AFFDL-TR-79-3032
Volume I

THE USAF STABILITY AND CONTROL DATCOM
Volume I, Users Manual
McDonnell Douglas Astronautics Company
St. Louis Division
St Louis, Missouri 63166
April 1979
Updated by
Public Domain Aeronautical Software
Santa Cruz CA 95061
December 1999

AIR FORCE FLIGHT DYNAMICS LABORATORY
AIR FORCE WRIGHT AERONAUTICAL LABORATORIESAIR FORCE SYSTEMS COMMAND
WRIGHT-PATTERSON AIR FORCE BASE, OH 45433

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From AFFDL-TR-79-3032
THE USAF STABILITY AND CONTROL DIGITAL DATCOM
Volume I, Users Manual

SECTION 1
INTRODUCTION
In preliminary design operations, rapid and economical estimations of
aerodynamic stability and control characteristics are frequently required. The
extensive application of complex automated estimation procedures is often
prohibitive in terms of time and computer cost in such an environment. Similarinefficiencies accompany hand-calculation procedures, which can require
expenditures of significant man-hours, particularly if configuration trade studies
are involved, or if estimates are desired over a range of flight conditions. The
fundamental purpose of the USAF Stability and Control Datcom is to provide asystematic summary of methods for estimating stability and control characteristics
In preliminary design applications. Consistent with this philosophy, thedevelopment of the Digital Datcom computer program is an approach to provide
rapid and economical estimation of aerodynamic stability and control
characteristics.
Digital Datcom calculates static stability, high-lift and control device, anddynamic-derivative characteristics using the methods contained in Sections 4
through 7 of Datcom. The computer program also offer a trim option thatcomputes control deflections and aerodynamic data for vehicle trim at subsonic
Mach numbers.
The program has been developed an a modular basis as illustrated in Figure 1.
These modules correspond to the primary building blocks referenced in the
program executive. The modular approach was used because it simplified program
development, testing, and modification or expansion.
This report is the User's Manual for the USAF Stability and Control Digital
Datcom. Potential users are directed to Section 2 for an overview of program
capabilities. Section 3 provides input definitions, with basic configurationgeometry modeling techniques presented in Section 4. Analyses of special
configurations are treated in Section 5. Section 6 discusses the available output
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data. The appendices discuss namelist coding rules, airfoil section characteristic
estimation methods with supplemental data, and a list of geometric andaerodynamic variables available as supplemental output. A self-contained user's
kit is included to aid the user in setting up inputs to the program.
Even though the development of Digital Datcom was pursued with the sole
objective of translating the Datcom methods into an efficient, user-oriented
computer program, differences between Datcom and Digital Datcom do exist.
Such is the primary subject of Volume II, Implementation of Datcom Methods,
which contains the correspondence between Datcom methods and program
formulation. This volume also defines the program implementation requirements.
The listing of the computer program is contained on microfiche as a supplement tothis report. Modifications, extensions, and limitations of Datcom methods asincorporated in Digital Datcom are discussed throughout the report.
Users should refer to Datcom for the limitations of methods involved. However,potential users are forewarned that Datcom drag methods are not recommended for
performance. Where more than one Datcom method exists, Volume II indicates
which method or methods are employed in Digita1 Datcom.
Direct all program inquiries to AFFDL FGC, Wright-Patterson Air Force Base,OH 45433; phone (513) 255-4315.

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MASTER ROUTINES
Main Programs

Performs the executive functions of organizing and directing theoperations performed by other program components.

Executive Subroutines

Performs user-oriented non-method operations such as ordering input
data, logic switching, input error analysis, and output format selection.

Utility Subroutines

Performs standard mathematical tasks repetitively performed by method
subroutines.

METHOD MODULESSUBSONIC

TRANSONIC

SUPERSONICSPECIAL
CONFIGURATIONSMODULE 1
MODULE 3
MODULE 5
MODULE 7
CHARACTERISTICS AT CHARACTERISTICS AT CHARACTERISTICS AT LOW ASPECT RATIOANGLE OF ATTACKANGLE OF ATTACKANGLE OF ATTACK
WING-BODY AT
SUBSONIC SPEEDSMODULE 2
CHARACTERISTICS IN
SIDESLIP

MODULE 4
CHARACTERISTICS IN
SIDESLIP

MODULE 6
CHARACTERISTICS IN
SIDESLIP

MODULE 10
DYNAMIC DERIVATIVESMODULE 8
AERODYNAMIC
CONTROL
EFFECTIVENESS AT
HYPERSONIC SPEEDSMODULE 9
TRAVERSE-JET
CONTROL
EFFECTIVENESS AT
HYPERSONIC SPEEDSMODULE 11
HIGH LIFT AND CONTROL DEVICESMODULE 7
TRIM OPTIONFIGURE 1 - DIGITAL DATCOM MODULES

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SECTION 2PROGRAM CAPABILITIESThis section has been prepared to assist the potential user in his decision processconcerning the applicability of the USAF Stability and Control Digital Datcom tohis particular requirements. For specific questions dealing with method validity
and limitations, the user is strongly encouraged to refer to the USAF Stability andControl Datcom document. Much of the flexibility inherent in the Datcom
methods has been retained by allowing the user to substitute experimental or
refined analytical data at intermediate computation levels. Extrapolations beyond
the normal range of the Datcom methods are provided by the program; however,each time an extrapolation is employed, a message is printed which identifies thepoint at which the extrapolation is made and the results of the extrapolation.
Supplemental output is available via the “dump” and “partial output” options
which give the user access to key intermediate parameters to aid verification or
adjustment of computations. The following paragraphs discuss primary program
capabilities as well as selected qualifiers and limitations.
2.1 ADDRESSABLE CONFIGURATIONS
In general, Datcom treats the traditional body-wing-tail geometries including
control effectiveness for a variety of high-lift /control devices. High-lift/control
output is generally in terms of the incremental effects due to deflection. The user
must integrate these incremental effects with the “basic” configuration output.
Certain Datcom methods applicable to reentry type vehicles are also available.
Therefore, the Digital Datcom addressable geometries include the “basic”
traditional aircraft concepts (including canard configurations), and uniquegeometries which are identified as “special” configurations. Table 1 summarizes
the addressable configurations accommodated by the program.

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CONFIGURATIONPROGRAM REMARKSBODY

PRIMARILY BODIES OF REVOLUTION, OR CLOSE APPROXIMATIONS,ARE TREATED. TRANSONIC METHODS FOR MOST OF THE
AERODYNAMIC DATA DO NOT EXIST. THE RECOMMENDED
PROCEDURE REQUIRES FAIRING BETWEEN SUBSONIC AND
SUPERSONIC DATA USING AVAILABLE DATA AS A GUIDE.

WING,
HORIZONTAL TAILSTRAIGHT TAPERED, CRANKED, OR DOUBLE DELTA PLANFORMS ARE
TREATED. EFFECTS OF SWEEP, TAPER AND INCIDENCE ARE
INCLUDED. LINEAR TWIST IS TREATED AT SUBSONIC MACH
NUMBERS. DIHEDRAL EFFECTS ARE PRESENT IN THE LATERALDIRECTIONAL DATA.

BODY-WING BODYHORIZONTALLONGITUDINAL METHODS REFLECT ONLY A MID-WING POSITION.
LATERAL DIRECTIONAL SOLUTIONS CONSIDER HIGH AND LOW-WING
POSITIONS.

WING-BODY-TAIL

THE VARIOUS GEOMETRY COMBINATIONS ARE GIVEN IN TABLE 2.
WING DOWNWASH METHODS ARE RESTRICTED TO STRAIGHTTAPERED PLANFORMS. EFFECTS OF TWIN VERTICAL TAILS ARE
INCLUDED IN THE STATIC LATERAL DIRECTIONAL DATA AT SUBSONIC
MACH NUMBERS.

NON-STANDARD
GEOMETRIESNON-STANDARD CONFIGURATIONS ARE SIMULATED USING “BASIC”
CONFIGURATION TECHNIQUES. STRAKES CAN BE RUN VIA A DOUBLEDELTA WING. A BODY-CANARD-WING IS INPUT AS A WING-BODYHORIZONTAL TAIL. THE FORWARD LIFTING SURFACE IS INPUT AS A
WING AND THE AFT SURFACE AS A HORIZONTAL TAIL.

SPECIAL
CONFIGURATIONSLOW ASPECT RATIO WING OR WING-BODY CONFIGURATIONS(LIFTING BODIES) ARE TREATED AT SUBSONIC SPEEDS. TWODIMENSIONAL FLAP AND TRANSVERSE JET EFFECTS ARE ALSOTREATED AT HYPERSONIC SPEEDS.

TABLE 1 - ADDRESSABLE CONFIGURATIONSPage 6 of67

2.2 BASIC CONFIGURATION DATA
The capabilities discussed below apply to basic configurations, i.e., traditional
body-wing-tail concepts. A detailed summary of output as a function ofconfiguration and speed regime is presented in Table 2. Note that transonic output
can be expanded through the use of data substitution (Sections 3.2 and 4.5).
Typical output for these configurations are presented in section 6.
2.2.1 Static Stability Characteristics
The longitudinal and lateral-directional stability characteristics provided by the
Datcom and the Digital Datcom are in the stability-axis system. Body-axis normalforce and axial-force coefficients are also included in the output for convenienceof the user. For those speed regimes and configurations where Datcom methods
are available, the Digital Datcom output provides the longitudinal coefficients CD,
CL, Cm, CN, and CA, and the derivatives
C L α , C m α , C Y β , C n β , C lβ

Output for configurations with a wing and horizontal tail also includes downwash
and the local dynamic-pressure ratio in the region of the tail. Subsonic data that
include propeller power, jet power, or ground effects are also available. Power andground effects are limited to the longitudinal aerodynamic characteristics.
Users are cautioned that the Datcom does not rigorously treat aerodynamics in thetransonic speed regime, and a fairing between subsonic and supersonic solutions isoften the recommended procedure. Digital Datcom uses linear and nonlinear
fairings through specific points; however, the user may find another fairing more
acceptable. The details of these fairing techniques are discussed in Volume II,
Section 4. The partial output option, discussed in Section 3.5, permits the user toobtain the information necessary for transonic fairings. The experimental data
input option allows the user to revise the transonic fairings on configuration
components, perform parametric analyses on test configurations, and apply bettermethod results (or data) for configuration build-up.
Datcom body aerodynamic characteristics can be obtained at all Mach numbers
only for bodies of revolution. Digital Datcom can also provide subsoniclongitudinal data for cambered bodies of arbitrary cross section as shown in Figure
6. The cambered body capability is restricted to subsonic longitudinal-stability
solutions.
Straight-tapered and nonstraight-tapered wings including effects of sweep, taper,and incidence can be treated by the program. The effect of linear twist can betreated at subsonic Mach numbers. Dihedral influences are included in lateralPage 8 of67

directional stability derivatives and wing
wake location used in the calculation of longitudinal data. Airfoil section
characteristics or a required input, although most of these characteristics may begenerated using the Airfoil Section Module (Appendix B). Users are advised to be
mindful of section characteristics which are sensitive to Reynolds number,particularly in cases where very low Reynolds number estimates are of interest. A
typical example would be pretest estimates for small, laminar flow wind tunnels
where Reynolds numbers on the order of 100,000 are common.
Users should be aware that the Datcom and Digital Datcom employ turbulent
skin friction methods in the computation of friction drag values. Estimates forcases involving significant wetted areas in laminar flow will require adjustment by
the user.
Computations of wing-body longitudinal characteristics assume, in manycases, that the configuration is of the mid-wing type. Lateral-directional analyses
do account for other wing locations. Users should consult the Datcom for specificdetails.
Wing-body-tail configurations which may be addressed are shown in Table 2.
These capabilities permit the user to analyze complete configurations, including
canard and conventional aircraft arrangements. Component aerodynamic
contributions and configuration build-up data are available through the use of the“BUILD” option described In Section 3.5. Using this option, the user can isolatecomponent aerodynamic contributions in a similar fashion to break down datafrom a wind tunnel where. such information in of value in obtaining an overall
understanding of a specific configuration.
Twin vertical panels can be placed either on the wing or horizontal tail.
Analysis can be performed with both twin vertical tail panels and a conventionalvertical tail specified though interference effects between the three panels is notcomputed. The influence of twin vertical tails is included only in the lateraldirectional stability characteristics at subsonic speeds.
2.2.2 Dynamic Stability Characteristics
The pitch, acceleration, roll and yaw derivatives of
C L q , C m q , C L α• , C m α• , C l p , C Y p , C n p , C n r , an d C l r
are computed for each component and the build-up configurations shown in Table
2. All limitations discussed in Section 7 of the USAF Stability and Control
Datcom are applicable to digital Datcom as well. The experimental data option of
the program (Section 4.5) permits the user to substitute experimental data for key
parameters involved in dynamic derivative solutions, such as body dCL/d andwing-body dCL/d. Any improvement in the accuracy of these key parameters will
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produce significant improvement in the dynamic stability estimates. Use ofexperimental data substitution for this purpose is strongly recommended.
2.2.3 High-Lift and Control Characteristics
High-lift devices that can be analyzed by the Datcom methods include jet flaps,
split, plain, single-slotted, double-slotted, fowler, and leading edge flaps and slats.
Control devices, such as trailing-edge flap-type controls and spoilers, can also betreated. In general terms, the program provides the incremental effects of high lift
or control device deflections at zero angle of attack.
The majority of the high-lift-device methods deal with subsonic lift, drag, andpitching-moment effects with flap deflection. General capabilities for jet flaps,
symmetrically deflected high-lift devices, or trailing-edge control devices include
lift, moment, and maximum-lift increments along with drag-polar increments andhinge-moment derivatives. For translating devices the lift-curve slope is also
computed. Asymmetrical deflection of wing control devices can be analyzed forrolling and yawing effectiveness. Rolling effectiveness may be obtained for allmovable differentially-deflected horizontal stabilizers. The speed regimes wherethese capabilities exist are shown in Table 3.
Control modes employing all-moving wing or tail surfaces can also be addressedwith the program. This is accomplished by executing multiple cases with a varietyof panel incidence angles.
2.2.4 Trim Option
Trim data can be calculated at subsonic speeds. Digital Datcom manipulates
computed stability and control characteristics to provide trim output (static Cm=
0.0). The trim option is available in two modes. One mode treats configurations
with a trim control device on the wing or horizontal tail. Output is presented as afunction of angle of attack and consists of control deflection angles required totrim and the associated longitudinal aerodynamic characteristics shown in Table 3.
The second mode treats conventional wing-body-tail configurations where thehorizontal-tail is all-movable or “flying.” In this case, output as a function of angleof attack consists of horizontal-stabilizer deflection (or incidence) angle required
to trim; untrimmed stabilizer CL, CD, Cm, and hinge-moment coefficients; trimmed
stabilizer CL, CD, and hinge moment coefficients; and total wing-body-tail CL and
CD. Body-canard-tail configurations may be trimmed by calculating the stability
characteristics at a variety of canard incidence angles and manually calculating thetrim data. Treatment of a canard configuration is addressed in Table 1.

