Workshop on Digital Datcom and Simulink

DHA
Suffa University
D e p a r t m e n t o f M e c h a n i c a l E n g i n e e r i n g
Project TitleWorkshop: Digital DATCOM
Dr. Bilal Ahmed Siddiqui

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Suffa University
1. Introduction to Aerodynamics
2. Stability and Control
3. USAF DATCOM
4. Digital DATCOM
5. Enhancements
6. Missile DATCOM
7. Sample Cases
8. Some Practice
9. What’s More?
2
Contents

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Suffa University
Aerodynamics
• Aerodynamics : interplay of air & bodies trying to
move through it
– Air resists some motion…and aids some motion
– Important to understand this interplay to harness it
• Formal study of aerodynamics began 300 years
ago, so it is a relatively young science!
• Informally, we have been harnessing wind since a
long time…….really long

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Ancient Egyptian Aerospace!
https://en.wikipedia.org/wiki/Helicopter_hieroglyphs [1290 BC]

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But Preserved History is different…
• In 1799, Sir George Cayley became the first person to identify the
four aerodynamic forces of flight
• In 1871, Francis Herbert Wenham constructed the first wind
tunnel, allowing precise measurements of aerodynamic forces.
• In 1889, Charles Renard became the first person to predict the
power needed for sustained flight.
• Otto Lilienthal was the first to propose thin, curved airfoils that
would produce high lift and low drag.
• However, interestingly the Wright brothers, mechanics – not
engineers- found most of the initial work flawed and did
something else….a hundred years after Cayley.

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27th of Ramadhan….a Friday
• On Dec 15, 1903 (corresponding to Hijri date
above), Wilbur and Orville Wright made
history after failing to achieve it for 3 years.
• Cycle mechanics, enthusiastic about flight,
they designed airplanes based on
aerodynamic data published by Leinthall and
Langley.
• All attempts were splendid failures, so they
began to doubt the crude theories of their
times.
• They built themselves a windtunnel and
started testing airfoils. To their surprise, they
obtained reliable results and the rest is
history.

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The [W]Right [Bi]Plane

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Forces and Moments in Flight
• Straight level flight means constant velocity and altitude.
• There are four main forces which govern straight level flight
• For level flight, Lift=Weight and Thrust=Drag
• In other words,
𝑇
𝑊
=
𝐿
𝐷
in level flight
• Except weight, all other variables depend
on Aerodynamics.
• Aerodynamics also causes moments in all three axes

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Aerodynamic Moments
• Aerodynamics also causes moments in all three axes.
• Performance of aircraft depends on aerodynamics!

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Source of all Aerodynamic Forces
&Moments
• No matter how complex the body shape and flow, the
aerodynamic forces and moments on the body are due to only
two basic sources:
a) Pressure distribution p over the body surface
b) Shear stress distribution τ over the body surface
• Pressure varies with velocity of air over the surface and acts
normal to it. For incompressible, inviscid flow, it follows Bernoulli
principle.
• Shear stress is due to friction in the boundary layer and acts
tangent to the surface. Typically 𝜏 ≪ 𝑝

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Net Effect of Pressure and Shear
Distribution
• Each body shape and flow condition creates unique p & τ
distribution
• The net effect of the p and τ distributions integrated over
the complete body surface is a resultant aerodynamic force
R and moment M on the body.
• Far ahead of the body, the flow is undisturbed and called
free stream.
• V∞ = free stream velocity=flow velocity far ahead of the
body.

