mae 331 lecture 12
DESCRIPTION
aviationsTRANSCRIPT
Linearized Lateral-Directional
Equations of MotionRobert Stengel, Aircraft Flight Dynamics MAE 331, 2010
• Spiral, Dutch roll, and roll modes
• Stability derivatives
Roll-mode responseof roll angle
Spiral-moderesponse ofcrossrange
Spiral-mode responseof yaw angle
Dutch-roll-mode responseof side velocity
Dutch-roll-moderesponse of rolland yaw rates
Copyright 2010 by Robert Stengel. All rights reserved. For educational use only.http://www.princeton.edu/~stengel/MAE331.html
http://www.princeton.edu/~stengel/FlightDynamics.html
6-Component
Lateral-DirectionalEquations of Motion
State Vector,6 components
Nonlinear Dynamic Equations
!v = Y / m + gsin! cos" # ru + pw
!yI= cos" sin$( )u + cos! cos$ + sin! sin" sin$( )v + # sin! cos$ + cos! sin" sin$( )w
!p = IzzL + I
xzN # I
xzIyy# I
xx# I
zz( ) p + Ixz
2+ I
zzIzz# I
yy( )%& '(r{ }q( ) ÷ IxxIzz# I
xz
2( )
!r = IxzL + I
xxN # I
xzIyy# I
xx# I
zz( )r + Ixz
2+ I
xxIxx# I
yy( )%& '( p{ }q( ) ÷ IxxIzz# I
xz
2( )!! = p + qsin! + r cos!( ) tan"
!$ = qsin! + r cos!( )sec"
x1
x2
x3
x4
x5
x6
!
"
########
$
%
&&&&&&&&
= xLD6=
v
y
p
r
'
(
!
"
########
$
%
&&&&&&&&
=
Side Velocity
Crossrange
Body ) Axis Roll Rate
Body ) Axis Yaw Rate
Roll Angle about Body x Axis
Yaw Angle about Inertial x Axis
!
"
########
$
%
&&&&&&&&
Douglas A-4
4- Component
Lateral-Directional
Equations of Motion
State Vector,
4 components
Nonlinear Dynamic Equations, neglecting crossrange and yaw angle
!v = Y / m + gsin! cos" # ru + pw
!p = IzzL + I
xzN # I
xzIyy# I
xx# I
zz( ) p + Ixz
2+ I
zzIzz# I
yy( )$% &'r{ }q( ) ÷ IxxIzz# I
xz
2( )
!r = IxzL + I
xxN # I
xzIyy# I
xx# I
zz( )r + Ixz
2+ I
xxIxx# I
yy( )$% &' p{ }q( ) ÷ IxxIzz# I
xz
2( )!! = p + qsin! + r cos!( ) tan"
x1
x2
x3
x4
!
"
#####
$
%
&&&&&
= xLD4=
v
p
r
'
!
"
#####
$
%
&&&&&
=
Side Velocity
Body ( Axis Roll Rate
Body ( Axis Yaw Rate
Roll Angle about Body x Axis
!
