The information on this page is taken verbatim from the user's manual

On this page there are references to section 3.5 and Appendix B. These are are in the user manual (see reference page).

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. Similar inefficiencies 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 a systematic summary of methods for estimating stability and control characteristics in preliminary design applications. Consistent with this philosophy, the development 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, and dynamic derivative characteristics using the methods contained in Sections 4 through 7 of the printed Datcom. The computer program also offer a trim option that computes control deflections and aerodynamic data for vehicle trim at subsonic Mach numbers.

The program has been developed an a modular basis. 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 section has been prepared to assist the potential user in his decision process concerning the applicability of the USAF Stability and Control Digital Datcom to his particular requirements. For specific questions dealing with method validity and limitations, the user is strongly encouraged to refer to the USAF Stability and Control 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 the point 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.

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 unique geometries which are identified as "special" configurations. Table 1 summarizes the addressable configurations accommodated by the program.

The capabilities discussed below apply to basic configurations, i.e., traditional body-wing-tail concepts. A detailed summary of output as a function of configuration 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 of the user's manual). Typical output for these configurations are presented in section 6 of the report.

The longitudinal and lateral-directional stability
characteristics provided by the Datcom and the Digital Datcom are
in the stability-axis system. Body-axis normal-force and
axial-force coefficients are also included in the output for
convenience of the user. For those speed regimes and configurations
where Datcom methods are available, the Digital Datcom output
provides the longitudinal coefficients C_{D},
C_{L}, C_{m}, C_{N}, and C_{A}, and
the derivatives dC_{L}/dα, dC_{m}/dα,
dC_{Y}/dβ, dC_{n}/dβ, and
dC_{l}/dβ. 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 and ground effects are limited to the longitudinal
aerodynamic characteristics.

Users are cautioned that the Datcom does not rigorously treat aerodynamics in the transonic speed regime, and a fairing between subsonic and supersonic solutions is often 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 to obtain 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 better method 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 subsonic longitudinal 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 be treated at subsonic Mach numbers. Dihedral influences are included in lateral-directional stability derivatives and wing wake location used in the calculation of longitudinal data. Airfoil section characteristics are a required input, although most of these characteristics may be generated 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 for cases involving significant wetted areas in laminar flow will require adjustment by the user.

Computations of wing-body longitudinal characteristics assume, in many cases, that the configuration is of the mid-wing type. Lateral-directional analyses do account for other wing locations. Users should consult the Datcom for specific details.

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 isolate component aerodynamic contributions in a similar fashion to break down data from 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 conventional vertical tail specified though interference effects between the three panels is not computed. The influence of twin vertical tails is included only in the lateral-directional stability characteristics at subsonic speeds.

The pitch, acceleration, roll and yaw derivatives of
dC_{L}/dq, dC_{m}/dq, dC_{L}/d(alpha-dot),
dC_{m}/d(alpha-dot), dC_{l}/dp, dC_{Y}/dp,
dC_{n}/dp, dC_{n}/dr, dC_{l}/dr 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
dC_{L}/d(alpha) and wing-body dC_{L}/d(alpha-dot).
Any improvement in the accuracy of these key parameters will
produce significant improvement in the dynamic stability estimates.
Use of experimental data substitution for this purpose is strongly
recommended.

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 be treated. 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, and pitching-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 and hinge-moment derivatives. For translating devices the lift-curve slope is also computed. Asymmetrical deflection of wing control devices can be analyzed for rolling and yawing effectiveness. Rolling effectiveness may be obtained for all-movable differentially-deflected horizontal stabilizers. The speed regimes where these capabilities exist are shown in Table 3.

Control modes employing all-moving wing or tail surfaces can also be addressed with the program. This is accomplished by executing multiple cases with a variety of panel incidence angles.

Trim data can be calculated at subsonic speeds. Digital Datcom
manipulates computed stability and control characteristics to
provide trim output (static C_{m}= 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 a
function of angle of attack and consists of control deflection
angles required to trim and the associated longitudinal aerodynamic
characteristics shown in Table 3. The second mode treats
conventional wing-body-tail configurations where the
horizontal-tail is all-movable or "flying." In this case, output as
a function of angle of attack consists of horizontal-stabilizer
deflection (or incidence) angle required to trim; untrimmed
stabilizer C_{L}, C_{D}, C_{m}, and
hinge-moment coefficients; trimmed stabilizer C_{L},
C_{D}, and hinge moment coefficients; and total
wing-body-tail C_{L} and C_{D}. Body-canard-tail
configurations may be trimmed by calculating the stability
characteristics at a variety of canard incidence angles and
manually calculating the trim data. Treatment of a canard
configuration is addressed in Table 1.

The capabilities discussed below apply to the three special configurations illustrated in Figure 2.