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Components of Aerodynamic Force R
• Let distance between leading and trailing edges be “chord”=c
• “Angle of attack” is the angle between 𝑉∞ and c
• R can be resolved into two sets of components: either wrt 𝑉∞ or c
•
𝐿 = 𝑙𝑖𝑓𝑡 = 𝑐𝑜𝑚𝑝𝑜𝑛𝑒𝑛𝑡 𝑝𝑒𝑟𝑝𝑒𝑛𝑑𝑖𝑐𝑢𝑙𝑎𝑟 𝑡𝑜 𝑉∞
𝐷 = 𝑑𝑟𝑎𝑔 = 𝑐𝑜𝑚𝑝𝑜𝑛𝑒𝑛𝑡 𝑝𝑎𝑟𝑎𝑙𝑙𝑒𝑙 𝑡𝑜 𝑉∞
•
𝑁 = 𝑛𝑜𝑟𝑚𝑎𝑙 𝑓𝑜𝑟𝑐𝑒 = 𝑐𝑜𝑚𝑝𝑜𝑛𝑒𝑛𝑡 𝑝𝑒𝑟𝑝𝑒𝑛𝑑𝑖𝑐𝑢𝑙𝑎𝑟 𝑡𝑜 𝑐
𝐴 = 𝑎𝑥𝑖𝑎𝑙 𝑓𝑜𝑟𝑐𝑒 = 𝑐𝑜𝑚𝑝𝑜𝑛𝑒𝑛𝑡 𝑝𝑎𝑟𝑎𝑙𝑙𝑒𝑙 𝑡𝑜 𝑐
• But, {L,D} and {N,A} are related through 𝛼
𝐿 = 𝑁 cos𝛼 − 𝐴 𝑠𝑖𝑛α
𝐷 = 𝑁 𝑠𝑖𝑛𝛼 + 𝐴 𝑐𝑜𝑠𝛼

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Source of {N,A} and {L,D}
• Pressure p, shear τ , surface slope θ are
functions of path length s.
• For a unit span l=1, the forces are
𝑁 = −
𝐿𝐸
𝑇𝐸
𝑝 𝑢 𝑐𝑜𝑠𝜃 + 𝜏 𝑢 𝑠𝑖𝑛𝜃 𝑑𝑠 𝑢
+
𝐿𝐸
𝑇𝐸
𝑝𝑙 𝑐𝑜𝑠𝜃 − 𝜏𝑙 𝑠𝑖𝑛𝜃 𝑑𝑠𝑙
𝐴 =
𝐿𝐸
𝑇𝐸
−𝑝 𝑢 𝑠𝑖𝑛𝜃 + 𝜏 𝑢 𝑐𝑜𝑠𝜃 𝑑𝑠 𝑢
+
𝐿𝐸
𝑇𝐸
𝑝𝑙 𝑠𝑖𝑛𝜃 + 𝜏𝑙 𝑐𝑜𝑠𝜃 𝑑𝑠𝑙

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Source of Aerodynamic Moment
• The aerodynamic moment exerted on the body depends on the point about
which moments are taken.
• Moments that tend to increase α (pitch up) are positive, and moments that tend
to decrease α (pitch down) are negative.
• Moment about the leading edge is simply the forces x moment arms.
𝑀𝐿𝐸 =
𝐿𝐸
𝑇𝐸
𝑝 𝑢 𝑐𝑜𝑠𝜃 + 𝜏 𝑢 𝑠𝑖𝑛𝜃 𝑥 − 𝑝 𝑢 𝑠𝑖𝑛𝜃 − 𝜏 𝑢 𝑐𝑜𝑠𝜃 𝑦 𝑑𝑠 𝑢
+
𝐿𝐸
𝑇𝐸
−𝑝𝑙 𝑐𝑜𝑠𝜃 + 𝜏𝑙 𝑠𝑖𝑛𝜃 𝑥 + 𝑝𝑙 𝑠𝑖𝑛𝜃 + 𝜏𝑙 𝑐𝑜𝑠𝜃 𝑦 𝑑𝑠𝑙
• In equations above, x, y and θ are known functions of s for given shape.
• A major goal of aerodynamics is to calculate p(s) and τ(s) for a given body shape
and freestream conditions (𝑉∞ and α) aerodynamic forces/moments