"
#####
$
%
&&&&&
Eurofighter Typhoon
Lateral-Directional Equations
of Motion Assuming Steady,Level Longitudinal Flight
Nonlinear Dynamic Equations, assuming steady, level, longitudinal flight
(constant longitudinal variables)
!v = YB/ m + gsin! cos"
N# ru
N+ pw
N
= YB/ m + gsin! cos$
N# ru
N+ pw
N
!p =IzzLB+ I
xzN
B( )IxxIzz# I
xz
2( )
!r =IxzLB+ I
xxN
B( )IxxIzz# I
xz
2( )!! = p + r cos!( ) tan"N
= p + r cos!( ) tan$N
qN= 0
!N= 0
"N= #
N
Lockheed F-117
Lateral-Directional Force
and Moments
YB = CYB
1
2!NVN
2S; Body " Axis Side Force
LB = ClB
1
2!NVN
2Sb; Body " Axis Rolling Moment
NB = CnB
1
2!NVN
2Sb; Body " Axis Yawing Moment
Linearized Equationsof Motion
Body-Axis Perturbation
Equations of Motion
!!x
LD(t) = F
LD!x
LD(t) +G
LD!u
LD(t) + L
LD!w
LD(t)
! !v(t)
!p(t)
!r(t)
!"(t)
#
$
%%%%%
&
'
(((((
=
) f1
)v) f
1
) p) f
1
)r) f
1
)"
) f2
)v) f
2
) p) f
2
)r) f
2
)"
) f3
)v) f
3
) p) f
3
)r) f
3
)"
) f4
)v) f
4
) p) f
4
)r) f
4
)"
#
$
%%%%%%%%%%%
&
'
(((((((((((
!v(t)
!p(t)
!r(t)
!"(t)
#
$
%%%%%
&
'
(((((
+
) f1
)*A) f
1
)*R) f
2
)*A) f
2
)*R) f
3
)*A) f
3
)*R) f
4
)*A) f
4
)*R
#
$
%%%%%%%%%%
&
'
((((((((((
!*A(t)
!*R(t)
#
$%%
&
'((+
) f1
)vwind
) f1
) pwind) f
2
)vwind
) f2
) pwind) f
3
)vwind
) f3
) pwind) f
4
)vwind
) f4
) pwind
#
$
%%%%%%%%%%%
&
'
(((((((((((
!vwind!pwind
#
$%%
&
'((
!u1
!u2
"
#
$$
%
&
''=
!(A
!(R
"
#$
%
&' =
Aileron Perturbation
Rudder Perturbation
"
#
$$
%
&
''
!w1
!w2
"
#
$$
%
&
''=
!(A
!(R
"
#$
%
&' =
SideWind Perturbation
Vortical Wind Perturbation
"
#
$$
%
&
''
• Perturbations from wings-level flight vN= 0; p
N= 0; r
N= 0; !
N= 0
!v!p
!r!"
#
$
%%%%%
&
'
(((((
=
Side Velocity Perturbation
Body ) Axis Roll Rate Perturbation
Body ) Axis Yaw Rate Perturbation
Roll Angle about Body x Axis Perturbation
#
$
%%%%%
&
'
(((((
Typical
Lateral-Directional Initial-Condition
Linearized Response
Roll-mode responseof roll angle
Spiral-moderesponse ofcrossrange
Spiral-moderesponse of yawangle
Dutch-roll-mode responseof side velocity
Dutch-roll-moderesponse of rolland yaw rates
Dimensional Stability-
and-Control Derivatives! f
1
!v! f
1
! p! f
1
!r! f
1
!"
! f2
!v! f
2
! p! f
2
!r! f
2
!"
! f3
!v! f
3
! p! f
3
!r! f
3
!"
! f4
!v! f
4
! p! f
4
!r! f
4
!"
#
$
%%%%%%%%%%%
&
'
(((((((((((
=
Yv Yp + wN( ) Yr ) uN( ) gcos*N
Lv Lp Lr 0
Nv Np Nr 0
0 1 tan*N 0
#
$
%%%%%%
&
'
((((((
! f1
!"A
! f1
!"R
! f2
!"A
! f2
!"R
! f3
!"A
! f3
!"R
! f4
!"A
! f4
!"R
#
$
%%%%%%%%%%
&
'
((((((((((
=
Y"A Y"R
L"A L"R
N"A N"R
0 0
#
$
%%%%%
&
'
(((((
! f1
!vwind
! f1
! pwind
! f2
!vwind
! f2
! pwind
! f3
!vwind
! f3
! pwind
! f4
!vwind
! f4
! pwind
"
#
$$$$$$$$$$$
%
&
'''''''''''
=
Yv Yp
Lv Lp
Nv Np
0 0
"
#
$$$$$
%
&
'''''
StabilityMatrix
ControlEffect Matrix
DisturbanceEffect Matrix
v
p
r
!