Datcom provides methods which apply to lifting reentry vehicles
at subsonic speeds. Digital Datcom output provides longitudinal
coefficients C_{D}, C_{L}, C_{m},
C_{N}, and C_{A} and the derivatives
dC_{L}/dα, dC_{m}/dα,
dC_{Y}/dβ, dC_{n}/dβ,
dC_{l}/dβ.

The USAF Stability and Control Datcom contains some special control methods for high-speed vehicles. These include hypersonic flap methods which are incorporated into Digital Datcom. The flap methods are restricted to Mach numbers greater than 5, angles of attack between zero and 20 degrees and deflections 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- and axial-force coefficients, associated hinge moments, and center-of-pressure location. These data are found from the local pressure distributions on the flap and in regions forward of the flap. The analysis includes the effects of flow separation due to windward flap deflection by providing estimates for separation induced-pressures forward of the flap and reattachment on the flap. Users may specify laminar or turbulent boundary layers.

Datcom provides a procedure for preliminary sizing of a two-dimensional transverse-jet control system in hypersonic flow, assuming that the nozzle is located at the aft end of the surface. The method evaluates the interaction of the transverse 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.

There are several operational considerations the user needs to understand in order to take maximum advantage of Digital Datcom.

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 the requirement are given in Figure 3. The speed reference is input as either Mach number or velocity, and the atmospheric conditions as either altitude or freestream pressure and temperature. The specific reference and atmospheric conditions are then used to calculate Reynolds number.

The program may loop on speed reference and atmospheric conditions three different ways, as given by the variable LOOP in Figure 3. In this discussion, and in Figure 3, the speed reference is referred to as Mach number, and atmospheric conditions as altitude. The three options for program looping on Mach number and altitude are listed and discussed below.

- LOOP = 1 - Vary Mach and altitude together. The program executes at the first 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, and must 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, and continues until each altitude has been selected. Atmospheric conditions oust be 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, and continues until each Mach number has been selected. Atmospheric conditions must be input for this option and NMACH times NALT flight conditions are executed.

Aerodynamic stability methods are defined in Datcom as a function of vehicle configuration and Mach regime. Digital Datcom logic determines the configuration being analyzed by identifying the particular input namelists that are present within a case (see Section 3). The Mach regime is normally determined according to the following criteria:

Mach Number (M) | Mach Regime | |
---|---|---|

M <= 0.6 | Subsonic | |

0.6 < M < 1.4 | Transonic | |

M >= 1.4 | Supersonic | |

M >= 1.4 and the hypersonic flag is set |
Hypersonic |

These limits were selected to conform with most Datcom methods. However, some methods are valid for a larger Mach number range. Some subsonic methods are valid up to a Mach number of 0.7 or 0.8. The user has the option to increase the subsonic 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 the variable 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 the range is exceeded. The Mach regimes are then defined as follows:

Mach Number (M) | Mach Regime | |
---|---|---|

M <= STMACH | Subsonic | |

STMACH < M < TSMACH | Transonic | |

M >= TSMACH | Supersonic | |

M >= TSMACH and the hypersonic flag is set |
Hypersonic |

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 name and variable name spelling, checks the numerical inputs for syntax errors, and checks 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.

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:

- An airfoil section designation (for NACA, double wedge, circular arc, or hexagonal airfoils)
- Section upper and lower Cartesian coordinates, or
- 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 the section lift curve slope, c_{lalpha}.
In addition a 5 per cent 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 B in the printed manual.

The airfoil section is assumed to be parallel to the free stream. Skewed airfoils can be 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.

- For subsonic Mach numbers, the module computes the airfoil subsonic section 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.
- 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.
- 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 be accomplished by supplying section ordinates at non-dimensional chord stations (x/c of 0.0, 0.001, 0.002, and 0.003.
- If the leading edge radius is not specified in the airfoil section input, the user must insure that the first and second coordinate points lie on the leading edge radius. For sharp nosed airfoils the user must specify a zero leading edge radius.
- The computational algorithm can be sensitive to the "smoothness" of the input coordinates. Therefore, the user should insure that the input data contains no unintentional fluctuations. Considering that Datcom procedures are preliminary design methods, it is at least as important to provide smoothly varying coordinates as it is to accurately define the airfoil geometry.

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 the nature of the configuration.
- Twin vertical tail methods are only applicable to lateral stability parameters at 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 radii and 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 be applied 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 of high lift and control devices on downwash is not calculated. The effects of multiple 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 HTPLNF causes Digital Datcom to assume that the configuration has a horizontal 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.

- Go to the Digital Datcom Main Page.
- Go to the page of references for the Digital Datcom program. This page has links to the user's manual for Digital Datcom and the original documents.
- Go to the download page to download the Digital Datcom Program.