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Getting rid of the
Dimensions…convenience
• It will become clear later that it is of benefit to non-dimensionalize forces
and moments.
• Let 𝜌∞ be the free stream air density and S and l be reference area and
reference length respectively.
• Dynamic pressure, 𝑄∞ =
1
2
𝜌∞ 𝑉∞
2
• Lift coefficient, CL =
L
Q∞S
• Drag coefficient, CD =
D
Q∞S
• Lift coefficient, CL =
M
Q∞S𝑙
• Axial and normal force coefficients are similarly defined.
Coefficients makes the math
manageable. An aircraft with a 50m2
wing area and weight of 10,000kg at sea
level cruise will have a lift coefficient of
0.3 at a speed of 100m/s, rather than a
lift of 9.8x104 N.

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Suffa University
Reference Area and Length
• In these coefficients, the reference area S and
reference length l are chosen to pertain to the
given geometric body shape
• E.g., for an airplane wing, S is the planform
area, and l is the mean chord length c.
• for a sphere, S is the cross-sectional area, and
l is the diameter
• Particular choice of reference area and length
is not critical
• But, when using force and moment coefficient
data, we must always know what reference
quantities the particular data are based upon.

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Parameters which influence Aerodynamics
• By intuition, the resultant aerodynamic force R should depend on
– Freestream velocity 𝑉∞ (faster air  more force and moment)
– Freestream density 𝜌∞ (denser air  more force and moment)
– Freestream viscosity 𝜇∞ (viscosity  shear stress)
– Size of the body (more reference area and length  more force and moment)
– Compressibility of air (density changes if flow speed is comparable to speed of sound 𝑎∞)
– Angle of attack 𝛼
Therefore,
𝐿 = 𝑔𝑙(𝜌∞, 𝑉∞, 𝑆, 𝑐, 𝜇∞, 𝑎∞, 𝛼)
𝐷 = 𝑔 𝑑(𝜌∞, 𝑉∞, 𝑆, 𝑐, 𝜇∞, 𝑎∞, 𝛼)
M= 𝑔 𝑚(𝜌∞, 𝑉∞, 𝑆, 𝑐, 𝜇∞, 𝑎∞, 𝛼)
for some nonlinear functions 𝑔𝑙, 𝑔 𝑑 and 𝑔 𝑚.
• This is not useful as this means a huge combination of parameters needs to be tested or
simulated to find the relationships necessary for designing aerodynamic vehicles and
products.

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Dimensional Analysis
• Fortunately, we can simplify the problem and considerably reduce our time and effort
by first employing the method of dimensional analysis.
• Dimensional analysis is based on the obvious fact that in an equation dealing with the
real physical world, each term must have the same dimensions.
• So it is equivalent to find relationships between dimensionless groups of parameters
rather than the parameters themselves.
• We can show that force and moment coefficients depend on the Reynold and Mach
numbers and flow angles only, i.e.
𝐶𝑙 = 𝑓𝑙(𝑀∞, 𝑅𝑒∞, 𝛼)
𝐶 𝑑 = 𝑓𝑑(𝑀∞, 𝑅𝑒∞, 𝛼)
𝐶 𝑚 = 𝑓𝑚(𝑀∞, 𝑅𝑒∞, 𝛼)
i.e. fl, fd and fm.
• Notice that all parameters are dimensionless!
• Notice that we have reduced the number of parameters from 7 to just 3!
• There may be other “similarity parameters” other than these 3, depending on problem.
𝑅𝑒∞ =
𝜌∞ 𝑉∞ 𝑐
𝜇∞
𝑀∞ =
𝑉∞
𝑎∞

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Suffa University
Flow Similarity
• Consider two different flow fields over two different bodies. By definition, different flows
are dynamically similar if:
1. Streamline patterns are geometrically similar.
2. Distributions of
𝑉
𝑉∞
,
𝑝
𝑝∞
,
𝑇
𝑇∞
etc., throughout the flow field are the same when plotted against
non-dimensional coordinates.
3. Force coefficients are the same.
• If nondimensional pressure (CP) and shear stress distributions (
𝜏
𝑄∞
) over different bodies
are the same, then the force and moment coefficients will be the same.
• In other words, two flows will be dynamically similar if:
1. The bodies and any other solid boundaries are geometrically similar for both
flows.
2. The similarity parameters are the same for both flows.
• This is a key point in the validity of wind-tunnel testing.