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2.3 SPECIAL CONFIGURATION DATA
The capabilities discussed below apply to the three special configurations
illustrated in Figure 2.
2.3.1 Low-Aspect-Ratio Wings and Wing-Body Combinations
Datcom provides methods which apply to 1ifting reentry vehicles at subsonic
speeds. Digital Datcom output provides longitudinal coefficients CD, CL, Cm, CN,
and CA and the derivatives
C L α , C m α , C Y β , C n β , C lβ
2.3.2 Aerodynamic Control at Hypersonic Speeds
The USAF Stability and Control Datcom contains some special control methods
for high-speed vehicles. These include hypersonic flap methods which areincorporated into Digital Datcom. The flap methods are restricted to Mach
numbers greater than 5, angles of attack between zero and 20 degrees anddeflections into the wind. A two-dimensional flow field is determined and oblique
shock relations are used to describe the flow field.
Data output from the hypersonic control-flap methods are incremental normal- andaxial-force coefficients, associated hinge moments, and center-of-pressure
location. These data are found from the local pressure distributions on the flap andin regions forward of the flap. The analysis includes the effects of flow separation
due to windward flap deflection by providing estimates for separation inducedpressures forward of the flap and reattachment on the flap, Users may specify
laminar or turbulent boundary layers.
2.3.3 Transverse-jet Control Effectiveness
Datcom provides a procedure for preliminary sizing of a two-dimensionaltransverse-jet control system in hypersonic flow, assuming that the nozzle islocated at the aft end of the surface. The method evaluates the interaction of thetransverse jet with the local flow field. A favorable interaction will produce
amplification forces that increase control effectiveness.
The Datcom method is restricted to control jets located on windward surfaces in a
Mach number range of 2 to 20. In addition, the method is invalid for altitudes
where mean free paths approach the jet-width dimension.
The transverse control jet method requires a user-specified time history of local
flow parameters and control force required to trim or maneuver. With these data,
the minimum jet plenum pressure is then employed to calculate the nozzle throat
diameter and the jet plenum pressure and propellant weight requirements to trim
or maneuver the vehicle.
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Page 13 of67

2.4 OPERATIONAL CONSIDERATIONSThere are several operational considerations the user needs to understand in orderto take maximum advantage of Digital Datcom.
2.4.1 Flight Condition Control
Digital Datcom requires Mach number and Reynolds number to define the flight
conditions. This requirement can be satisfied by defining combinations of Mach
number, velocity, Reynolds number, altitude, and pressure and temperature. The
input options for speed reference and atmospheric conditions that satisfy therequirement are given in Figure 3, The speed reference is input as either Mach
number or velocity, and the atmospheric conditions as either altitude or freestreampressure and temperature. The specific reference and atmospheric conditions arethen used to calculate Reynolds number.
The program may loop on speed reference and atmospheric conditions threedifferent ways, an given by the variable LOOP in Figure 3. In this discussion, andin Figure 3, the speed reference is referred to as Mach number, and atmosphericconditions as altitude. The three options for program looping on Mach number andaltitude are listed and discussed below.
•••

LOOP = 1 - Vary Mach and altitude together. The program executes at thefirst Mach number and first altitude, the second Mach number and second
altitude, and continues for all the flight conditions. In the input data,
NMACH must equal NALT and NMACH flight conditions are executed.
This option should be selected when the Reynolds number is input, andmust be selected when atmospheric conditions are not input.
LOOP = 2 - Vary Mach number at fixed altitude. The program executes
using the first altitude and cycles through each Mach number in the input
list, the second altitude and cycles through each Mach number, andcontinues until each altitude has been selected. Atmospheric conditions oustbe input for this option and NMACH times NALT flight conditions are
executed.
LOOP = 3 - Vary altitude at fixed Mach number. The program executes
using the first Mach number and cycles through each altitude in the input
list, the second Mach number and cycles through each altitude. andcontinues until each Mach number has been selected. Atmosphericconditions must be input for this option and NMACH times NALT flight
conditions are executed.

2.4.2 Mach Regimes
Aerodynamic stability methods are defined in Datcom as a function of vehicle
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configuration and Mach regime. Digital Datcom logic determines theconfiguration being analyzed by identifying the particular input namelists that arepresent within a case (see Section 3). The Mach regime is normally determinedaccording to the following criteria:
Mach Number (M)Mach Regime
M < 0.6
Subsonic
0.6 < M < 1.4
Transonic
M > 1.4
SupersonicM > 1.4
Hypersonicand the hypersonicflag is set (see Figure 3)
These limits were selected to conform with most Datcom methods. However, somemethods are valid for a larger Mach number range. Some subsonic methods arevalid up to a Mach number of 0.7 or 0.8. The user has the option to increase thesubsonic Mach number limit using the variable STMACH described In Section
3.2. The program will permit this variable to be in the range: 0.6 < STMACH <0.99. In the same fashion, the supersonic Mach limit can be reduced using thevariable TSMACH. The program will permit this value be in the range: 1.01 <TSMACH < 1.40. The program will default to the limits of each variable if therange is exceeded. The Mach regimes are then defined as follows:
Mach Number (M)M < STMACH
STMACH < M < TSMACH
M > TSMACH
M > TSMACH
and the hypersonicflag is set

Mach Regime
Subsonic
Transonic
SupersonicHypersonic

2.4.3 Input Diagnostics
There to an input diagnostic analysis module in Digital Datcom which scans all of
the input data cards prior to program execution. A listing of all input data is given
and any errors are flagged. It checks all namelist cards for correct namelist nameand variable name spelling, checks the numerical inputs for syntax errors, andchecks for legal control cards. The namelist and control cards are described in
Section 3.
This module does not “fix up” input errors. It will, however, insert a namelist
termination if it is not found. Digital Datcom will attempt to execute all cases as
input by the user even if errors are detected.
2.4.4 Airfoil Section Module
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The airfoil section module can be used to calculate the required geometric and.
aerodynamic input parameters for virtually any user defined airfoil section. This
module substantially simplifies the user's input preparation.
An airfoil section is defined by one of the following methods:
1.
An airfoil section designation (for NACA, double wedge, circular arc, or
hexagonal airfoils)
2.
Section upper and lower cartesian coordinates, or3.
Section mean line and thickness distribution.
The airfoil section module uses Weber's method (References 2 to 4) to calculate
the inviscid aerodynamic characteristics. A viscous correction is applied to thesection lift curve slope, cg . In addition a 5% correlation factor (suggested in
Datcom, page 4.1.1.2-2) is applied to bring the results in line with experimental
data. The airfoil section module methods are discussed in Appendix A.
The airfoil section is assumed to be parallel to the free stream. Skewed airfoils canbe handled by supplying the section coordinates parallel to the free stream. The
module will calculate the characteristics of any input airfoil, so the user must
determine whether the results are applicable to his particular situation. Five
general characteristics of the module should be noted.
1.
For subsonic Mach numbers, the module computes the airfoil subsonicsection characteristics and the results can be considered accurate for Mach
numbers less than the crest critical Mach number. Near crest critical Mach
number, flow mixing due to the upper surface shock will make the boundary
layer correction invalid. Compressibility corrections also become invalid.
The module also computes the required geometric variables at all speeds,
and for transonic and supersonic speeds these are the only required inputs.
Mach equals zero data are always supplied.
2.
Because of the nature of the solution, predictions for an airfoil whose
maximum camber is greater than 6% of the chord will lose accuracy.
Accuracy will also diminish when the maximum airfoil thickness exceeds
approximately 12% of the chord, or large viscous interactions are present
such as with supercritical airfoils.
3.
When section cartesian coordinates or mean line and thickness distribution
coordinates are specified, the user must adequately define the leading edge
region to prevent surface curve fits that have infinite slope. This can beaccomplished by supplying section ordinates at non-dimensional chordstations (x/c of 0.0, 0.001, 0.002, and 0.003.
4.
If the leading edge radius is not specified in the airfoil section input, theuser must insure that the first and second coordinate points lie on theleading edge radius. For sharp nosed airfoils the user must specify a zero
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5.

leading edge radius.
The computational algorithm can be sensitive to the “smoothness” of the
input coordinates. Therefore, the user should insure that the input datacontains no unintentional fluctuations. Considering that Datcom proceduresare preliminary design methods, it is at least as important to provide
smoothly varying coordinates. as it is to accurately define the airfoil
geometry.

2.4.5 Operational Limitations
Several operational limitations exist in Digital Datcom. These limitations are listed
below without extensive discussion or justification. Some pertinent operational
techniques are also listed.
•

•
•

•

•

•
•

•

•

The forward lifting surface is always input as the wing and the aft lifting
surface as the horizontal tail. This convention is used regardless of thenature o! the configuration.
Twin vertical tail methods are only applicable to lateral stability parametersat subsonic speeds.
Airfoil section characteristics are assumed to be constant across the airfoil
span, or an average for the panel. Inboard and outboard panels of cranked or
double-delta planforms can have their individual panel leading edge radiiand maximum thickness ratios specified separately.
If airfoil sections are simultaneously specified for the same aerodynamic
surface by an NACA designation and by coordinates, the coordinate
information will take precedence.
Jet and propeller power effects are only applied to the longitudinal stability
parameters at subsonic speeds. Jet and propeller power effects cannot beapplied simultaneously.
Ground effect methods are only applicable to longitudinal stability
parameters at subsonic speeds.
Only one high lift or control device can be analyzed at a time. The effect ofhigh lift and control devices on downwash is not calculated. The effects ofmultiple devices can be calculated by using the experimental data input
option to supply the effects of one device and allowing Digital Datcom to
calculate the incremental effects of the second device.
Jet flaps are considered to be symmetrical high lift and control devices. The
methods are only applicable to the longitudinal stability parameters at
subsonic speeds.
The program uses the input namelist names to define the configuration
components to be synthesized. For example, the presence of namelist
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HTPLNF causes Digital Datcom to assume that the configuration has ahorizontal tail.
Should Digital Datcom not provide output for those configurations for which
output is expected, as shown in Table 2, limitations on the use of a Datcom
method has probably been exceeded. In all cases users should consult the Datcom
for method limitations.

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67

SECTION 3
DEFINITION OF INPUTSThe Digital Datcom basic input data unit is the “case.” A “case” is a set of input
data that defines a configuration and its flight conditions. The case consists of
inputs from up to four data groups.
•
•

•

•

Group I inputs define the flight conditions and reference dimensions.
Group II inputs specify the basic configuration geometry for conventional
configurations, defining the body, wing and tail surfaces and their relative
locations.
Group III inputs specify additional configuration definition, such asengines, flaps, control tabs, ground effects or twin vertical panels. This
input group also defines those “special” configurations that cannot be
described using Group II inputs and include low aspect ratio wing andwing-body configurations, transverse jet control and hypersonic flaps.
Group IV inputs control the execution of the case, or job for multiple cases,
and allow the user to choose some of the special options, or to obtain extra
output.

3.1 INPUT TECHNIQUE
Two techniques are generally available for introducing input data into a Fortrancomputer program: namelist and fixed format. Digital Datcom employs thenamelist input technique for input Groups I, II and III since it is the mostconvenient and flexible for this application. Its use reduces the possibility of input
errors and increases the utility of the program as follows:•
Variables within a namelist may be input in any order;•
Namelist variables are not restricted to particular card columns;•
Only required input variables need be included; and•A variable may be included more than once within a namelist, but the lastvalue to appear will be used.
Namelist rules used In the program and applicable to CDC and IBM systems arepresented In Appendix A. The user should adhere to them when preparing inputs
for Digital Datcom. To aid the user in complying with the general namelist rules,
examples of both correct and incorrect namelist coding are included in Appendix
A.
All namelist input variables (and program data blocks) are initialized “UNUSED”
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(1.0E-60 on CDC systems) prior to case execution. Therefore, omission of
pertinent input variables may result in the “UNUSED” value to be used in
calculations. However, the “UNUSED” value is often used as a switch forprogram control, so the user should not indiscriminately use dummy inputs.
All Digital Datcom numeric constants require a decimal point. The Fortranvariable names that are implied INTEGERS (name begins with I, J, K, L, M, or N)
are declared REAL and must be specified in either “E” or “F” format(X.XXXEYY or X.XXX).
Group IV inputs are the “case control cards.” Though they are input in a fixed
format, their use has the characteristic of a namelist. since (with the exception of
the case termination card) they can be placed in any order or location in the input
data. Descriptions and limitations of each of the available control cards are
discussed in Section 3.5.
Table 4 defines the namelists and control cards that can be input to the program.
Since not all namelist inputs are required to define a particular problem or
configuration, those namelists required for various analyses are summarized inTables 5 through 7. Use of these tables will save time in preparing namelist inputs
for a specific problem.
The user has the option to specify the system of units to be used, English or
Metric. Table 8 summarizes the systems available, and defines the case control
card required to invoke each option. For clarity, the namelist variable description
charts which follow have a column titled “Units” using the following
nomenclature:ldenotes units of length: feet, inches, meters, or centimetersA
denotes units of area: ft2, in2, m2, or cm2
Deg denotes angular measure in degrees, or temperature in degrees
Rankine or degrees KelvinFdenotes units of force; pounds or Newtonstdenotes units of time; seconds.
Specific input parameters, geometric illustrations, and supporting data areprovided throughout the report. To aid the user in reading these figures, thecharacter ‘0' defines the number zero and the character ‘O’ the fifteenth letter in
the alphabet.