"
#
$
$
$
$
%
&
'
'
'
'
=
Side Velocity
Body ( Axis Roll Rate
Body ( Axis Yaw Rate
Roll Angle about Body x Axis
"
#
$
$
$
$
%
&
'
'
'
'
! !v(t)
!!p(t)
!!r(t)
! !"(t)
#
$
%%%%%
&
'
(((((
=
Yv Yp + wN( ) Yr ) uN( ) gcos*N
Lv Lp Lr 0
Nv Np Nr 0
0 1 tan*N 0
#
$
%%%%%%
&
'
((((((
!v(t)
!p(t)
!r(t)
!"(t)
#
$
%%%%%
&
'
(((((
+
Y+A Y+R
L+A L+R
N+A N+R
0 0
#
$
%%%%%
&
'
(((((
!+A(t)
!+R(t)
#
$%%
&
'((+
Yv Yp
Lv Lp
Nv Np
0 0
#
$
%%%%%
&
'
(((((
!vwind!pwind
#
$%%
&
'((
• Rolling and yawing motions
Body-Axis Perturbation
Equations of Motion
Stability Axes
Stability Axes• Alternative set of body axes
• Nominal x axis is offset from the body centerline bythe nominal angle of attack, !N
Transformation fromOriginal Body Axesto Stability Axes
HB
S=
cos!N
0 sin!N
0 1 0
" sin!N
cos!N
#
$
%%%
&
'
(((
!u
!v
!w
"
#
$$$
%
&
'''S
= HB
S
!u
!v
!w
"
#
$$$
%
&
'''B
!p
!q
!r
"
#
$$$
%
&
'''S
= HB
S
!p
!q
!r
"
#
$$$
%
&
'''B
• Side velocity (!v) and pitch rate (!q) are unchanged by the transformation
Stability-Axis Lateral-Directional Equations
• Rotate body axes to stability axes
• Replace side velocity by sideslip angle
• Revise state order
!" #!v
VN
!v(t)
!p(t)
!r(t)
!"(t)
#
$
%%%%%
&
'
(((((Body)Axis
*
!v(t)
!p(t)
!r(t)
!"(t)
#
$
%%%%%
&
'
(((((Stability)Axis
!v(t)
!p(t)
!r(t)
!"(t)
#
$
%%%%%
&
'
(((((Stability)Axis
*
!+(t)
!p(t)
!r(t)
!"(t)
#
$
%%%%%
&
'
(((((Stability)Axis
!"(t)
!p(t)
!r(t)
!#(t)
$
%
&&&&&
'
(
)))))Stability*Axis
+
!r(t)
!"(t)
!p(t)
!#(t)
$
%
&&&&&
'
(
)))))Stability*Axis
Stability-Axis Lateral-Directional Equations
!!r(t)
! !"(t)!!p(t)
! !#(t)
$
%
&&&&&
'
(
)))))S
=
Nr N" Np 0
Yr
VN*1
+
,-.
/0Y"
VN
Yp
VN
gcos1 N
VN
Lr L" Lp 0
tan1 N 0 1 0
$
%
&&&&&&&
'
(
)))))))S
!r(t)!"(t)!p(t)!#(t)
$
%
&&&&&
'
(
)))))S
+
N2A N2R
Y2A
VN
Y2R
VN
L2A L2R
0 0
$
%
&&&&&&
'
(
))))))S
!2A(t)!2R(t)
$
%&&
'
())+
N" Np
Y"
VN
Yp
VN
L" Lp
0 0
$
%
&&&&&&&
'
(
)))))))S
!"wind
!pwind
$
%&&
'
())
!x1
!x2
!x3
!x4
"
#
$$$$$
%
&
'''''
=
!r!(
!p
!)
"
#
$$$$$
%
&
'''''
=
Stability * Axis Yaw Rate Perturbation
Sideslip Angle Perturbation
Stability * Axis Roll Rate Perturbation
Stability * Axis Roll Angle Perturbation
"
#
$$$$$
%
&
'''''
Why Modify the Equations?
• Dutch-roll mode is primarily described by stability-axis yawrate and sideslip angle
• Roll and spiral mode are primarily described by stability-axisroll rate and roll angle
• Linearized equations allow the three modes to be studied
Stable Spiral
Unstable Spiral
Roll
Dutch Roll, top Dutch Roll, front
Why Modifythe Equations?
FLD
=FDR
FRS
DR
FDR
RSFRS
!
"
##
$
%
&&=
FDR
small
small FRS
!