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Wait a Minute…Too much Math 
• Hey, this was supposed to be fun

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So how to solve these equations?
• Obviously, this is pretty darn hard mathematics!
• So how to solve these equalities and inequalities?
• No closed form analytic solution (except simplest cases)
• One way is the wind tunnel!
1/3 scale model of space shuttle in
NASA’s 40-foot-by-80-foot WT
Mercedes-Benz’s Aeroacoustic Wind Tunnel Educational Wind Tunnel

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Computational Fluid Dynamics (CFD)
• An alternate technique is to divide the flow into small
boxes (grid) and solve the full Navier Stokes (or
simplifications) equations at every point numerically.
• This is now possible with high speed computing.
• But it is not really as useful as thought.
• In the end, it is really Colorful Fluid Dynamics. A lot of
colors which may mean something…or nothing.
• Needs a lot of calibration and EXPERTISE.
• Also, not feasible for trade studies and initial design.
• Slow solutions. One design iteration can take days.

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Between Wind Tunnels and CFD
• So, wind tunnels are expensive and time
consuming.
• So is CFD. Both requires experts to interpret
results.
• You really can’t DESIGN your aircraft in WT/CFD.
• Here is where “engineering solutions” come in.
• Ultimately somewhere between an “educated
guess” and JUGAAR!

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Engineering Aerodynamic Softwares
• For preliminary aircraft design, we need quick, somewhat
crude solutions. Ball park estimates will do.
• The USAF saw that need, and decided to compile data on
aerodynamic predictions. Contract: McDonnel Douglas.
• There were some approximate analytic solutions.
• There were some correlations.
• There were half a century of wind/water tunnel tests
• All of this was compiled in two volumes called USAF Stability
and Control Data Compendium (or DATCOM!) in 1978

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What is DATCOM?
• DATCOM is a collection, correlation, codification, and recording of best knowledge,
opinion, and judgment in the area of aerodynamic stability and control prediction
methods.
• Used for
– Conceptual and Preliminary aircraft design
– Evaluate changes resulting from proposed engineering fixes
– For making simulators.
• For any given configuration and flight condition, a complete set of stability and control
derivatives can be determined without resort to outside information.
• Methods range from very simple and easily applied techniques to quite accurate and
thorough procedures.
• The book is intended to be used for preliminary design purposes before WT/CFD.
• It is not easy to sift through two volumes of 3000 pages though!

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Digital DATCOM
• USAF Stability and Control Digital DATCOM is a computer
program that implements the methods contained in
the USAF DATCOM.
• First version in 1978. Implemented in FORTRAN IV.
• It calculates the stability, control and dynamic derivative
characteristics of fixed-wing aircraft.
• It uses text based input and text based output.
• Some GUIs have recently come out, but are not very stable.
• The program was declassified around the year 2000.
http://www.pdas.com/datcomdownload.html

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What can DATCOM do?
• Datcom requires two basic inputs
– Flight Conditions (Altitude, Speed, Re, M, Flow angles)
– Aircraft Geometry (Wings, Tails, Fuselage dimensions etc)
– Optionally propulsion data (jet/propeller can also be input)
• It calculates:
– Lift, Drag, Moments, Center of Pressure, Flow angle derivatives
(stability derivatives)
– Output can be for whole aircraft or components
•CL
•CD
•Cm
•CN
•CA
•CLα
•Cmα
•CYβ
•Cnβ
•Clβ