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GROUP I INPUT DATA

Namelist input data to define the case flight conditions and reference dimensions
are shown in Figures 3 and 4.
Namelist FLTCON, Figure 3, defines the case flight conditions. The user may opt
to provide Mach number and Reynolds number per unit length for each case to becomputed. In this case, input preparation requires that the user compute Reynolds
number for each Mach number, and altitude combination he desires to run.
However, the program has a standard atmosphere model, which accurately
simulates the 1962 Standard Atmosphere for geometric altitudes from -16,404 feetto 2,296,588 feet, that can be used to eliminate the Reynolds number input
requirement and provides the user the option to employ Mach number or velocity
as the flight speed reference. The user may specify Mach numbers (or velocities)
and altitudes for each case and program computations will employ the atmospheremodel to determine pressure, temperature, Reynolds number and other required
parameters to support method applications.
Also incorporated is the provision for optional inputs of pressure and temperatureby the user. The program will override the standard atmosphere and compute flow
condition parameters consistent with the pressure and temperature inputs. This
option will permit Digital Datcom applications such as wind tunnel model
analyses at test section conditions.
The five input combinations which will satisfy the Mach number and Reynolds
number requirements are summarized in Figure 3. If the NACA control card isused. the Reynolds number and Mach number must be defined using the variables
RNNUB and MACH.
Other optional inputs include vehicle weight and flight path angle (“WT” and“GAMMA”). These parameters are of particular interest when using the Trim
Option (Section 3.5). The trim flight conditions are output as an additional line of
output with the trim data and the steady flight lift coefficient is output with theuntrimmed data.
Use of the variable LOOP enables the user to run cases at fixed altitude withvarying Mach number (or velocity), at fixed Mach number (or velocity) at varying
altitudes, or varying speed and altitude together.
Nondimensional aerodynamic coefficients generated by Digital Datcom may bebased on user-specified reference area and lengths. These reference parameters
are input via namelist OPTINS, Figure 4. If the reference area is not specified, it isset equal to the theoretical planform area of the wing. This wing area includes thefuselage area subtended by the extension of the wing leading and trailing edges tothe body center line. The longitudinal reference length, if not specified in
OPTINS, in set equal to the theoretical wing mean aerodynamic chord. The lateral
Page 22 of67

reference length is set equal to the wing span when it is not user specified.
Reference parameters contained in OPTINS must be specified for body-alone
configurations since the default reference parameters are based on wing geometry.
It is suggested that values near the magnitude of body maximum cross-sectional
area be used for the reference area and body maximum diameter for thelongitudinal and lateral reference lengths.
The output format generally provides at least three significant digits in the solution
when user specified reference parameters are of the same order of magnitude as
the default reference parameters. If the user specifies reference parameters that areorders of magnitude different from the wing area or aerodynamic chord, some
output data can overflow the output format or print only zeros. This may happen
in rare instances and would require readjustment of the reference parameters.

Page 23 of

67

NOTE: The INERTA namelist is a new addition to the Digital Datcom program, addedby Holy Cows, Inc. in May 2011. As such, this namelist will not be recognized by otherversions of the Digital Datcom program. These parameters are strictly passed through theDigital Datcom program into the JSBSim output file generated by Datcom+ Pro, to allow
the JSBSim model files generated by Datcom+ Pro to be executed without additional
inputs.
Figure 4A INERTA inputs
Variable
Name
IXX
IYY
IZZ
IXY
IXZ
IYZ
NFUELTXFUELTYFUELTZFUELTFUELRD
FUELCP
FUELQT

Array
Dimension
10
10
10
10
10
10

ACTYPE-METHOD-

Definition
Ixx Moment of Inertia
Iyy Moment of Inertia
Izz Moment of Inertia
Ixy Cross Product of Inertia (usually insignificant)Ixz Cross Product of Inertia
Iyz Cross Product of Inertia (usually insignificant)Number of fuel tanks, maximum of 10
X location of fuel tanks, maximum of 10 values
Y location of fuel tanks, maximum of 10 values
Z location of fuel tanks, maximum of 10 values
Effective fuel tank radius, maximum of 10 values
Fuel tank maximum quantity, maximum of 10 values
Fuel tank current quantity, maximum of 10 values
Not required if IXX, IYY, and IZZ are known. This tells
the inertia estimation equations how to calculate theinertias.
Aircraft type:
0 - Glider
1 - Light Single
2 - Light Twin
3 - WW II Fighter ( or subsonic Racer/aerobatic)
4 - Single-Engine Transonic of Supersonic Fighter
5 - Two-Engine Transonic or Supersonic Fighter
6 - Two-Engine Transonic Transport
7 - Three-engine Transonic Transport
8 - Four+-engine Transonic Transport
9 - Multi-engine Prop Transport
If no aircraft type is entered and primary moments ofinertia are not present, the program will attempt to guess
which aircraft type to use.
Estimation method
1 - Roskam equations, also used in Aeromatic(default)

Page 26A of 67

Unitsmass · length^2
mass · length^2
mass · length^2
mass · length^2
mass · length^2
mass · length^2
lengthlengthlengthlengthweightweight--3.3 GROUP II INPUT DATA
Namelist data to define basic configuration geometry is shown in Figures 5
through 8. Those “special” configurations (Figure 2) are defined using Group III
namelists.
The namelist SYNTHS defines the basic configuration synthesis parameters. The
user has the option to apply a scale factor to his geometry which permits full scaleconfiguration dimensions to be input for an analysis of a wind tunnel model. The
program will use the scale factor to scale the input data to model dimensions. The
variable used is “SCALE.”
The body configuration is defined using the namelist BODY (Figure 6). The
variable METHOD enables the user to select either the traditional Datcom
methods for body CL, Cm and CD at low angles of attack (default), or Jorgensen’s
method, which is applicable from zero to 180 degrees angle of attack. Jorgensen’s
method can be used by selecting “METHOD=2” subsonically or supersonically.
Users are encouraged to consult the Datcom for details concerning these methods.
Digital Datcom will accept an arbitrary origin for the body coordinate system, i.e.,body station “zero” is or required to be at the fuselage nose.
The planform geometry of each of the aerodynamic surfaces are input using thenamelist WGPLNF, HTPLNF, VTPLNF and VFPLNF shown in Figure 7. The
section aerodynamic characteristics for these surfaces are input using either thesection characteristics namelists WGSCHR, HTSCHR, VTSCHR and VFSCHR(Figure 4) and/or the NACA control card discussed in Section 3.5. Airfoil
characteristics are assumed constant for each panel of the planform.
The USAF Datcom contains three methods for the computation of forward lifting
surface downwash field effects on aft lifting surface aerodynamics. They are
given in detail in Section 4.4 of Datcom, and their regimes of primary applicability
are summarized in Figure 9. The user is cautioned not to apply the empirically
based subsonic Method 2 outside the bounds listed in Figure 9. Method 1 isrecommended as an optional approach for the bw/bh regime of 1.0 to 1.5.
By default, Digital Datcom selects Method 3 for bw/bh less than 0.5 and Method 1
for span ratios greater than or equal to 1.5.
Using the variable DWASH In namelist WGSCHR, the user has the option of
applying Method 1 or 2. Method 2 is applicable at subsonic Mach numbers andspan ratios of 1.25 to 3.6.
Aspect ratio classification is required to employ the Datcom straight tapered wing
solutions for wing or tail lift in the subsonic and transonic Mach regimes.
Classification of lifting surface aspect ratio as either high or low results in theselection of appropriate methods for computation. The USAF Datcom uses aclassification parameter, which depends upon planform taper ratio and leading
Page 27 of67

edge sweep (Table 9). It also notes an overlap regime where the user may employ
either the low or high aspect ratio methods. Digital Datcom allows the user tospecify the aspect ratio method to be used in this overlap regime using theparameter ARCL in the section namelists. High aspect ratio methods areautomatically selected for unswept, untapered wings with aspect ratios of 3.5 ormore if ARCL is not input.
Transonically, several parameters need to be defined to obtain the panel liftcharacteristics. Those required variables are summarized In Figures 10 and 11 andare input using the experimental data substitution namelist EXPRnn. Additionally,
intermediate data may be available, for example (dCl/d)/CL, which requires
experimental data to complete. By use of the experimental data input namelist
EXPRnn, data can be made available to complete these second-levelcomputations. as shown in Figure 10.
The namelist EXPRnn can also be used to substitute selected configuration datawith known test results for some Datcom method output and build a new
configuration based on existing data. This option is most useful for theoretically
expanding a wind tunnel test data base for analysis of non-tested configurations.

Page 28 of67Additions to Figure 5

Figure 5a - Vertical Stabilizer terms

Symbol Namelist Variable Description
Units
ΦV
SYNTHS PHIV Cant angle of the vertical stabilizer from vertical. deg
Positive is outward, but Datcom does notdistinguish between negative and positive cant.
Datcom only calculates changes in Cnβ, Clβ, CyβΦF
SYNTHS
PHIF
Cant angle of the vertical fin from vertical.
deg
Positive is outward, but Datcom does notdistinguish between negative and positive cant.
Datcom only calculates changes in Cnβ, Clβ, CyβYVSYNTHSYV
Lateral offset of the vertical stabilizer from thelaircraft centerline. If present, Datcom assumes 2
vertical stabilizers, and all terms in the VTPLNF
namelist are duplicated. No interaction effects
between twin vertical stabilizers are calculated.
YF
SYNTHSYF
Lateral offset of the vertical fin from the aircraftlcenterline. If present, Datcom assumes 2 verticalfins, and all terms in the VFPLNF namelist are
duplicated. No interaction effects between twin
vertical stabilizers are calculated.

April 20113.4 GROUP III INPUT DATAThe namelists required for additional or “special” configuration definition arepresented in Figures 12 through 22, and Tables 10 through 12. Specifically, thenamelists PROPWR, JETPWR, GRNDEF, TVTPAN, ASYFLP and CONTAB
enable the user to “build upon” the configuration defined through Group II inputs.
The remaining namelists LARWB, TRNJET and HYPEFF define “stand alone”
configurations whose namelists are not used with Group II inputs.
The inputs for propellor power or jet power effects are made through namelists
PROPWR and JETPWE, respectively. The number of engines allowable is one or
two and the engines may be located anywhere on the configuration. The
configuration must have a body and a wing defined and, optionally, a horizontal
tail and a vertical tail. Since the Datcom method accounts for incremental
aerodynamic effects due to power, configuration changes required to account for
proper placement of the engine(s) on the configuration (e.g., pylons) are not taken
into account.
Twin vertical panels, defined by namelist TVTPAN, can be defined on either thewing or horizontal tail. Since the method only computes the incremental lateralstability results, “end-plate” effects on the longitudinal characteristics are notcalculated. If the twin vertical panels are present on the horizontal tail, and a
vertical tail or ventral fin is specified, the mutual interference among the panels is
not computed.
Inputs for the high lift and control devices are made with the namelists SYMFLP,ASYFLP and CONTAB. In general, the eight flap types defined using SYMFLP(variable FTYPE) are assumed to be located on the most aft lifting surface, either
horizontal tail or wing if a horizontal tail is not defined. Jet flaps, also definedusing SYMFLP, will always be located on the wing, even with the presence of ahorizontal tail. Control tabs (namelist CONTAB) are assumed to be mounted an aplain trailing edge flap (FTYPE=1); therefore, for a control tab analysis namelists
CONTAB and SYMFLP (with FTYPE=1) must both be input. For ASYFLP
namelist inputs, the spoiler and aileron devices (STYPE of 1., 2., 3. or 4.) aredefined for the wing, even with the presence of a horizontal tail, whereas the
all-moveable horizontal tail (STYPE=5.0) is, of course, a horizontal tail device.

Page 30 of673.5 GROUP IV INPUT DATA
Case control cards are provided to give the user case control and optional
input/output flexibility.
All Datcom control cards must start in card Column 1. The control card namecannot contain any embedded blanks unless the name consists of two words; they
are then separated by a single blank. All but the case termination card (NEXT
CASE) may be inserted anywhere within a case (including the middle of any
namelist). Each control card is defined below and examples of their usage areillustrated in the example problems of Section 7.
3.5.1 Case Control
NAMELIST - When this card is encountered, the content of each applicablenamelist is dumped for the case in the input system of units. This option isrecommended if there is doubt about the input values being used, especially when
the SAVE option has been used.
SAVE - When this control card is present in a case, input data for the case arepreserved for use in following case. Thus, data encountered in the following casewill update the saved data. Values not input in the new case will remainunchanged. Use of the SAVE card also allows minimum inputs for multiple casejobs. The total number of appearances of all namelists in consecutive SAVE cases
cannot exceed 300; this includes multiple appearances of the same namelist. An
error message is printed and the case is terminated if the 300 namelist limit isexceeded. Note, if both SAVE and NEXT CASE cards appear in the last input
case, the last case will be executed twice.
The NACA, DERIV and DIM control cards are the only control cards affected by
the SAVE card; i.e., no other control cards can be saved from case to case.
DIM FT, DIM IN, DIM M, DIM CM
When any of these cards are encountered, the input and output data are specified
in the stated system of units. (See Table 8.) DIM FT is the default.
NEXT CASE - When this card is encountered, the program terminates the readingof input data and begins execution of the case. Case data are destroyed following
execution of a case unless a SAVE card is present. The presence of this cardbehind last input case is optional.
3.5.2 Execution Control
TRIM - If this card is included in the case input, trim calculations will be
performed for each subsonic Mach number within the case. A vehicle may betrimmed by deflecting a control device on the wing or horizontal tail or by
deflecting an all-movable horizontal stabilizer.
Page 35 of67

DAMP - The presence of this card in a case will provide dynamic derivative
results (for addressable configurations) In addition to the standard static-derivative
output (see Figure 25).
NACA - This card provides in NACA airfoil section. designation (or supersonic
airfoil definition) for use in the airfoil section module. It is used in conjunction
with, or in place of, the airfoil section characteristics namelists, Figure 8. The
airfoil section module calculates the airfoil section characteristics designated in
Figure 8, and is executed if either a NACA control card is present or the variable
TYPEIN is defined in the appropriate section characteristic namelist (WGSCHR,HTSCHR, VTSCHR or VFSCHR). Note that if airfoil coordinates and the NACA
card are specified for the same aerodynamic surface, the airfoil coordinate
specification will be used. Therefore, if coordinates have been specified in aprevious case and the SAVE option is in effect, “TYPEIN” must be set equal to“UNUSED” for the presence of an NACA card to be recognized for that
aerodynamic surface. The airfoil designated with card will be used for bothpanels of cranked or double-delta planforms.
The form of this control card and the required parameters are given below.
Card Column(s) input(s)1 thru 4
NACA

PurposeThe unique letters NACA
designate that at airfoil
is to be defined5
6

Any delimiter
W, H, V, or F

7
8

Any delimiter
1, 4, 5, 6, S

9
10 thru 80

Any delimeter
Designation

Planforms for which theairfoil designation applies
Wing(W), Horizontal tail (H),
Vertical Tail (V), or ventral Fin (F)Type of airfoil section;
1-series (1), 4-digit (4),5-digit (5), 6-series (6).
or supersonic (S)Input designation; columns
are free-field (blanks are ignored)