"
##
$
%
&&'
FDR
~ 0
~ 0 FRS
!
"
##
$
%
&&
Effects of Dutch roll perturbations
on Dutch roll motion
Effects of Dutch roll perturbations
on roll-spiral motion
Effects of roll-spiral perturbationson Dutch roll motion
Effects of roll-spiral perturbations on
roll-spiral motion
... but are the off-diagonal blocks really small?
Dassault Rafale
Stability, Control, andDisturbance Matrices
FLD =FDR FRS
DR
FDRRS
FRS
!
"
##
$
%
&&=
Nr N' Np 0
Yr
VN(1
)
*+,
-.Y'
VN
Yp
VN
gcos/ N
VN
Lr L' Lp 0
tan/ N 0 1 0
!
"
#######
$
%
&&&&&&&
GLD
=
N!A N!R
Y!A
VN
Y!R
VN
L!A L!R
0 0
"
#
$$$$$$
%
&
''''''
LLD
=
N! Np
Y!
VN
Yp
VN
L! Lp
0 0
"
#
$$$$$$$
%
&
'''''''
!x1
!x2
!x3
!x4
"
#
$$$$$
%
&
'''''
=
!r!(
!p
!)
"
#
$$$$$
%
&
'''''
!u1
!u2
"
#
$$
%
&
''=
!(A
!(R
"
#$
%
&'
!w1
!w2
"
#
$$
%
&
''=
!(A
!(R
"
#$
%
&'
Lateral-DirectionalStability Derivatives
“Weathercock Stability”: Yaw AccelerationSensitivity to Sideslip Angle, N"
N! "#C
n
#!$V 2
2
%&'
()*Sb
Izz
= Cn!
$V 2
2
%&'
()*Sb
Izz
• Vertical tail, fuselage, and wing areprincipal contributors to directionalstability– Side force contributions times
respective moment arms
– Non-dimensional stability derivative
> 0 for stability
• Dimensional stability derivative
Cn!" Cn!( )
Vertical Tail+ Cn!( )
Fuselage+ Cn!( )
Wing+ Cn!( )
Propeller
Contributions toDirectional Stability, N"
Cn!( )Fuselage
="2K VolumeFuselage
SbK = 1! dmax
Lengthfuselage
"#$
%&'1.3
Cn!( )
Vertical Tail
" aFin#vt
SFinlvt
Sb= a
Fin#vtVvt
• Vertical tail
Vvt=SFinlvt
Sb= Vertical Tail Volume Ratio
• Fuselage
!vt= !
elas1+ "# "$( ) V
Tail
2
VN
2
%
&'(
)*
aFin = (–) Lift slope of isolated fin
referenced to fin area, SFin
VTail = Airspeed at vertical tail;“scrubbing” lowers VTail,propeller slipstreamincreases VTail
#, Sidewash: Deflection of flowfrom freestream direction dueto wing-fuselage interference
Contributions toDirectional Stability, N"
Cn!( )Wing
= 0.75CLN" + fcn #,AR,$( )CLN
2
Cn!( )Propeller• Propeller fin effect
– Function of blade geometry,number of blades, thrustcoefficient, and advance ratio
– J = V/nd
– n = rpm
– d = propeller diameter
" = Sweep angle
AR = Aspect ratio
$ = Taper ratio
% = Dihedral angle
• Wing– Differential drag
– Interference effects
see NACA-TR-819
Side Acceleration Sensitivityto Sideslip Angle, Y"
• Dimensional stability derivative
Y! " CY!