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Let’s Begin with the Input File
• Remember we are talking about FORTRAN, so data must be entered in the correct column.
– FORTRAN uses Control Cards (objects) and Namelists (functions).
– Namelists start with $ sign. They star after one space.
– Inputs to namelists can be entered one space after a namelist title or beginning with the third space of a new line.
– Control Cards come alone on a line. The start in the first column.
• Easier to edit old files than making one from scratch.
• Begin by first naming the aircraft or “case” to be run.
• This is done by typing CASEID, spacing once, and typing the desired name.
• CASEID is a control card, and thus the “C” should be the first letter on the line and no characters
should follow the case name on the line.
• Next, the system of units for the input may be specified by DIM control card :
– DIM M – kilogram-meter-second
– DIM CM – centimeter-gram-second
– DIM FT – foot-pound-second
– DIM IN – inch-pound-second
• Angles are entered as degrees, always!

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• Next, flight conditions should be entered.
• This is done by calling out the FLTCON namelist. A namelist must be called out
on a new line by spacing once and entering the $ symbol followed by the
namelist title.
• The inputs under the FLTCON namelist includes
– MACH – Mach number
– VINF – Airspeed in units of length (as chosen by the DIM control card) per unit time
– NALPHA – number of angles of attack to be evaluated
– ALSCHD – angles of attack to be evaluated, written sequentially
– GAMMA – flight path angle
– ALT – altitudes to be evaluated
– WT – aircraft weight
• Alternatively, Reynold Number can be entered.

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What does it look like so far?
The “picture” so far

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Geometry….Fuselage
• Next, geometry of the aircraft must be entered.
• Fuselage is defined by the BODY namelist.
• It can be defined by a maximum of 20 longitudinal stations.
– NX – number of stations used to define the body
– X(1) – longitudinal location of station
– R(1) – planform half-width of the fuselage in the spanwise direction
– ZU(1) – location of upper vertical surface of fuselage with respect to an
arbitrary reference plane
– ZL(1) – location of lower vertical surface of fuselage with respect to an
arbitrary reference plane
• Alternatively, a cylindrical fore, mid and aft body is all that is needed

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Namelist BODY

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Geometry….Lifting Surfaces
• Wing, horizontal and vertical tails all have similar inputs.
• Use WGPLNF, VTPLNF, and HTPLNF namelists.
• All the necessary inputs to define a straight tapered planform are as follows:
– SSNPE – exposed semi-span
– SSPN – theoretical semi-span
– CHRDR – root chord
– CHRDTP – tip chord
– SAVSI – sweep angle
– DHDADI – dihedral angle
– TWISTA – twist angle
– CHSTAT – reference chord station for inboard for panel sweep angle
– TYPE – type of wing planform (1.0=straight tapered planform)
• It is also possible to make cranked wings (see manual)

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WGPLNF, VTPLNF, and HTPLNF

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Wing Sections (Airfoils)
• Airfoil can be entered using the NACA control card. (begin column 01)
• NACA 23012 airfoil for a vertical tail is given by:
– NACA-V-5-23012
• V specifies that the airfoil is for the vertical tail. A W, H, or F in the same
place would specify the wing, horizontal tail, or ventral fin airfoil
respectively.
• The 5 specifies the type of airfoil, in this case the 5-digit airfoils.
• Other options are 1, 4, 6, and S for 1-series, 4-digit, 6-series NACA
airfoils, and supersonic airfoils respectively.
• Last input is the airfoil designation.
• You can also enter your own or exotic airfoils (See manual)

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The “Picture” so far…

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Locating the Components on the Fuse
• SYNTHS namelist used to place wings, tail, center of gravity.
– XCG – longitudinal location of center of gravity
– ZCG – vertical location of center of gravity
– XW – longitudinal location of theoretical wing apex
– ZW – vertical location of theoretical wing apex
– XH – longitudinal location of theoretical horizontal tail apex
– ZH – vertical location of theoretical horizontal tail apex
– XV – longitudinal location of theoretical vertical tail apex
– ZV – vertical location of theoretical vertical tail apex
– ALIW – wing incidence angle
– ALIH – horizontal tail incidence angle
•