Only fifteen (15) characters are accepted in the airfoil designation. The vocabulary
consists of the numbers zero (0) through nine (9), the letter “A”, and the special
characters comma, period, hyphen and equal sign. Any characters input that are
Page 36 of67

not in the vocabulary list will be interpreted as the number zero (0).
Section designation input restrictions inherent to the Airfoil Section Module arepresented in Table 13.
3.5.3 Output Control
CASEID - This card provides a case identification that. is printed as part of theoutput headings. This identification can be any user defined case title, and must
appear in card columns 7 through 80.
DUMP NAME1, NAME2, ... - This card is used to print the contents of the named
arrays in the foot-pound-second system of units. The arrays that can be listed anddefinition of their contents are given in Appendix C. For example, if the control
card read was “DUMP FLC, A” the flight conditions array FLC and the wingarray A would be printed prior to the conventional output. If more names aredesired than can fit in the available space on one card, additional dump cards may
be included.
DUMP CASE - This card is similar to the “DUMP NAME1, ...” control card.
When this card is present in a case, all the arrays (defined in Appendix C) that areused during case execution are printed prior to the conventional out put. The
values in the arrays are in the foot-pound-second system of units.
DUMP INPUT - This card is similar to the “DUMP CASE” card except that it
forces a dump of all input data blocks used for the case.
DUMP IOM - This card is similar to the “DUMP CASE” card except that all the
output arrays for the case are dumped.
DUMP ALL - This card is similar to the “DUMP CASE” card. Its use dumps all
program arrays, even if not used for the case.
DERIV RAD - This card causes the static and dynamic stability der±vatives to be
output in radian measure. The output will be in degree measure unless this flag isset. The flag remains set until a DERIV DEG control card is encountered, even if“NEXT CASE” cards are subsequently encountered.
DERIV DEG - This card causes the static and dynamic stability derivatives to be
output in degree measure. The remaining characteristics of this control card are
the same as the DERIV RAD card. DERIV DEG is the default.
PART - This card provides auxiliary and partial outputs at each Mach number inthe case (see Section 6.1.8). These outputs are automatically provided for all cases
at transonic Mach numbers.
BUILD - This control card provides configuration build-up data. Conventional
static and dynamic stability data are output for all of the applicable basic
configuration combinations shown in Table 2.
PLOT - This control card causes data generated by the program to be written to
logical unit 13, which can be retained for input to the Plot Module (described inPage 37 of

67

Volume III). The format of this plot file is described in Section 5 of Volume III.

Page 38 of67

3.4. REPRESENTATIVE CASE SETUP
Figures 23 and 24 illustrate a typical case setup utilizing the namelists and control
cards described. Though namelists (and control cards) may appear in any order(except for NEXT CASE), users are encouraged to provide inputs in the datagroups outlined In this section in order to avoid one of the most common input
errors - neglecting an important namelist input. The user's kit (Appendix D) has
been assigned to assist the user in eliminating many common input errors, and its
use is encouraged.

Page 40 of

67SECTION 4BASIC CONFIGURATION MODELING TECHNIQUES
4.1 COMPONENT CONFIGURATION MODELING
Use of the Datcom methods requires engineering judgement and experience toproperly model a configuration and interpret results. The same holds true in theuse of the Digital Datcom program. As a convenience to the user, the program
performs intermediate geometric computations (e.g., area and aspect ratio)required in method applications. The user can retrieve the values used for key
geometric parameters by means of the PART and/or DUMP options, Section 3.5.
The geometric inputs to the Digital Datcom program are relatively simple except
for the judgement required in best representing a particular configuration. Thissection describes some geometry modeling techniques to appropriately model aconfiguration.
4.1.1 Body Modeling
The basic body geometry parameters required (regardless of speed regime) consistof the longitudinal coordinates with corresponding planform half widths,
peripheries, and/or cross-sectional areas. These values are usually used in a linear
sense (e.g., the trapezoidal rule is used to integrate for planform area). Thisimplies that body-shape parameters are linearly connected. However, geometricderivatives, such as (dS/dx)i, are obtained from quadratic interpolations. Proper
modeling techniques which reflect a knowledge of method implementation, whenused in conjunction with the PART and DUMP options, greatly enhance theprogram capability and accuracy.
Body methods for lift-curve slope, pitching-moment slope and drag coefficient in
the transonic, supersonic, and hypersonic speed regimes require the body to besynthesized from a combination of body segments. The body segments consist of anose segment, an afterbody segment, and a tail segment. However, in these speed
regimes, lift and pitching-moment coefficients versus angle of attack are defined
as functions of the body planform characteristics, and therefore are not necessarily
a function of the body-segment parameters.
The program performs the configuration synthesis computations as described
below. The body input parameters R, P, and S (defined in Figure 6) can reflect
actual body contours. Digital Datcom will interpolate the R array at x=lN, x=lN+lA,
and the last input x for dN, d1, and d2, respectively. Using the shape parameters
Bnose and Btail it will synthesize an “equivalent” body from the various possibilities
shown in Figure 6. For example, in the center body X = lN x=lN, x=lN+lA, to X =lN + la, will be treated as a cylinder with a fineness ratio of 2la/(dN+d1), the nose
will be the shape specified by Bnose with a fineness ratio of lN/dN, etc. Thus, it is up
to the user to choose lN, lA, Bnose, and Btail to derive a reasonable approximation ofPage 43 of67

the actual body.
Digital Datcom requires synthesized body configurations to be either nose-alone,nose-afterbody, nose-afterbody-tail, or nose-tail (see Figure 6). The shape of the
body segments is restricted as follows: nose and tail shapes must be either anogive or cone, afterbodies must be cylindrical while tails may be either boat-tailed
or flared. Additional body namelist inputs are required to define these bodysegments and consist of nose- and tail-shape parameters BNOSE and BTAIL andnose and afterbody length parameters BLN and BLA. In the hypersonic speed
regime, the effects of nose bluntness may be obtained by specifying DS, the nose
bluntness diameter.
For an example of inputs for BLN (length of nose) and BLA (length of cylindricalafterbody) as required in speed regimes other than subsonic, the reader is directed
to Figure 6. Body diameters at the various segment intersections, dN, dl, and d2,
are obtained from linear interpolation. The tail length, lBT, is obtained bysubtracting segments BLN and BLA from the total body length.
Most Digital Datcom analyses assume bodies are axisymmetric. Users may obtain
limited results for cambered bodies of arbitrary cross section by specifying the
BODY namelist optional inputs ZU and ZL. This option is restricted to thelongitudinal stability results in the subsonic speed regime. At speeds other thansubsonic, ZU and ZL values are ignored and axisymmetric body results areprovided. It ih recommended that the reference plane for ZU and ZL inputs bechosen near the base area centroid.
The body modeling example problem (Section 7. problem 1) was selected
specifically to illustrate modeling techniques and relevant program operations.
They include:•
Choice of longitudinal coordinates that reflect body curvature and criticalbody intersections, i.e.. wing-body intersection, and body segmentation, if
required.
•
Subsonic cambered body modeling.
•
Use of the DUMP option so that key parameters can be obtained with theaid of Appendix C.
4.1.2 Wing-Tail Modeling
Input data for wings, horizontal tails, vertical tails and ventral fins have beenclassified as either planform data or as section characteristic data, as shown in
Figures 7 and 8 of Section 3. Twin-vertical panel planform input data is shown in
Figure 15.
Classification of nonstraight-tapered wings and horizontal tails as either cranked(aspect ratio > 3) or double delta (aspect ratio < 3) is relevant to only the subsonic
speed regime. In this speed regime, the appropriate lift and drag prediction
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methods depend on the classification of the lifting surface. Digital Datcom
executes subsonic analyses according to the user-specified classification
regardless of the surface aspect ratio. However, if the surface is inappropriately
designated, a warning message is printed.
Dihedral angle inputs are used primarily in the lateral stability methods. The
longitudinal stability methods reflect only the effects of dihedral in the downwash
and ground effect calculations. The direct effects of dihedral on the primary lift ofhorizontal surfaces are not defined in Datcom and are therefore not included inDigital Datcom.
Digital Datcom wing or horizontal tail alone analysis requires the exposedsemispan and the theoretical semispan to be set to the same value in namelist
WGPLNF and HTPLNF. The input wing root chord should be consistent with thechosen semispan. The reference parameters in namelist OPTINS should be used to
specify reference parameters corresponding to other than the theoretical wing
planform. If the reference parameters are not specified, they are evaluated using
the theoretical wing inputs and the reference area is set as the wing theoreticalarea, the longitudinal reference length as the wing mean aerodynamic chord, andthe lateral reference length is set as the wing span.
Horizontal tail input parameters SVWB, WVB, and SVHB, as well as vertical tail
input parameters SHB, SEXT, and RLPH, are required only for the supersonic andhypersonic speed regimes. They are used in calculation of lateral-stability
derivatives. If these data are not input, the program will calculate them, but will
fail if any part of the exposed root chord lies off of the body; lateral stability
calculations are not performed if this occurs.
Two-dimensional airfoil section characteristic data for wings and tails are input
via namelists WGSCHR, HTSCHR, VTSCHR, and VFSCHR, or may be
calculated using the airfoil section module. On occasion, the section
characteristics cannot be explicitly defined because airfoil sections either varywith span (an average airfoil section may be specified), or the planform is notstraight tapered and has different airfoil sections between the panels. In suchcircumstances, inputs should be estimated after reviewing existing airfoil test data.
Sensitivity of program results to the estimated section characteristics can bereadily evaluated by performing parametric studies utilizing the SAVE and NEXT
CASE options defined in Section 3.5. Users are warned that airfoil sensitivities doexist for low Reynolds numbers, i.e., on the order of 100,000. These namelists canalso be used to specify the aspect ratio criteria using “ARCL” (Table 9).
Planform geometry, section characteristic parameters, and synthesis dimensions
for twin vertical panels are input via namelist TVTPAN. The effects of suchpanels are reflected in only the subsonic lateral-stability output. The panels may belocated either on the wing or on the horizontal tail.
Page 45 of67

4.2 MULTIPLE COMPONENT MODELING
Combinations of aerodynamic components must be synthesized in namelist
SYNTHS. However, the program makes no cross checks in assembly ofcomponents for configuration analysis. The user must confirm the geometry inputs
to assure consistency of dimensions and component locations in total
configuration representation.
4.2.1 Wing-Body-Tail-body Modeling
Body values employed in wing-body computations are not the same as body-alone
results but are obtained by performing body-alone analysis for that portion of the
body forward of the exposed root chord of the wing. User supplied body data,
input via the namelist EXPRnn, will be used in lieu of the “nose segment” data
calculated. Carryover factors are a function of the ratio of body diameter to wing
span, as obtained from the wing input data, i.e., the body diameter is taken as
twice the difference of the exposed semispan and the theoretical semispan. Hence,
the body radius input in namelist BODY does not affect the interference
parameters.
4.2.2 Wing-Body-Tail Modeling
A conventional “aircraft” configuration is modeled using the body, wing,horizontal tail, and vertical tail modeling techniques previously described. Wing
downwash data are required to complete analysis of configurations with a wing
and horizontal tail. Subsonic and supersonic downwash data are calculated forstraight-tapered wings. For other wing planforms, or at transonic Mach numbers,
the downwash data (qH/q, , and d/d) must be supplied using the experimental
data substitution option, though two alternatives are suggested:
6.
Actual data, or from a wing-body-tail configuration which has an“equivalent” straight tapered wing, or
7.
Defining an “equivalent” straight tapered wing and substituting the wingbody results obtained from the previous Digital Datcom run to obtain thebest analytical estimate of the configuration.
8.
Body-canard-wing configurations are simulated using the standard body-wing-tailinputs. The forward surface (canard) is input as the wing, and the aft lifting
surface as the horizontal tail. Digital Datcom checks the relative span of the wing
and horizontal tail to determine if the configuration is a conventional wing-bodytail or a canard configuration.
4.2.3 Configuration Build-up Considerations
Section 3.5 describes multiple case control cards which simplify inputs forparametric and configuration build-ups. There are a few items to keep in mind.
Page 46 of

67

The effect of omitting an input variable or setting its value to zero may not be thesame, since all inputs are initialized to “UNUSED”, 1.0E-60 for CDC computers.
However, the “UNUSED” value say be used to give the effect of an input variablebeing omitted. For example, If “KSHARP” in namelist WGSCHR was specified ina previous SAVE case, a subsequent case could specify “KSHARP = 1.0E-60"(for CDC computers) which would result in KSHARP being omitted in thesubsequent case. In many places Digital Datcom uses the presence of a namelist
for program control. For example, the program assumes a body has been input if
the namelist BODY exists in a case. The effects of a presence of a namelist,
through case input or a SAVE card, cannot be eliminated even if all input volume
are set to “UNUSED.” The only exception to this rule involves high-lift andcontrol input. Either namelist SYMFLP or ASYFLP way be specified in a case,but not both. In a case sequence involving namelist SYMFLP and a SAVE card,followed by another case where ASYFLP is specified, the ASYFLP analysis will
be performed and the previous SYMFLF input ignored.
4.3 DYNAMIC DERIVATIVES
Digital Datcom computes dynamic derivatives for body, wing, wing-body, end
wing-body-tall configurations for subsonic, transonic, and supersonic speeds. In
addition, body-alone derivatives are available at hypersonic speeds. There is no
special namelist input associated with dynamic derivatives. Use of the DUMP
control card discussed in Section 3.5 will initiate computation. If experimental
data are input, the dynamic derivative methods will employ the relevant
experimental data. Dynamic derivative solutions are proviled for basic geometry
only. and the effects of high-lift and control devices are not recognized.
The experimental data option of the program permits the user to substituteexperimental data for key static stability parameters involved in dynamic
derivative solutions such as body CL, wing-body CL, etc. Any improvement in theaccuracy of these parameters will produce significant improvement in the dynamic
stability estimates. Use of experimental data substitution for this purpose isstrongly recommended.
4.4 TRIM OPTION
Digital Datcom provides a trim option allows users to obtain longitudinal trim
data. Two types of capability are provided: control device on wing or tail (Section
3.4) and the all-movable horizontal stabilizer. Trim with a control device on thewing or tail is activated by the presence of the namelist SYMFLP (Section 3.4)
and TRIM control card (Section 3.5) In the same case. Output consists of
aerodynamic increments associated with each flap deflection; similar output isprovided at trim deflection angles. The trim output is generated as follows theundeflected total configuration moment at each angle of attack is compared with
Page 47 of67

the incremental moments generated from SYMFLP input. Once the incremental
moment is reached, the corresponding deflection angle is the trim deflection angle.
The trim deflection is then used as the independent variable in table look-ups forthe remaining increments, such as CL and CD. The user should specify a liberal
range of flap deflection angles when using the control device trim option.
4.5 SUBSTITUTION OF EXPERIMENTAL DATA
Users have the option of substituting certain experimental data that will be used inlieu of Digital Datcon results. The experimental data are used in subsequentconfiguration analysis, e.g., body data are used in the wing-body and wing-bodytail calculations. Experimental data are input via namelist EXPRnn, Figure 11. All
specified parameters must be based on the same reference area and length used byDigital Datcom.
In the transonic Mach regime, some Datcom methods are available that require
user supplied data to complete the calculations. For example, Datcom methods are
given that define wing Ct,/CL and CDL/CL 2 although methods are not available
for CL. If the wing lift coefficient is supplied using experiental data substitution,
CL and CD can be calculated at each angle of attack for which CL is given. The
additional transonic data that can be calculatcd, and the “experimental” datarequired, are defined in Figure 10.

