN=8.8,

BTAIL=1.0, BLA=19.7,

ITYPE=1.0, METHOD=1.0$$

**********************************

*

Wing planform variables

pg 37-38

**********************************

*

CHRDR

Chord root

*

CHRDBP

Chord at breakpoint. Not required for straight

*

tapered planform.

*

CHRDTP

Tip chord

*

SSPN

Semi-span theoretical panel from theoretical root chord

*

SSPNE

Semi-span exposed panel, See diagram on pg 37.

*

SSPNOP

Semi-span outboard panel. Not required for straight

*

tapered planform.

*

SA

VSI

Inboard panel sweep angle

*

SA

VSO

Outboard panel sweep angle

*

CHSTAT

Reference chord station for inboard and outboard panel

*

sweep angles, fraction of chord

*

TWISTA

Twist angle, negative leading edge rotated down (from

*

exposed root to tip)

*

SSPNDD

Semi-span of outboard panel with dihedral

*

DHDADI

Dihedral angle of inboard panel

*

DHDADO

Dihedral angle of outboard panel. If DHDADI=DHDADO only

*

input DHDADI

*

TYPE

1.0 - Straight tapered planform

*

2.0 - Double delta planform (aspect ratio <= 3)

*

3.0 - Cranked planform (aspect ratio > 3)

$$WGPLNF CHRDR=9.4,

CHRDTP=3.01,

SSPN=25.85,

SSPNE=23.46,

SAVSI=1.3,

CHSTA

T=0.25, TWISTA=-3.0,

DHDADI=3.6,

TYPE=1.0$$

< p>
******************************************** **

*

Wing Sectional Characteristics Parameters *

pg 39-40

6

*************************** *******************

*

The section aerodynamic characteristics for these surfaces are

*

input using either the sectional characteristics namelists WGSCHR,

*

HTSCHR, VTSCHR and VFSCHR and/or the NACA control cards. Airfoil

*

characteristics are assummed constant for each panel of the planform.

*

*

To avoid having to input all the airfoil sectional characteristics,

*

you can specify the NACA airfoil designation. Starts in Column 1.

*

* NACA x y zzzzzz

*

*

where:

*

column 1-4

NACA

*

5

any deliminator

*

6

W, H, V

, or F

Planform for which the airfoil

*

designation applies:

Wing, Horizontal

*

tail, Vertical tail, or Ventral fin.

*

7

any deliminator

*

8

1,4,5,6,S

Type of airfoil section: 1-series,

*

4-digit, 5-digit, 6-series, or Supersonic

*

9

any deliminator

*

10-80

Designation, columns are free format, blanks are ignored

*

*

TOVC

Maximum airfoil section thickness fraction of chord

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

DELTAY

Difference between airfoil ordinates at 6% and 15% chord,

*

percent chord (% correct ???)

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

XOVC

Chord location of maximum airfoil thickness, fraction of chord

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

CLI

Airfoil section design lift coefficient

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

ALPHAI

Angle of attack at section design lift coefficient, deg

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

7

*

CLALPA

Airfoil section lift curve slope dCl/dAlpha, per deg (array 20)

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

CLMAX

Airfoil section maximum lift cofficient (array 20)

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

CMO

Section zero lift pitching moment coefficient

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

LERI

Airfoil leading edge radius, fraction of chord

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

LERO

RLE for outboard panel, fraction of chord

*

[Required input].

*

Not required for straight tapered planforms.

*

CAMBER

Cambered airfoil flag flag

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

TOVCO

t/c for outboard panel

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

Not required for straight tapered planforms.

*

XOVCO

(x/c)max for outboard panel

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

Not required for straight tapered planforms.

*

CMOT

Cmo for outboard panel

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

Not required for straight tapered planforms.

*

CLMAXL

Airfoil maximum lift coefficient at mach = 0.0

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

CLAMO

Airfoil section lift curve slope at Mach=0.0, per deg

*

[Not required for subsonic speed regime. Required input

8

*

for transonic speed regime, user supplied or computed if

*

NACA card supplied]

*

TCEFF

Planform effective thickness ratio, fraction of chord

*

[Not required for subsonic speed regime. Required input

*

for transonic speed regime, user supplied or computed if

*

NACA card supplied]

*

KSHARP

Wave-drag factor for sharp- nosed airfoil section, not

*

input for round-nosed airfoils

*

[Not required for subsonic speed regime. Required input

*

for transonic speed regime, user supplied or computed if

*

NACA card supplied]

*

SLOPE

Airfoil surface slope at 0,20,40,60,80 and 100% chord, deg.

*

Positive when the tangent intersects the chord plane forward

*

of the reference chord point

*

[Not required for subsonic speed regime. Required input

*

for transonic speed regime, user supplied or computed if

*

NACA card supplied]

*

ARCL

Aspect ratio classification (see table 9, pg 41)

*

[Optional input]

*

XAC

Section Aerodynamic Center, fraction of chord

*

[Optional input, computed by airfoil section module if airfoil

*

defined with NACA card or section coordinates]

*

DW

ASH

Subsonic downwash method flag

*

= 1.0

use DATCOM method 1

*

= 2.0

use DATCOM method 2

*

= 3.0

use DATCOM method 3

*

Supersonic, use DATCOM method 2

*

[Optional input]

*

See figure 9 on page 41.

*

YCM

Airfoil maximum camber, fraction of chord

*

[Required input, user supplied or computed by airfoil

*

section module if airfoil defined with NACA card or

*

section coordinates]

*

CLD

Conical camber design lift coefficient for M=1.0 design

*

see NACA RM A55G19 (default to 0.0)

*

[Required input]

*

TYPEIN

Type of airfoil section coordinates input for airfoil

*

section module

*

= 1.0

upper and lower surface coordinates (YUPPER and YLOWER)

*

= 2.0 Mean line and thickness distribution (MEAN and THICK)

*

[Optional input]

*

NPTS

Number of section points input, max = 50.0

*

[Optional input]

*

XCORD

Abscissas of inputs points, TYPEIN=1.0 or 2.0, XCORD(1)=0.0

9