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Reference Areas and Lengths
• Finally, we may want to put our own reference areas and
lengths (can be left if you want DATCOM to calculate the
same)

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Flaps, Control Surfaces
• Flaps and elevators can be modeled using the SYMFLP and ailerons by ASYFLP
namelists, which output results for symmetrical and asymmetrical flap
deflections, respectively.
– FTYPE – type of flaps (SYMFLP only)
– NDELTA – number of flap deflections to be evaluated
– DELTA(1) – flap deflections listed sequentially (maximum of 9, SYMFLP only)
– CHRDFI – inboard flap chord length
– CHRDFO – outboard flap chord length
– SPANFI – spanwise location of flap inboard panel
– SPANFO – spanwise location of flap outboard panel
– STYPE – control surface type (1.0=flap spoiler, 2.0=plug spoiler, 3.0=spoiler-slot
deflection, 4.0=plain flap aileron, 5.0=all moveable tail, ASYFLP only)
– DELTAL(1) – left flap deflection angles listed sequentially (maximum of 9, ASYFLP only)
– DELTAR(1) – right flap deflection angles

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Control Surfaces

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Some other Options
• You may also want dynamic deratives (with pitch rate, angle
of attack rate etc)
• Use DAMP control card
• DERIV RAD and DERIV DEG output these derivatives in radian
and degree

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Complete File

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Processing the Input file
• Save the file as ***.inp
• Then run datcom.exe and process

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Output
• Let’s see the output in Class.
• You can import the ouput to Matlab using Aerospace Toolbox
• alldata = datcomimport('astdatcom.out', true, 0);
• Plotting Lift Curve Moments
h1 = figure;
for k=1:2
subplot(2,1,k)
plot(data.alpha,permute(data.cl(:,k,:),[1 3 2]))
end

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Plotting the Input Aircraft
• Bill Galbraith of HolyCows sells DATCOM+ for $100 doing this

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Intermission – Equations of Motion

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Equations of Motion
• So, these are 12 nonlinear ODEs.
• There is no analytical solution.
• We need to use numerical integration methods (Adams, RK..)
• Luckily, this is all programmed in Simulink graphical
programming language.

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Simulink Aerospace Toolbox
• Aerospace Blockset™ software extends Simulink® with blocks for
– modeling and simulating
• aircraft,
• spacecraft,
• rocket,
• propulsion systems,
• unmanned airborne vehicles.
– aerospace standards,
– modeling equations of motion
– navigation,
– gain scheduling,
– visualization,
– unit conversion

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So, let us implement the first step

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Next Set Some Aerodynamic Variables
• Use conversions for flow angles (𝛼, 𝛽)
• Use standard atmosphere and gravity models for (𝑔, 𝑀)

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Tidy Things Up a bit
• Create subsystems

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A detailed example
• Lightweight Airplane Simulator Design
– Four seater monoplane: Skyhogg
• We will use Simulink for rapid
– Aircraft design
– Modeling
– Simulation
– Control Design

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Aerodynamics
• The designed geometry is
modeled in Datcom
• Datcom provides
aerodynamic stability and
control derivatives and
coefficients at specified
flight conditions.
• Import in Matlab using
statdyn = datcomimport('SkyHogg.out');

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Let’s get this Aerodynamics in Simulink

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Next also put the Elevator Model

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Equations of Motion Once again

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Actuator and Sensor Models

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Environmental Models

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Trim and Linearize
• Trim and Linearization can be done to find the operating
point at specified speed and altitude, for example.
• You Analysis>Control System Design> Linear Analysis

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Linear System Analysis

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Control System Design
• Dual Loop Control
– Time scale separation

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Altitude Control

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Pizzaz: Integration with FlighGear

Workshop on Digital Datcom and Simulink

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