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SECTION 5
ADDITIONAL CONFIGURATION MODELING TECHNIQUES
5.1 HIGH-LIFT AND CONTROL CONFIGURATIONSControl-device input data for symmetrical and asymmetrical deflections arecontained in namelist SYMFLP and ASYFLP respectively. Analysis is limited toeither symmetrical or asymmetrical results in any one case. Multiple case runs
involving SAVE cards, may interchange symmetrical and asymmetrical analyses
from case to case. Only one control device, on either the wing or horizontal tail,may be analyzed per case. If a wing or wing-body case is run, flap input
automatically refers to the wing geometry. However, if a wing-body-horizontaltail case is input, flap input data refer to the horizontal tail. Multiple-device
analysis must be performed manually by using the experimental-data input option.
Symmetrical and asymmetrical flap analysis (namelists SYMFLP and ASYFLP)
are not performed in the hypersonic speed regime (hypersonic flap effectivenessinputs are made via namelist HYPEFF). No distinction is made between high lift
devices and control devices within the program. For instance, trim data may beobtained with any device for which the pitching moment increment is output, withthe exception of leading edge flaps. Jet flap analysis assume the flaps are on thewing and the increments are for a wing-body configuration.
5.2 POWER AND GROUND EFFECTS
Input parameters required to calculate the effects of propeller power, jet power,
and ground proximity on the subsonic longitudinal-stability results are input via
namelists PKOPWR, JETPWR, and GRNDEF. The effects of power or groundproximity on the static longitudinal stability results way be obtained for any wingbody or wing-body-horizontal. tail-and/or vertical
tail configuration. Output consists of lift, drag, and pitching moment coefficientsthat include the effects of power or ground proximity. Ground effect output may
be obtained at a maximum of ten different ground heights. It should be noted thatthe of ground height usually become negligible when the ground height exceeds
the wing span.
The effects of ground proximity on a wing-body configuration with symmetricalflaps can be calculated for as many as nine flap deflections at each ground height.
The required data are input via namelists GRNDEF and SYMFLP.
5.3 LOW-ASPECT-RATIO WING OR WING-BODY
The Datcom provides special methods to analyze low aspect ratio wing and wingbody combinations (lifting-body vehicles) in the subsonic speed regime.
Parameters required to calculate the subsonic longitudinal and lateral results forlifting bodies are input via namelist LARWB. Digital Datcom output provides
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longitudinal coefficients CL, CD, CN, CA, and Cm and the derivatives
C L α , C m α , C Y β , an d C l β

5.4 TRANSVERSE-JET CONTROL EFFECTIVENESS
A flat plate equipped with a transverse-jet control system and corresponding input
data requirements for namelist TRNJET is shown in Figure 21. The free stream
Mach number, Reynolds number, and pressure are defined via namelist FLTCON,
Figure 3. Estimates for the required control force can be made on the assumptionthat the center of pressure is at the nozzle. The predicted center of pressure
location is calculated by the program and obtained by dumping the JET array. If
the calculated center of pressure location disagrees with the assumption, arefinement of input data may be necessary
5.5 FLAP CONTROL EFFECTIVENESS AT HYPERSONIC SPEEDS
A flat plate with a flap control is shown in Figure 22 along with input namelist
HYPFLP. Force and moment data are predicted assuming a two-dimensional flow
field. Oblique shock relations are used in describing the flow field.

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SECTION 6
DEFINITION OF OUTPUT
Digital Datcom results are output at the Mach numbers specified in namelist
FLTCON. At each Mach number, output consists of a general heading, reference
parameters, input error messages, array dumps, and specific aerodynamic
characteristics as a function of angle of attack and/or flap deflection angle.
Separate output formats are provided for the following sets of related aerodynamic
data: static longitudinal and lateral stability, dynamic derivatives, high lift andcontrol, trim option, transverse-jet effectiveness, and control effectiveness at
hypersonic speeds. Since computer output is limited symbolically, definitions forthe output symbols used within the related output sets are given. The Datcom
engineering symbol follows the output symbol notation when appropriate, Unlessotherwise noted, all results are presented in the stability axis coordinate system.
6.1 STATIC AND DYNAMIC STABILITY OUTPUT
The primary outputs of Digital Datcom are the static and dynamic stability data for
a configuration. An example of this output is shown in Figure 25. Definitions of
the output notations are given below.
6.1.1 General headings
Case identification information is contained in the output heading and consists ofthe following: the version of Datcom from which the program methodologies arederived, the type of vehicle configuration (e.g. body alone or wing-body) for
which aerodynamic characteristics are output, and supplemental user-specified
case identification information if the CASEID control card is used.
6.1.2 Reference Parameters
Reference parameters and flight-condition output are defined as follows:•
MACH NUMBER - Mach at which output was calculated. This parameter isuser-specified In namelist FLTCON, or calculated from the altitude andvelocity inputs.
•
ALTITUDE - Altitude (if user input) at which Reynolds number was
calculated. This optional parameter is user specified in namelist FLTCON.
•VELOCITY - Freestream velocity (if user input) at which Mach number andReynolds number was calculated. This optional parameter is user specified
in namelist FLTCON.
•PRESSURE - Freestream atmospheric pressure at which output wascalculated (function of altitude). This parameter can also be user specified
in namelist FLTCON.
•TEMPERATURE - Freestream atmospheric temperature at which output
was calculated (function of altitude). This parameter can also be user
specified in namelist FLTCON.
Page 51 of

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•
•

•

•

•

•

REYNOLDS NO. - This flight condition parameter is the Reynolds number
per unit length and is user-specified (or input) in namelist FLTCON.
REFERENCE AREA - Digital Datcom aerodynamic characteristics are
based on this reference area. It is either user-specified in namelist OPTINS
or is equal to the planform area of the theoretical wing.
REFERENCE LENGTH - LONGITUDINAL - The Digital Datcom pitching
moment coefficient is based on this reference length. It is either user-specified in namelist OPTINS or is equal to the mean aerodynamic chord of thetheoretical wing.
REFERENCE LENGTH - LATERAL - The Digital Datcom yawingmoment and rolling-moment derivatives are based on this reference length.
It is either user-specified in namelist OPTINS or is set equal to the wing
span.
MOMENT REFERENCE CENTER - The moment reference center locationfor vehicle moments (and rotations). It is user-specified in namelist
SYNTHS and output as XCG(HORIZ) and ZCG(VERT).
ALPHA - This is the angle-of-attack array that is user specified in namelist
FLTCON. The angles are expressed in degrees.

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6.1.3 Static Longitudinal and Lateral Stability
Not all of the static aerodynamic- characteristics shown in Figure 25 are calculated
for each combination of vehicle configuration and speed regime, Digital Datcom
methods are not always available. Aerodynamic characteristics that are available
as output from Digital Datcom are presented in Table 2 as a function of vehicle
configuration and speed regime. Additional constraints are imposed on somederivatives; the user should consult the Methods Summary in Section 1 of the
USAF Stability and Control Datcom Handbook. The stability derivatives are
expressed per degree or per radian at the users option (see Section 3.5).
•
CD - Vehicle drag coefficient based on the reference area and presented an afunction of angle of attack. If Datcom methods are available to calculate
zero-lift drag but not to calculate CD versus alpha, the value of CD isprinted as output at the first alpha. CD is positive when the drag is an aft
acting load.
•
CL - Vehicle lift coefficient based on the reference area and presented as afunction of angle of attack. CL is positive when the lift is an up acting load.
•
CM - Vehicle pitching-moment coefficient based on the reference area andlongitudinal reference length and presented as a function of angle of attack.
Positive pitching moment causes a nose-up vehicle rotation.
•
CN - Vehicle (body axis) normal-force coefficient based on the reference
area and presented as a function of angle of attack. CN is positive when thenormal force is in the +Z direction. Refer to Figure 5 for Z-axis definition.
•
CA - Vehicle (body axis) axial-force coefficient based on the reference areaand presented as a function of angle of attack. CA Is positive when the axial
force is in the +X direction. Refer to Figure 5 for X-axis definition.
•
XCP - The distance between the vehicle moment reference center and thecenter of pressure divided by the longitudinal reference length. Positive is alocation forward of the center of gravity. If output is given only for the firstangle of attack, or for those cases where pitching moment (CM)is notcomputed, the value(s) define the aerodynamic-center location; i.e., XCP=dCm/dCL - (XCG-Xac) /c•
CLA - Derivative of lift coefficient with respect to alpha. If CLA is output
versus angle of attack, these values correspond to numerical derivatives ofthe lift curve. When a single value of CLA is output at the first angle ofattack, this output is the linear-lift-region derivative. CLA is based on thereference area.
•
CMA - Derivative of the pitching-moment coefficient with respect to alpha.
If CMA is output versus angle of attack, the values correspond to numericalderivatives of the pitching-moment curve. When a single value of CMA isPage 54 of

67

•

•

•

•

•

•

output at the first angle of attack, this output is the linear-lift-region
derivative. CMA is based on the reference area and longitudinal reference
length.
CYB - Derivative of side-force coefficient with respect to sideslip angle.
When CYB in defined independent of the angle of attack, output is printed
at the first angle of attack. CYB is based on the reference area.
CNB - Derivative of yawing-moment coefficient with respect to sideslip
angle. When CNB is defined independent of angle of attack, output isprinted at the first angle of attack. CNB is based on the reference area andlateral reference length.
CLB - Derivative of rolling-moment coefficient with respect to sideslip
angle presented as a function of angle of attack. CLB is based on thereference area and lateral reference length.
• Q/QINF - Ratio of dynamic pressure at the horizontal tail to the freestream
value presented an a function of angle of attack. When a single value of
Q/QINF is output at the first angle of attack, this output is the linear-liftregion value.
EPSILON - Downwash angle at horizontal tail expressed in degrees.
Downwash angle has the same algebraic sign as the lift coefficient. Positive
downwash implies that the local angle of attack of the horizontal tail is lessthan the free-stream angle of attack.
D(EPSLON)/D(ALPHA) - Derivative of downwash angle with respect toangle of attack. When a single value of D(EPSLON)/ D(ALPHA) is output
at the first angle of attack, it corresponds to the linear-lift-region derivative.

6.1.4 Dynamic Derivatives
Not all of the dynamic derivatives shown in Figure 25 are calculated for eachcombination of vehicle configuration and speed regime because of Datcom
limitations. Aerodynamic characteristics that are available as output from Digital
Datcom are presented in Table 2 as a function of vehicle configuration and speed
regime. See the Datcom Handbook, Section 1, for additional restrictions. Dynamic
stability derivatives are expressed per degree or per radian at the users option (seeSection 3.5).
•••

CLQ - Vehicle pitching derivative based on the product of reference areaand longitudinal reference length.
CMQ - Vehicle pitching derivative based on the product of reference area
and the square of the longitudinal reference length.
CLAD - Vehicle acceleration derivative based on the product of reference
area and longitudinal reference length.
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67

•
•
•
•
•
•

CMAD - Vehicle acceleration derivative based on the product of reference
area and the square of the longitudinal reference length.
CLP - Vehicle rolling derivative based on the product of reference area andthe square of the lateral reference length.
CYP - Vehicle rolling derivative based an the product of reference area andlateral reference length.
CNP - Vehicle rolling derivative based on the product of reference area andthe square of the lateral reference length.
CNR - Vehicle yawing derivative based on the product of reference area andthe square of the lateral reference length.
CLR - Vehicle rolling derivative based on the product of reference area andthe square of the lateral reference length.

6.1.5 High Lift and Control
This output consists of two basic categories: symmetrical deflection of high lift
and/or control devices, and asymmetrical control surfaces. The high lift/control
data follow the same sign convention an the static aerodynamic coefficients.
Available output is presented in Table 3 as a function of speed regime and control
type. Users are urged to consult the Datcom for limitations and constraints
imposed upon these characteristics. Output obtained from symmetrical flap
analysis are as follows.
•
DELTA - Control-surface streamwise deflection angle. Positive trailingedge down. Values of this array are user-specified in namelist SYMFLP.
•
D(CL) - Incremental lift coefficient in the linear-lift angle-of-attack range
due to deflection of control surface. Based on reference area and presented
as a function of deflection angle.
•
D(CM) - Incremental pitching-moment coefficient due to control surface
deflection valid in the linear lift angle-of-attack range. Based on the productof reference area and longitudinal reference length. Output is a function ofdeflection angle.
•
D(CL MAX) - Incremental maximum-lift coefficient. Based on reference
area and presented as a function of deflection angle.
•
D(CD MIN) - Incremental minimum drag coefficient due to control or flap
deflection. Based on reference area and presented as a function of deflection
angle.
•
D(CDI) - Incremental induced-drag coefficient due to flap deflection basedon reference area and presented as a function of angle-of-attack anddeflection angle.
•(CLA)D - Lift-curve slope of the deflected, translated surface based onreference area and presented as a function of deflection angle.
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•

•

(CH)A - Control-surface hinge-moment derivative due to angle of attack
based on the product of the control surface area and the control surface
chord, ScCc. A positive hinge moment will tend to rotate the flap trailingedge down.
(CH)D - Control-surface hinge-moment derivative due to control deflection
based on the product of the control surface area and the control surface
chord. A positive hinge moment will tend to rotate the flap trailing edge
down.

Output obtained from asymmetrical control surfaces are given below. Left andright are related to a forward facing observer:
•

•

•
•
•
•
•

•

DELTAL - Left lifting surface streamwise control deflection angle. Positive
trailing edge down. Values in this array are user-specified in namelist
ASYFLP.
DELTAR - Right lifting-surface streamwise control deflection angle .
Positive trailing edge down. Values in this array are user-specified in
namelist ASYFLP.
XS/C - Streamwise distance from wing leading edge to spoiler lip. Values inthis array are input via namelist ASYFLP, Figure 19.
HS/C - Projected height of spoiler measured from and normal to airfoil
mean line. Values in this array are input via namelist ASYFLP.
DD/C - Projected height of deflector for spoiler-slot-deflector control.
Values in this array are input via namelist ASYFLP.
DS/C - Projected height of spoiler control. Values in this array are input via
namelist ASYFLP.
(CL) ROLL - Incremental rolling moment coefficient due to asymmetricaldeflection of control surface based on the product of reference area andlateral reference length. Positive rolling moment is right wing down.
CN - Incremental yawing-moment coefficient due to asymmetricaldeflection of control surface based on the product of reference area andlateral reference length. Positive yawing moment is nose right.

