# F-16 Subsonic Aerodynamics Model (a la Garza)

### F-16 Aero Data file. Based on Morelli's adaptation of Stevens and Lewis' F-16 example [1] described in Garza & Morelli's TM [2]. Obtained from E. A. Morelli in the form of Matlab scripts [3] & [4]

Author: Bruce Jackson

Created: 28-MAR-2002

Reference 1. Stevens, Brian L. and Lewis, Frank L.: Aircraft Control and Simulation. ISBN O-471-61397-5, 1992.

Reference 2. Garza, Fredrico R.; and Morelli, Eugene A.: A Collection of Nonlinear Aircraft Simulations in MATLAB. NASA TM-2003-212145, JAN-2003.

Reference 3. Morelli, Eugene A.: f16_aero.m. 17-JUN-1995.

Reference 4. Morelli, Eugene A.: f16_aero_setup.m. 17-JUN-1995.

# Signal Definitions

SymbolNameUnitsSignInitial valueDepends on
Absolute value of angle-of-sideslip, deg.
absCl0
Absolute value of rolling moment coefficient
absCn0
Absolute value of yawing moment coefficient
Angle_of_Attack_degαdeg
Instantaneous true angle-of-attack, in degrees
Angle_of_Sideslip_degβdeg +wind in right ear
Instantaneous true angle-of-sideslip, in degrees
b2vbspan
tvt
Wing span over 2 times velocity. Used for coefficient normalization.
bspanft`30.`
Length of aerodynamic span, ft
cbarft`11.32`
Length of aerodynamic chord, ft
Cl0 +right wing downabsCl0
absCl0
beta
Basic coefficient of moment around the X-body direction (roll) (+RWD)
Cl1clt
dail
dclda
dcldr
drdr
Initial value of rolling moment, less damping terms
ClDAper deg +right/+left
Control power derivative: rolling moment per deg of aileron deflection
ClDRper deg +right/+left
Control power derivative: rolling moment per deg of rudder deflection
CLL +RWDb2v
cl1
clp
clr
p
r
Total coefficient of moment around the body X-axis (rolling moment)
CLM +ANUcmq
cmt
cq2v
cz
xcg
xcgr
Total coefficient of moment around the body Y-axis (pitching moment)
CLN +ANRb2v
bspan
cbar
cn1
cnp
cnr
cy
p
r
xcg
xcgr
Total coefficient of moment around the body Z-axis (yawing moment)
Damping derivative: rolling moment increment per radian/sec of roll rate
Damping derivative: rolling moment increment per radian/sec of yaw rate
Cm0 +nose up
Basic coefficient of moment around the Y-body direction (pitch) (+ANU)
Damping derivative: pitching moment increment per radian/sec of pitch rate
Cncnt
dail
dcnda
dcndr
drdr
Initial value of rolling moment coefficient without damping terms
Cn0 +nose rightabsCn0
absCn0
beta
Basic coefficient of moment around the Z-body direction (yaw) (+RWD)
CnDAper deg +right/+left
Control power derivative: yawing moment per deg of aileron deflection
CnDRper deg +right/+left
Control power derivative: yawing moment per deg of rudder deflection
Damping derivative: yawing moment increment per radian/sec of roll rate
Damping derivative: yawing moment increment per radian/sec of yaw rate
cq2vcbar
q
tvt
Chord times pitch rate over 2 times velocity. Used to normalize aero coefficients.
CX +FWDcq2v
cxq
cxt
Total coefficient of force along the body X-axis.
CX0 +forward
Basic coefficient of force in the X-body direction (+FWD). Function table based on angle of attack and elevator deflection.
Damping derivative: axial force increment per radian/sec of pitch rate
CY +RIGHTb2v
cy0
cyp
cyr
p
r
Total coefficient of force along the body Y-axis
CY0 +rightbeta
dail
drdr
Basic coefficient of force in the Y-body direction (+right). Function of sideslip angle, aileron and rudder deflection.
Damping derivative: lateral force increment per radian/sec of roll rate
Damping derivative: lateral force increment per radian/sec of yaw rate
CZ +DOWNcq2v
cz1
czq
Total coefficient of force along the body Z-axis.
CZ0 +down
Basic coefficient of force in the Z-body direction (+DWN). Function table based on angle of attack and elevator deflection.
CZ1 +downbeta
czt
del
rtd
Coefficient of force in the Z-body direction, modified from basic CZ0 to include effects of sideslip and elevator deflection.
Damping derivative: vertical force increment per radian/sec of pitch rate
dailail
Normalized aileron deflection.
delel
Normalized elevator deflection (note range is beyond breakpoint range)
Aileron deflection angle, in degrees (+LWD)
delta elevatorδedeg +trailing edge down
Elevator deflection angle, in degrees (+ED)
delta rudderδrdeg +TEL
Rudder deflection angle, in degrees (+TEL)
drdrrdr
Normalized rudder deflection.
Conversion constant from radians to degrees
True_Airspeed_f_p_sft/sec
True airspeed, ft/sec
tvtft/secvt
Twice velocity.
Xcgfract +aft
CG location relative to wing leading edge, expressed as a fraction of aerodynamic chord length
xcgrfract`0.35`
Default location of center of gravity relative to wing leading edge, expressed as a fraction of aerodynamic chord length.