6.1.6 Trim Option
The Digital Datcom trim option provides subsonic longitudinal characteristics atthe calculated trim deflection angle of the control device. The trim calculations
assume unaccelerated flight; i.e., the static pitching moment is set to zero without
accounting for any contribution from a non-zero pitch rate. Trim output is also
provided for an all-movable horizontal stabilizer at subsonic speeds. These data
include untrimmed stabilizer coefficients CD, CL, Cm, and the hinge moment
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coefficients, stabilizer trim incidence and trimmed stabilizer coefficients CD, Cm,
and the hinge-moment coefficient; wing-body-tail CD and CL with stabilizer at
trim deflection angle. Additional Digital Datcom symbols used in output are as
follows:••

HM - Stabilizer hinge-moment coefficient based on the product of reference
area and longitudinal reference length. Positive hinge moment will tend torotate the stabilizer leading edge up and trailing edge down.
ALIHT - Stabilizer incidence required to trim expressed in degrees. Positive
incidence, or deflection, is trailing edge down.

The all-movable horizontal stabilizer trim output is presented as a function ofangle of attack

6.1.7 Control at Hypersonic Speeds
Two types of control analyses are available at hypersonic speeds. They aretransverse-jet control and flap effectiveness.
Data output from the hypersonic flap methods are incremental normal- and axialforce coefficients, associated hinge moments, and center-of-pressure location.
These data are found from the local pressure distributions on the flap and inregions forward of the flap, The analysis includes the effects of flow separation
due to windward flap deflection. This is done by providing estimates forseparation induced pressures forward of the flap and reattachment on the flap. The
users may specify laminar or turbulent boundary layers.
The transverse control jet method requires a user-specified time history of local
flow parameters and control force required to trim or maneuver. With these data,
the minimum jet plenum pressure necessary to induce separation is calculated.
This minimum jet plenum pressure is then employed to calculate the nozzle throat
diameter and the jet plenum pressure and propellant weight requirements to trim or
maneuver the vehicle. Typical output can be seen in example problem 10.
6.1.8 Auxiliary and Partial Output
Auxiliary outputs consist of drag breakdown data, and basic configurationgeometric properties. Partial outputs consist of component and vortex interference
factors, effect of geometric parameters (e.g., dihedral and wing twist) on static anddynamic characteristics, canard effective-downwash data for transonic fairings andintermediate data that require user supplied data to complete. Typical output isshown in Figure 2b.
6.1.9 Effective Downwash
Datcom methods for configurations where the forward lifting-surface span is lessPage 58 of

67

than 1.5 times the aft lifting-surface span do not explicitly provide estimates for
either the downwash angle or gradient. However., Digital Datcom provides“effective-values” for those quantities. The canard effective downwash angle andgradient are defined as downwash data required to produce the correct wing-bodytail lift characteristics when applied to conventional configuration equations. The
effective downwash gradient, d/d, to found by equating the right hand sides ofDatcom equations 4.5.1.1-a and 4.5-1.1-b. The effective downwash angle, , isfound by equating the right hand sides of Datcom equations 4.5.1.2-a and 4.5.1.2b.

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6.2 DIGITAL DATCOM SYSTEM OUTPUT
Execution of Digital Datcom will produce a series of messages and data in
addition to the results previously discussed. This information falls into three
categories: input diagnostics and error analysis, extrapolation warning messages,
and Airfoil Section Module output. In addition to those outputs, an optional listingof the case input namelist data is available by using the NAMELIST control card(see Section 3.5).
Additional output may be obtained by using the DUMP and PART control cards.
When the DUMP option is exercised, the contents of user specified data blocks are
output prior to the conventional aerodynamic characteristics output. A list of thearrays and variables stored ft each data block is presented in Appendix C.
6.2.1 Input Error Analysis
An input diagnostic module (CONERR) checks all data in the input stream prior to
execution of any other Digital Datcom module. This module checks all namelist
and control cards and flags any errors. CONERR headings and error messages aredesigned to be self explanatory. All input cards are listed and any cards containing
errors have the appropriate message written immediately to the right of the card.
An explanation of the seven messages that can be generated by CONERR are
given in Table V . CONERR will not correct any errors and the program willattempt to execute each case using the data as input by the user.
Prior to case execution, additional input error analysis its conducted to insure that
all namelists essential to the case are present. This analysis will abort only thosecases missing an essential namelist. The messages that can be produced by thisanalysis are given in Table 15.
6.2.2 Extrapolation Messages
Extrapolation messages are produced when the independent variable range of theDatcom figures (nomographs/design charts) have been exceeded. Those messages
identify the number of the figure involved, the independent variable values
currently being used, the resultant value of the dependent variable, the type of
extrapolation that was used to generate the dependent variables and the name of
the table lookup routine and the subroutines that contains the figure. They are
printed primarily to alert users when the normal limit of Datcom figures has been
exceeded so that the user can determine the credibility of the results. The messages
are listed at the end of the ease output. Extrapolation message interpretation is
illustrated in Figure 27. The extrapolation messages are written to a computer
system “scratch tape” as they are generated. At the conclusion of the case they areread and sorted by figure number within each program overlay. In this way all,extrapolations for a single figure produced in a method nodule are output together
for convenience. Note that these extrapolation messages are not necessarily output
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in their order of occurrence in the program.

Page 62 of

67TABLE 14 - CONERR ERROR MESSAGESERROR MESSAGE

EXPLANATION** ERROR ** UNKNOWN NAMELIST NAMENAMELIST NAME NOT RECOGNIZED.

** MISSING NAMELIST TERMINATION ADDEDNAMELIST TERMINATION NOT FOUND.

** ERROR NO NAMELIST NAME FOLLOWING $FIRST NAMELIST CARD DOES NOT CONTAIN
A NAMELIST NAME.

** ERROR ** N*A N*B N*C N*D N*E N*FERROR FOUND ON THE CARD, N* DENOTESTHE NUMBER OF OCCURRENCES OF EACH
ERROR
A - UNKNOWN VARIABLE NAME
B - MISSING EQUAL SIGN FOLLOWINGVARIABLE NAME
C - NON-ARRAY VARIABLE HAS AN ARRAYDESIGNATION, (N)D - NON-ARRAY HAS MULTIPLE VALUESASSIGNED
E - ASSIGNED VALUES EXCEED ARRAYDIMENSIONF - SYNTAX ERROR** ILLEGAL CONTROL CARDCONTROL CARD NOT RECOGNIZED.

** ERROR ** N INCORRECT ARRAY NAMESON A DUMP CARD, "N" ARRAY NAMES WEREINCORRECT** ERROR 11** INCORRECT LIFTINGSURFACE DESIGNATION ON NACA CARD

COLUMN 6 OF THE NACA CARD DOES NOTCONTAIN W, H, V OR F.

Page 63 of

67TABLE 15 - CASE ERROR MESSAGESMESSAGEEXPLANATIONERROR ** FLAP INBOARD EDGE SPANI =XXX, IS INSIDE THE BODY AS DEFINED BY
SSPN AND SSPNE. SPANI IS REDEFINED,SPANI=SSPN-SSPNE=XXX.

THE FLAP INBOARD FLAP STATION, bi/2,
DEFINED IN NAMELIST SYMFLP OR ASYFLP
LIES INSIDE THE BODY AS DEFINED BY THE
TOTAL SPAN AND EXPOSED SPAN, b/2 AND
b*/2, IN THE PLANFORM NAMELIST.

ERROR-FLIGHT CONDITIONS NOT PRESENTMISSING NAME *FLTCON*NAMELIST "FLTCON'' NOT INPUTERROR-SYNTHESIS DATA MISSING-MISSING
NAME *SYNTHS*NAMELIST "SYNTHS" NOT INPUTERROR-WING PLANFORM PRESENT BUT
SECTION CHARACTERISTICS ABSENTMISSING NAME *WGSCHR*NAMELIST ''WGSCHR" OR "NACA-W"
CONTROL CARD NOT INPUTERROR-WING SECTION CHARACTERISTICSPRESENT BUT PLANFORM ABSENT-MISSING
NAME *WGPLNF*

NAMELIST "WGPLNF" NOT INPUTERROR-HORIZONTAL TAIL PLANFORM
PRESENT BUT SECTION CHARACTERISTICS
ABSENT-MISSING NAME *HTSCHR*NAMELIST "HTSCHR" OR "NACA - H"
CONTROL CARD NOT INPUTERROR-HORIZONTAL TAIL SECTION
CHARACTERISTICS PRESENT BUT
PLANFORM ABSENT-MISSING NAME*HTPLNF*NAMELIST "HTPLNF'' NOT INPUTERROR-VERTICAL TAIL PLANFORM
PRESENT BUT SECTION CHARACTERISTICS
ABSENT-MISSING NAME *VTSCHR*NAMELIST "VTSCHR" OR "NACA - V"
CONTROL CARD NOT INPUTERROR-VERTICAL TAIL SECTION
CHARACTERISTICS PRESENT BUT
PLANFORM ABSENT - MISSING NAME*VTPLNF*NAMELIST "VTPLNF” NOT INPUTERROR-VENTRAL FIN PLANFORM PRESENTBUT SECTION CHARACTERISTICS ABSENTMISSING NAME *VFSCHR*NAMELIST "VFSCHR" OR "NACA-F" CONTROLCARD NOT INPUTERROR-VENTRAL FIN SECTION
CHARACTERISTICS PRESENT BUT
PLANFORM ABSENT-MISSING NAME
*VFPLNF*NAMELIST "VFPLNF" NOT INPUTTHIS CASE ABORTED FOR THE ABOVEREASON(S), ALL NAMES REFER TONAMELIST NAMESTHIS CASE WILL NOT BE EXECUTED, THE
NEXT CASE WILL BE ATTEMPTED.

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6.2.3 Airfoil Section Module
The Airfoil Section Module is executed whenever airfoil section characteristics
are to be calculated. Output consists of section coordinates and a listing of thecalculated section characteristics.

Page 65 of

67SECTION 7
EXAMPLE PROBLEMS
Eleven sample problems have been selected to illustrate the modeling techniques
described in Section 4 as well as the use of the input namelist and control cards.
The paragraphs below describe each of the example problems selected forillustrating the program setup of the configurations described in Sections 4 and 5.
The input data for each example problem is presented. [The complete input andoutput for each case will be found on the CD-ROM]
7.1 EXAMPLE PROBLEM 1
Figure 28 shows three body configurations along with selected X coordinates
where shape parameters would be specified. Notice the concentration of points
used to define curvature and abrupt changes in body contours. Configuration (c)is chosen as the Problem 1 example to illustrate the body alone analysis at all
speed regimes.
Subsonic body analyses are obtained for an approximate axisymmetric body andfor a cambered body.
A summary of the four cases in problem 1 is given below:
Case No.

Configuration

Mach No.

Comments

1

Body

0.6

Axisymetric solution

2

Body

0.6

Cambered solution

3

Body

0.9.1.40,2.5

Supersonic analysis at Mach
No. 1.4 and 2.5

4

Body

2.5

Hypersonic analysis

This problem illustrates the use of the CASEID, DUMP CASE, SAVE, and NEXT
CASE control cards.

1

$FLTCON NMACH = 1.0, MACH(1)=0.60,
NALPHA = 11.0,
ALSCHD(1) = -6.0, -4.0, -2.0, 0.0, 2.0, 4.0, 8.0, 12.0, 16.0, 20.0,
RNNUB=4.28E6$$OPTINS SREF=8.85, CBARR=2.48, BLREF=4.28$$SYNTHS XCG=4.14, ZCG=-0.20$$BODY NX = 10.0,
X(1)=0.0,0.258,0.589,1.26,2.26,2.59,2.93,3.59,4.57,6.26,
R(1)=0.0,0.106,0.206,0.424,0.533,0.533,0.533,0.533,0.533,0.533,
S(1)=0.0,0.080,0.160,0.323,0.751,0.883,0.939,1.032,1.032,1.032,
P(1)=0.0,1.00,1.42,2.01,3.08,3.64,3.44,3.61,3.61,3.61$$BODY BNOSE=1.0, BLN=2.59, BLA=3.67$
CASEID APPROXIMATE AXISYMMETRIC BODY SOLUTION, EXAMPLE PROBLEM 1, CASE 1
SAVE
DUMP CASE
NEXT CASE$BODY ZU(1)= -.595,-.476,-.372,-.138, .200,.334,.343,.343,.343,.343,
ZL(1)= -.595, -.715, -.754, -.805, -.868, -.868, -.868, -.868, -.868$
CASEID ASYMMETRIC (CAMBERED) BODY SOLUTION, EXAMPLE PROBLEM 1, CASE 2
SAVENEXT CASE$FLTCON NMACH=3.0, MACH(1)=0.9,1.4,2.5, RNNUB=6.4E6, 9.96E6, 17.0E6$
SAVENEXT CASE
CASEID ASYMMETRIC (CAMBERED) BODY SOLUTION, EXAMPLE PROBLEM 1, CASE 3
NEXT CASE$FLTCON NMACH=1.0, MACH(1)=2.5, RNNUB=17.86E6, HYPERS=.TRUE.$$BODY DS=0.0$
CASEID HYPERSONIC BODY SOLUTION, EXAMPLE PROBLEM 1, CASE 4
NEXT CASE

Input file ex1.inp

2

7.2 EXAMPLE PROBLEM 2
Wing alone models for straight-tapered and non-straight-tapered planforms areshown in Figure 29. The root and tip airfoil sections differ as shown in Figure 30;
therefore average values of section data are used where appropriate. Calculationand determination of section input characteristics are from the procedure andfigures of Appendix B. These input variables are also summarized in Figure 30.
The configuration analysis consists of:Case No.

Configuration

Mach No.

Comments

1

Exposed wing

0.6,0.9,1.40
2.5

Straight-tapered-wing
dump A array

2

Exposed wing

0.6

Cranked wing3

Exposed wing

0.6

Double deltaThis problem also illustrates the control of program looping using the variable
LOOP in namelist FLTCON to obtain the flight conditions. Note that cases 2 and3 use the same inputs to FLTCON, but LOOP is changed from 2 to 3.

4

$FLTCON NMACH=4.0, MACH=0.6,0.9,1.4,2.5, LOOP=1.0, NALT=4.0,
ALT=0.0, 2000.0, 40000.0, 90000.0, HYPERS=.FALSE.,
NALPHA=11.0, ALSCHD=-6.0, -4.0, -2.0, 0.0, 2.0, 4.0,
8.0, 12.0, 16.0, 20.0, 24.0$$OPTINS SREF=8.05, CBARR=2.46, BLREF=4.28$$SYNTHS XW=3.61, ZW=-.98, ALIW=2.0,XCG=4.14$$WGPLNF CHRDTP=0.64, SSPNE=1.59, SSPN=1.59,CHRDR=2.90, SAVSI=55.0,
CHSTAT=0.0,
SWAFP=0.0, TWISTA=0.0, SSPNDD=0.0, DHDADI=0.0, DHDADO=0.0,
TYPE=1.0$$WGSCHR DELTAY=2.05, XOVC=0.4, CLI=0.127, ALPHAI=0.123,
CLALPA=0.1335,
TOVC=0.11,
CLMAX(1)=1.195, CMO=-0.0262, LERI=0.0134, CAMBER=.TRUE.,
CLAMO=0.105,
TCEFF=0.055$
CASEID STRAIGHT TAPERED EXPOSED WING SOLUTION, EXAMPLE PROBLEM 2, CASE1SAVE
DUMP A
NEXT CASE$FLTCON NMACH=2.0,MACH=0.6,2.5, LOOP=2.0, NALT=2.0, ALT=0.0, 90000.0$$SYNTHS XW=2.497, ZW=0.710$$WGPLNF SSPNOP=1.11, CHRDBP=2.24, CHRDR=4.01,
SAVSI=75.1, SAVSO=55.0, TYPE=3.0$$WGSCHR TOVC=0.1, LERI=0.011, LERO=0.0138, TOVCO=0.12,
XOVCO=0.40, CMOT=-0.0262$
CASEID EXPOSED CRANKED WING SOLUTION, EXAMPLE PROBLEM 2, CASE 2
SAVENEXT CASE$FLTCON LOOP=3.0$$WGPLNF TYPE=2.0$
CASEID EXPOSED DOUBLE DELTA WING SOLUTION, EXAMPLE PROBLEM 2, CASE 3
NEXT CASE

Input file ex2.inp

57.3 EXAMPLE PROBLEM 3
Pertinent data for Example Problem 3 are presented in Figure 31. The problem
consists of a wing-body-horizontal tail-vertical-tail configuration analyzed at
subsonic and transonic Mach numbers. Results are obtained for various
combinations of the vehicle components by using the BUILD option. The secondcase utilizes experimental body and wing-body data to update subsequent Digital
Datcom configuration analyses. The remaining cases lllustrate the use of the twin
vertical panel, propeller power and jet power inputs. A summary of the various
configurations analyzed is presented below.
Case No.

Configuration

1

Wing + body + vertical-tail + horizontal-tail configuration buildup2

Wing + body + vertical-tail + horizontal-tail with body and wingbody experimental data

3

Wing + body + vertical-tail + horizontal-tail + twin-vertical-panelswith body and wing-body experimental data

4

Wing + body + vertical-tail + horizontal-tail + twin-vertical-panels +propellor power with body and wing-body experimental data

5

Wing + body + vertical-tail + horizontal-tail + twin-vertical-panels +jet power with body and wing-body experimental data

8

BUILD$FLTCON NMACH=2.0, MACH(1)=0.6,0.8,
NALPHA=9.0, ALSCHD(1)=-2.0,0.0,2.0,4.0,8.0,12.0,16.0,20.0,24.0,
RNNUB(1)=2.28E6,3.04E6$$FLTCON NMACH=3.0, MACH(1)=0.6,0.8,1.5,
RNNUB(1)=4.26E6, 6.4E6,9.96E6$$OPTINS SREF=2.25, CBARR=0.822, BLREF=3.00$$SYNTHS XCG=2.60, ZCG=0.0, XW=1.70, ZW=0.0, ALIW=0.0, XH=3.93,
ZH=0.0, ALIH=0.0, XV=3.34, VERTUP=.TRUE.$$BODY NX=10.0,BNOSE=2.0,BTAIL=1.0,BLN=1.46,BLA=1.97,
X(1)=0.0, .175,.322,.530,.850,1.46,2.5,3.43,3.97,4.57,
S(1)=0.0,.00547,.022,.0491,.0872,.136,.136,.136,.0993,.0698,
P(1)=0.0,.262,.523,.785,1.04,1.305,1.305,1.305,1.12,.866,
R(1)=0.0, .0417,.0833,.125,.1665,.208,.208,.208,.178,.138$$WGPLNF CHRDTP=0.346, SSPNE=1.29, SSPN=1.5, CHRDR=1.16, SAVSI=45.0,
CHSTAT=0.25, SWAFP=0.0, TWISTA=0.0, SSPNDD=0.0,
DHDADI=0.0, DHDADO=0.0, TYPE=1.0$$WGSCHR TOVC=0.06, DELTAY=1.3, XOVC=0.4, CLI=0.0, ALPHAI=0.0,
CLALPA(1)=0.131, CLMAX(1)=.82, CMO=0.0, LERI=.0025, CLAMO=.105$$VTPLNF CHRDTP=.42, SSPNE=.63, SSPN=0.849, CHRDR=1.02, SAVSI=28.1,
CHSTAT=.25, SWAFP=0.0, TWISTA=0.0, TYPE=1.0$$VTSCHR TOVC=0.09, XOVC=0.4, CLALPA(1)=0.141, LERI=0.0075$$WGSCHR CLMAXL=0.78$$HTPLNF CHRDTP=0.253, SSPNE=0.52, SSPN=0.67, CHRDR=0.42, SAVSI=45.0,
CHSTAT=0.25, SWAFP=0.0, TWISTA=0.0, SSPNDD=0.0, DHDADI=0.0, DHDADO=0.0,
TYPE=1.0$$HTSCHR TOVC=0.06, DELTAY=1.3, XOVC=0.4, CLI=0.0, ALPHAI=0.0,
CLALPA(1)=.131, CLMAX(1)=0.82, CMO=0.0, LERI=.0025, CLAMO=.105$
CASEID CONFIGURATION BUILDUP, EX. PROBLEM 3, CASE 1
SAVENEXT CASE$EXPR01 CLAWB(1)=0.0575, CMAWB(1)=-0.0050,
CDWB(1)=.015,.014,.015,.019,.064,.141,.216,.302,.410,
CLWB(1)=-.115,0.0,.115,.23,.47,.65,.76,.81,.90,
CMWB(1)=.010,0.0,-.010,-.020,-.038,-.002,-.013,-.013,-.020,
CLAB(1)=.002, CMAB(1)=.0039,
CDB(1)=.012,.010,.012,.013,.014,.016,.020,.030,.047,
CLB(1)=-.004,0.0,.004,.008,.012,.020,.060,.085,.1,
CMB(1)=-.0078,.0078,.020,.038,.060,.083,.110,.140,.165$$EXPR02 CLAWB(1)=.06,CLAB(1)=.002,CMAB(1)=.0039,
ALPOW=0.0, ALPLW=8.0,ACLMW=12.01, CLMW=1.39,
ALPOW=0.0, ALPLW=6.2,ACLMW=10.10, CLMH=1.02$
CASEID INCL. BODY AND WING-BODY EXPERIMENTAL DATA, EX. PROB. 3,CASE 2
SAVENEXT CASE$TVTPAN BVP=0.4, BV=0.6, BDV=0.36, BH=1.10,
SV=0.360, VPHITE=20.0, VLP=1.04, ZP=0.0$
CASEID INCL. BODY AND WING-BODY EXPERIMENTAL DATA, EX. PROB.3, CASE 3
SAVENEXT CASE$FLTCON NMACH=1.0, MACH(1)=0.6, RNNUB(1)=2.28E6$$PROPWR AIETLP=2.0,NENGSP=1.0,THSTCP=0.15,
PHALOC=0.0, PHVLOC=0.0, PRPRAD=0.4,
ENGFCT=70.0, NOPBPE=4.0,BAPR75=18.0, YP=0.0, CROT=.FALSE.$
CASEID INCL. BODY AND WING-BODY EXPERIMENTAL DATA, EX. PROB.3, CASE 4
SAVENEXT CASE$FLTCON NMACH=1.0, MACH(1)=0.6, RNNUB(1)=2.28E6$

9

$JETPWR AIETLJ=2.0, NENGSJ=1.0, THSTCJ=0.35, JIALOC=0.0,
JEVLOC=0.0, JEALOC=0.5, JINLTA=3.0, JEANGL=15.0, JEVELO=4000.0,
AMBTMP=500.0, JESTMP=2000.0, JELLOC=0.0,
JETOTP=5000.0, AMBSTP=500.0, JERAD=2.0$
CASEID INCL. BODY AND WING-BODY EXPERIMENTAL DATA, EX. PROB. 3, CASE 5
NEXT CASE

Input file ex3.inp

10

7.4 EXAMPLE PROBLEM 4
Pertinent information for Example Problem 4 is presented in Figure 32. In thisexample, a wing-body-canard configuration is analyzed in the subsonic speed
regime (Case 1). Canard and wing section data are calculated using the Airfoil
Section Module (Appendix B). Case 2 illustrates the use of the supersonic airfoil
option of the Airfoil Section Module, nonzero body nose ordinate, vehicle scale
factor, and use of metric inputs. Note that since the NACA control cards are being
used, RNNUB and MACH must be used to define the flight conditions.
$FLTCON NMACH=1.0, MACH(1)=0.6,
NALPHA=5.0, ALSCHD(1)=0.0,5.0,10.0,15.0,20.0, RNNUB(1)=3.1E6$$OPTINS SREF=694.2, CBARR=18.07, BLREF=45.6$$SYNTHS XCG=36.68, ZCG=0.0$$BODY NX=19.0,BNOSE=2.0,BTAIL=1.0,BLN=1.46,BLA=1.97,
X(1)=0.0, 2.01, 5.49, 8.975, 12.47,
15.97, 19.47, 22.89, 26.49, 30.0,
33.51, 37.02, 40.53, 44.03, 47.53,
51.02, 54.52, 57.99, 60.0,
S(1)=0.0, 2.89,
7.42, 11.32, 14.64,
17.36, 19.49, 21.0, 21.91, 22.20,
21.90, 21.0, 19.49, 17.36, 14.64,
11.33, 7.42, 2.89, 0.0,
P(1)=0.0, 1.84, 4.72, 7.21, 9.32,
11.01, 12.41, 13.36, 13.94, 14.14,
13.94, 13.36, 12.41, 11.05, 9.32,
7.21, 4.72, 1.84 , 0.0,
R(1)=0.0, 0.293, 0.752, 1.15, 1.48,
1.76, 1.97, 2.13, 2.22, 2.25,
2.22, 2.13, 1.97, 1.76, 1.48,
1.15, 0.752, 0.293, 0.0$
NACA-W-6-65A004
NACA-H-6-65A004$WGPLNF CHSTAT=0.0,
SWAFP=0.0,TWISTA=0.0,SSPNDD=0.0,DHDADI=0.0,DHDADO=0.0,TYPE=1.0$$SYNTHS XW=8.064,ZW=0.0,ALIW=0.0$$WGPLNF CHRDTP=0.0,SSPNE=6.205,SSPN=8.01,CHRDR=13.87,SAVSI=60.0$$SYNTHS XH=29.42,ZH=0.0,ALIH=0.0$$HTPLNF SSPNE=21.34,SSPN=22.82,CHRDR=26.62,SAVSI=38.52,CHSTAT=0.0,
CHRDTP=3.80,
SWAFP=0.0,TWISTA=0.0,SSPNDD=0.0,DHDADI=0.0,DHDADO=0.0,TYPE=1.0,
SHB(1)=73.5,
SEXT(1)=73.5,RLPH(1)=47.3$
CASEID BODY PLUS WING PLUS CANARD, EXAMPLE PROBLEM 4, CASE 1
NEXT CASE
DIM M$FLTCON NMACH=1.0, MACH(1)=2.0, RNNUB=6.56E6,
NALPHA=5.0, ALSCHD(1)=0.0,5.0,10.0,15.0,20.0,
NALT=1.0, ALT(1)=27400.0$$OPTINS SREF=64.4933,CBARR=5.5077,BLREF=13.9111$$SYNTHS XCG=12.1800, ZCG=0.0, SCALE=0.30$$BODY NX=19.0,BNOSE=2.0,BTAIL=2.0,BLN=9.144,BLA=0.0,

12

X(1)=1.0,
1.613, 2.673, 3.736, 4.801,
5.868, 6.934, 8.004, 9.074, 10.144,
11.214, 12.284, 13.354, 14.420, 15.487,
16.551, 17.618, 18.675, 19.288,
S(1)=0.0,
0.268, 0.689, 1.052, 1.360,
1.613, 1.811, 1.951, 2.036, 2.062,
2.085, 1.951, 1.811, 1.613, 1.360,
1.053, 0.689, 0.268, 0.0,
P(1)=0.0,
0.561, 1.439, 2.198, 2.841,
3.368, 3.783, 4.072, 4.249, 4.310,
4.249, 4.072, 3.783, 3.368, 2.841,
2.198, 1.439, 0.561, 0.0,
R(1)=0.0,
0.089, 0.229, 0.351, 0.451,
0.536, 0.600, 0.649, 0.677, 0.686,
0.677, 0.649, 0.600, 0.536, 0.451,
0.351, 0.229, 0.089, 0.0$
NACA-W-S-3-30.0-2.5-20.0
NACA-H-S-1-50.0-2.5$WGPLNF CHSTAT=0.0,
SWAFP=0.0,TWISTA=0.0,SSPNDD=0.0,DHDADI=0.0,DHDADO=0.0,TYPE=1.0$$SYNTHS XW=3.4579,ZW=0.0,ALIW=0.0$$WGPLNF CHRDTP=0.0,SSPNE=1.8913,SSPN=2.4414,CHRDR=4.2276,SAVSI=60.0$$SYNTHS XH=9.9672,ZH=0.0,ALIH=0.0$$HTPLNF SSPNE=6.5044,SSPN=6.9555,CHRDR=8.1138,SAVSI=38.52,CHSTAT=0.0,
CHRDTP=1.1582,
SWAFP=0.0,TWISTA=0.0,SSPNDD=0.0,DHDADI=0.0,DHDADO=0.0,TYPE=1.0,
SHB(1)=6.8283,
SEXT(1)=6.8284,RLPH(1)=14.4170$
CASEID BODY PLUS WING PLUS CANARD, EXAMPLE PROBLEM 4, CASE 2
NEXT CASE

Input file: ex4.inp

13

7.5 EXAMPLE PROBLEM 5
The wing-body portion of the configuration used in Example Problem 3 ismodified by attaching plain trailing-edge flaps to the wing. This example problem
is used to illustrate partial outputs and dynamic derivative input and output. A
summary of Example Problem 5 analysis is as follows:
Case No.