# Function Descriptions

 Basic Cl absCl0 = f(absbeta, α) where: 0.0 <= absbeta <= 30.0; extrapolate neither -10.0 <= α <= 45.0; extrapolate neither Basic coefficient of rolling moment as a function of angle of attack and sideslip angle
 Basic Cm Cm0 = f(δe, α) where: -24.0 <= δe <= 24.0; extrapolate neither -10.0 <= α <= 45.0; extrapolate neither Basic coefficient of pitching-moment as a function of angle of attack and elevator
 Basic Cn absCn0 = f(absbeta, α) where: 0.0 <= absbeta <= 30.0; extrapolate neither -10.0 <= α <= 45.0; extrapolate neither Basic coefficient of yawing moment as a function of angle of attack and sideslip angle
 Basic CX CX0 = f(δe, α) where: -24.0 <= δe <= 24.0; extrapolate neither -10.0 <= α <= 45.0; extrapolate neither Basic coefficient of X-force table as a function of angle of attack and elevator
 Basic CZ CZ0 = f(α) where: -10.0 <= α <= 45.0; extrapolate neither Basic coefficient of Z-force table as a function of angle of attack
 Clp Clp = f(α) where: -10.0 <= α <= 45.0; extrapolate neither Damping derivative: rolling moment due to roll rate as a function of angle-of-attack
 Clr Clr = f(α) where: -10.0 <= α <= 45.0; extrapolate neither Damping derivative: rolling moment due to yaw rate as a function of angle-of-attack
 Cmq Cmq = f(α) where: -10.0 <= α <= 45.0; extrapolate neither Damping derivative: pitching moment due to pitch rate as a function of angle-of-attack
 Cnp Cnp = f(α) where: -10.0 <= α <= 45.0; extrapolate neither Damping derivative: yawing moment due to roll rate as a function of angle-of-attack
 Cnr Cnr = f(α) where: -10.0 <= α <= 45.0; extrapolate neither Damping derivative: yawing moment due to yaw rate as a function of angle-of-attack
 CXq CXq = f(α) where: -10.0 <= α <= 45.0; extrapolate neither Damping derivative: axial force due to pitch rate as a function of angle-of-attack
 CYp CYp = f(α) where: -10.0 <= α <= 45.0; extrapolate neither Damping derivative: lateral force due to roll rate as a function of angle-of-attack
 CYr CYr = f(α) where: -10.0 <= α <= 45.0; extrapolate neither Damping derivative: lateral force due to yaw rate as a function of angle-of-attack
 CZq CZq = f(α) where: -10.0 <= α <= 45.0; extrapolate neither Damping derivative: normal force due to pitch rate as a function of angle-of-attack
 dlda ClDA = f(β, α) where: -30.0 <= β <= 30.0; extrapolate neither -10.0 <= α <= 45.0; extrapolate neither Rolling moment increment due to aileron deflection as a function of angles of attack and sideslip
 dldr ClDR = f(β, α) where: -30.0 <= β <= 30.0; extrapolate neither -10.0 <= α <= 45.0; extrapolate neither Rolling moment increment due to rudder deflection as a function of angles of attack and sideslip
 dnda CnDA = f(β, α) where: -30.0 <= β <= 30.0; extrapolate neither -10.0 <= α <= 45.0; extrapolate neither Yawing moment increment due to aileron deflection as a function of angles of attack and sideslip
 dndr CnDR = f(β, α) where: -30.0 <= β <= 30.0; extrapolate neither -10.0 <= α <= 45.0; extrapolate neither Yawing moment increment due to rudder deflection as a function of angles of attack and sideslip

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