Configuration

Mach No.

Comments

1

Body + wing

0.6

PART, DAMP, DUMP
DYN

2Body + wing + plain
trailing-edge flaps

0.6

DUMP FCM

The Digital Datcom output data, including a dump of the DYN and FCM common
arrays may be found on the CD-ROM. The flap configuration is shown in Figure
33.
DIM FTPART$FLTCON NALPHA=9.0,
ALSCHD(1)=-2.0,0.0,2.0,4.0,8.0,12.0,16.0,20.0,24.0$$FLTCON NMACH=1.0, MACH(1)=0.6, RNNUB(1)=4.26E6$$OPTINS SREF=2.25, CBARR=0.822, BLREF=3.00$$SYNTHS XCG=2.60, ZCG=0.0, XW=1.70, ZW=0.0, ALIW=0.0$$BODY NX=10.0,BNOSE=2.0,BTAIL=1.0,BLN=1.46,BLA=1.97,
X(1)=0.0, .175,.322,.530,.850,1.46,2.5,3.43,3.97,4.57,
R(1)=0.0, .0417,.0833,.125,.1665,.208,.208,.208,.178,.138$$WGPLNF CHRDTP=0.346, SSPNE=1.29, SSPN=1.5, CHRDR=1.16, SAVSI=45.0,
CHSTAT=0.25, SWAFP=0.0, TWISTA=0.0, SSPNDD=0.0, DHDADI=0.0,
DHDADO=0.0, TYPE=1.0$$WGSCHR TOVC=.06, DELTAY=1.3, XOVC=0.4, CLI=0.0, ALPHAI=0.0,
CLALPA(1)=0.131,
CLMAX(1)=.82,CMO=0.0,LERI=.0025,CLAMO=.105$$WGSCHR CLMAXL=0.8, TCEFF=.03$
CASEID BODY-WING DAMPING DERIVATIVES, EXAMPLE PROBLEM 5, CASE 1
DAMP
SAVE
DUMP DYNNEXT CASE$SYMFLP NDELTA=6.0, DELTA(1)=0.,.10,.20,.30,.40,.60, PHETE=.0522,
CHRDFI=.2094, CHRDFO=.1554, SPANFI=0.208, SPANFO=.708, FTYPE=1.0,
CB=.01125, TC=.0225, PHETEP=.0391, NTYPE=1.0$
CASEID PLAIN FLAPS ON WING, EXAMPLE PROBLEM 5, CASE 2

15

DUMP FCM
NEXT CASE

16

7.6 EXAMPLE PROBLEM 6
The wing-body configuration of Example Problem 3 is used to illustrate aileronand spoiler input and output data. Figure 34 show the geometry.
$FLTCON NALPHA=9.0, ALSCHD=-2.0,0.0,2.0,4.0,8.0,12.0,16.0,20.0,24.0$$FLTCON NMACH=1.0, MACH(1)=0.6, RNNUB(1)=4.26E6$$OPTINS SREF=2.25, CBARR=0.822, BLREF=3.00$$SYNTHS XCG=2.60, ZCG=0.0, XW=1.70, ZW=0.0, ALIW=0.0$$BODY NX=10.0,BNOSE=2.0,BTAIL=1.0,BLN=1.46,BLA=1.97,
X(1)=0.0, .175,.322,.530,.850,1.46,2.5,3.43,3.97,4.57,
R(1)=0.0, .0417,.0833,.125,.1665,.208,.208,.208,.178,.138$$WGPLNF CHRDTP=0.346,SSPNE=1.29,SSPN=1.5,CHRDR=1.16,SAVSI=45.0,CHSTAT=0.25,
SWAFP=0.0,TWISTA=0.0,SSPNDD=0.0,DHDADI=0.0,DHDADO=0.0,TYPE=1.0$$WGSCHR TOVC=.06,DELTAY=1.3, XOVC=0.4, CLI=0.0, ALPHAI=0.0,
CLALPA(1)=0.131,
CLMAX(1)=.82,CMO=0.0,LERI=.0025,CLAMO=.105$$WGSCHR CLMAXL=0.8, TCEFF=.03$$ASYFLP NDELTA=5.0, DELTAL(1)=5.,10.,20.,30.,40.,
DELTAR(1)=-2.,-5.,-10.,-15.,-20.,
STYPE=4.0, CHRDFI=.1116, CHRDFO=.0692,
SPANFI=1.109, SPANFO=1.5, PHETE=.0522$
CASEID PLAIN FLAP AILERON, EXAMPLE PROBLEM 6, CASE 1
SAVENEXT CASE$ASYFLP STYPE=3.0, DELTAD(1)=.0130,.0261,.0380,.0513,.0630,.0750,
DELTAS(1)=.013, .0261, .030, .0513, .0630, .0750,
XSOC(1)=.6900,.6955,.6680,.6438,.6454,.6259,
XSPRME=0.55,
XSOC(1)=0.0357,0.0710,0.0956,0.1102,0.1365,0.1359$
CASEID SPOILER-SLOT-DEFLECTOR ON WING, EXAMPLE PROBLEM 6, CASE 2
NEXT CASE

Input file ex6.inp

18

7.7 EXAMPLE PROBLEM 7
The wing-body-tail configuration of Example Problem 3 is used to illustrate trim
control with an elevator on the horizontal tail. In addition, the effect of plain
trailing edge flaps on the wing (see Example Problem 5) is included via
experimental data input to illustrate a procedure for multiple high-lift and control
device analysis. The wing high lift increment is used to update wing-bodyundeflected totals via namelist EXPRnn.
The geometry is sketched in Figure 35.
$FLTCON NMACH=1.0,MACH(1)=.60, NALPHA=9.0,
ALSCHD(1)=-2.0,0.0,2.0,4.0,8.0, 12.0,16.0,20.0,24.0,
RNNUB(1)=2.28E6$$OPTINS SREF=2.25,CBARR=0.822,BLREF=3.0$$SYNTHS XCG=2.60, ZCG=0.0, XW=1.70, ZW=0.0, ALIW=0.0, XH=3.93,
ZH=0.0,ALIH=0.0, XV=3.34,VERTUP=.TRUE.$$BODY NX=10.,
X(1)=0.0, 0.175, 0.322, 0.530, 0.85,
1.46, 2.50, 3.43, 3.97, 4.57,
R(1)=0.0,
0.0417, 0.0833, 0.125, 0.1665,
0.208, 0.208, 0.208, 0.178, 0.138$$WGPLNF CHRDTP=0.346, SSPNE=1.29, SSPN=1.50, CHRDR=1.16, SAVSI=45.0,
CHSTAT=0.25, SWAFP=0.0, TWISTA=0.0, SSPNDD=0.0, DHDADI=0.0,
DHDADO=0.0, TYPE=1.0$$WGSCHR TOVC=0.060, DELTAY=1.30, XOVC=0.40, CLI=0.0, ALPHAI=0.0,
CLALPA(1)=0.131, CLMAX(1)=0.82, CMO=0.0, LERI=0.0025,
CLAMO=0.105$$WGSCHR CLMAXL=0.78$$VTPLNF CHRDTP=0.420, SSPNE=0.63, SSPN=0.849, CHRDR=1.02,
SAVSI=28.1,
CHSTAT=0.25, SWAFP=0.0, TWISTA=0.0, TYPE=1.0$$VTSCHR TOVC=0.09, XOVC=0.40, CLALPA(1)=0.141, LERI=0.0075$$HTPLNF CHRDTP=0.253, SSPNE=0.52, SSPN=0.67, CHRDR=0.42, SAVSI=45.0,
CHSTAT=0.25, SWAFP=0.0, TWISTA=0.0, SSPNDD=0.0, DHDADI=0.0,
DHDADO=0.0, TYPE=1.0$$HTSCHR TOVC=0.060, DELTAY=1.30, XOVC=0.40, CLI=0.0, ALPHAI=0.0,
CLALPA(1)=0.131, CLMAX(1)=0.82, CMO=0.0, LERI=0.0025,CLAMO=0.105$$SYMFLP FTYPE=1.0, NDELTA=9.0,
DELTA(1)=-60.0, -40.0, -20.0, -10.0, 0.0, 10.0, 20.0, 40.0, 60.0,
PHETE=0.0522, PHETEP=0.0523, SPANFI=0.18, SPANFO=0.670, CHRDFI=0.075,
CHRDFO=0.051, CB=0.0038, TC=0.0076, NTYPE=1.0,$$EXPR01 CLWB(1)=0.09,0.204,0.330,0.450,0.690,0.895,1.070,1.180,1.174$
TRIM
CASEID INCLUDES HIGH LIFT EFFECT ON WING, EXAMPLE PROBLEM 7
NEXT CASE

Input file ex7.inp

207.8 EXAMPLE PROBLEM 8
The all-movable horizontal trim case is illustrated using the configuration of
Example Problem 3. Note that a hinge-axis distance is specified in namelist
SYNTHS and a TRIM control card is present in the case.
$FLTCON NMACH=1.0, MACH(1)=0.6,
NALPHA=9.0, ALSCHD(1)=-2.0,0.0,2.0,4.0,8.0,12.0,16.0,20.0,24.0,
RNNUB(1)=2.28E6$$OPTINS SREF=2.25, CBARR=0.822, BLREF=3.00$$SYNTHS XCG=2.60, ZCG=0.0, XW=1.70, ZW=0.0, ALIW=0.0, XH=3.93,
ZH=0.0, ALIH=0.0, XV=3.34, VERTUP=.TRUE.$$SYNTHS HINAX=4.271$$BODY NX=10.0,
X(1)=0.0, .175,.322,.530,.850,1.46,2.5,3.43,3.97,4.57,
R(1)=0.0, .0417,.0833,.125,.1665,.208,.208,.208,.178,.138$$WGPLNF CHRDTP=0.346, SSPNE=1.29, SSPN=1.5, CHRDR=1.16, SAVSI=45.0,
CHSTAT=0.25, SWAFP=0.0, TWISTA=0.0, SSPNDD=0.0,
DHDADI=0.0, DHDADO=0.0, TYPE=1.0$$WGSCHR TOVC=0.06, DELTAY=1.3, XOVC=0.4, CLI=0.0, ALPHAI=0.0,
CLALPA(1)=0.131, CLMAX(1)=.82, CMO=0.0, LERI=0.0025, CLAMO=0.105$$WGSCHR CLMAXL=0.78$$VTPLNF CHRDTP=0.42, SSPNE=0.63, SSPN=0.849, CHRDR=1.02, SAVSI=28.1,
CHSTAT=.25, SWAFP=0.0, TWISTA=0.0, TYPE=1.0$$VTSCHR TOVC=0.09, XOVC=0.4, CLALPA(1)=0.141, LERI=0.0075$$HTPLNF CHRDTP=0.253, SSPNE=0.52, SSPN=0.67, CHRDR=0.42, SAVSI=45.0,
CHSTAT=0.25, SWAFP=0.0, TWISTA=0.0, SSPNDD=0.0, DHDADI=0.0,
DHDADO=0.0,
TYPE=1.0$$HTSCHR TOVC=0.06, DELTAY=1.3, XOVC=0.4, CLI=0.0, ALPHAI=0.0,
CLALPA(1)=.131, CLMAX(1)=0.82, CMO=0.0, LERI=.0025, CLAMO=.105$
CASEID ALL MOVEABLE HORIZONTAL TAIL, EXAMPLE PROBLEM 8
TRIMNEXT CASE

Input file ex8.inp

22

7.9 EXAMPLE PROBLEM 9
Problem 9 consists of a lifting body configuration with a delta planform, sharpleading edge, and symmetrical diamond cross section. Pertinent data for thisconfiguration are shown in Figure 36.
$FLTCON NMACH=1.0, MACH(1)=0.26, NALPHA=6.0,
ALSCHD=-5.0, 0.0, 5.0, 10.0, 15.0, 20.0, RNNUB(1)=1.868E6$$LARWB ZB=0.0, SREF=0.989, DELTEP=90.0, SFRONT=0.307, AR=1.076,
L=1.915, SWET=2.28, PERBAS=2.38, SBASE=0.307, HB=0.595, BB=1.03,
BLF=.FALSE., XCG=1.44, THETAD=15.0, ROUNDN=.FALSE., SBS=0.57,
SBSLB=0.0228, XCENSB=1.277, XCENW=1.277$
CASEID LIFTING BODY WITH SHARP LEADING EDGE, EXAMPLE PROBLEM 9NEXT CASE

Input file ex9.inp

23

7.10 EXAMPLE PROBLEM 10
This problem demonstrates the analysis of the transverse control jet in hypersonicflow located on a flat plate, as shown in Figure 37.
$FLTCON NMACH=1.0, MACH(1)=10.0, RNNUB(1)=1.0E7, PINF(1)=10.0,
HYPERS=.TRUE.,$$TRNJET TIME(1)=1.,2.,3.,4.,5.,
FC(1)=1000.,2000.,1000.,500.,200.,
ALPHA(1)=0.0, 3.0, 6.0, 9.0, 13.0,
LAMNRJ=.FALSE., .FALSE., .FALSE., .FALSE., .TRUE.,
ME=2.39, ISP=225.0, SPAN=2.0, PHE=30.0, GP=1.2, CC=90.0, LFP=10.0$
CASEID TRANSVERSE-JET SIZING, EXAMPLE PROBLEM 10
DUMP JET
NEXT CASE

Input file ex10.inp

25

7.11 EXAMPLE PROBLEM 11
The use of a hypersonic control flap is demonstrated in this example. Pertinent
geometry is shown in Figure 38.
$FLTCON NMACH=1.0, MACH(1)=10.0, RNNUB(1)=1.06E5,
NALPHA=5.0, ALSCHD(1)=0.0,5.0,10.0,15.0,20.0,
HYPERS=.TRUE.$$OPTINS SREF=1.0, CBARR=1.0$$HYPEFF ALITD=15000.0, XHL=8.0, TWOTI=3.122, CF=2.0,
HDELTA(1)=0.,2.,4.,6.,10.,12.,16.,20.,25.,30.,
LAMNR=.TRUE., HNDLTA=10.0$
CASEID FLAT PLATE WITH FLAP IN HYPERSONIC FLOW, EXAMPLE PROBLEM 11
NEXT CASE

Input file ex11.inp

27