#V 2
2m
$%&
'()S
CY!" CY!( )
Fuselage+ CY!( )
Vertical Tail+ CY!( )
Wing
CY!( )Vertical Tail
" aFin#vt
SFinS
CY!( )Fuselage
" $2SBase
S; SB =
%dBase2
4
CY!( )Wing
= $CDParasite,Wing$ k& 2
k =!AR
1+ 1+ AR2
• Fuselage, vertical tail, and wing are principal contributors
Yaw Damping Due to Yaw Rate, Nr
• Dimensional stability derivative
Nr! C
nr
"VN
2
2Izz
#
$
%
&
'
( Sb = Cn
ˆ r
b
2VN
#
$ %
&
' ( "V
N
2
2Izz
#
$
%
&
'
( Sb
= Cn
ˆ r
"VN
4Izz
#
$
%
&
'
( Sb2 < 0 for stability
• High wing-sweep anglecan lead to Nr > 0
Martin Marietta X-24B
Yaw Damping Due toYaw Rate, Nr
Cnr̂! Cnr̂( )
Vertical Tail+ Cnr̂( )
Wing
Cnr̂( )
Wing= k
0CL
2+ k
1CDParasite,Wing
k0 and k1 are functions of
aspect ratio and sweep angle
• Wing contribution
• Vertical tail contribution
! Cn( )
Vertical Tail= " C
n#( )Vertical Tail
rlvt
V( ) = " Cn#( )
Vertical Tail
lvt
b
$%&
'()
b
V
$%&
'()r
Cnr̂
( )vt
=*! C
n( )Vertical Tail
* rb2V( )
=*! C
n( )Vertical Tail
* r̂= "2 C
n#( )Vertical Tail
lvt
b
$%&
'()
r̂ =rb
2VN
Approximate Rolland Spiral Modes
!!p
! !"
#
$%%
&
'((=
Lp 0
1 0
#
$%%
&
'((
!p
!"
#
$%%
&
'((+
L)A
0
#
$%%
&
'((!)A
!RS (s) = s s " Lp( )#S = 0
#R = Lp
• Characteristic polynomial has real roots
• Roll rate is damped by Lp
• Roll angle is a pure integral of roll rate
Roll Damping Due to Roll Rate, Lp
Lp ! Clp
"VN2
2Ixx
#
$%&
'(Sb = Clp̂
b
2VN
#
$%&
'("VN
2
2Ixx
#
$%&
'(Sb
= Clp̂
"VN4Ixx
#
$%&
'(Sb
2
Clp̂( )Wing
=! "Cl( )
Wing
! p̂= #
CL$
12
1+ 3%1+ %
&'(
)*+
• Wing with taper
• Thin triangular wingClp̂( )
Wing= !
"AR
32
Clp̂! Clp̂( )
Vertical Tail+ Clp̂( )
Horizontal Tail+ Clp̂( )
Wing
• Vertical tail, horizontal tail, and wing areprincipal contributors
< 0 for stability
Roll Damping Dueto Roll Rate, Lp
• Vertical tail
• Tapered horizontal tail p̂ =pb
2VN
Clp̂( )ht=! "Cl( )
ht
! p̂= #
CL$ht
12
Sht
S
%&'
()*1+ 3+1+ +
%&'
()*
• pb/2VN describes helix anglefor a steady roll
Clp̂( )vt=! "Cl( )
vt
! p̂= #
CY$vt
12
Svt
S
%&'
()*1+ 3+1+ +
%&'
()*
Primary Lateral-DirectionalControl Derivatives
N!R = Cn!R
"VN
2
2Izz
#
$%&
'(Sb
L!A = Cl!A
"VN
2
2Ixx
#
$%&
'(Sb
Comparison of Fourth-and Second-OrderDynamic Models
Second-Order Models of
Lateral-Directional Motion
• Approximate Spiral-Roll Equation
• Approximate Dutch Roll Equation
!!xDR
=!!r! !"
#
$%%
&
'(()
Nr
N"
Yr
VN
*1+
,-.
/0Y"
VN
#
$
%%%%
&
'
((((
!r!"
#
$%%
&
'((+
N1R
Y1R
VN
#
$
%%%
&
'
(((!1R +
N"
Y"
VN
#
$
%%%%
&
'
((((
!"wind
!!xRS =!!p
! !"
#
$%%
&
'(()
Lp 0
1 0
#
$%%
&
'((
!p
!"
#
$%%
&
'((+
L*A
0
#
$%%
&
'((!*A +
Lp
0
#
$%%
&
'((!pwind
• 2nd-order-model eigenvalues are close to those of the 4th-order model
• Eigenvalue magnitudes of Dutch roll and roll roots are similar
Bizjet Fourth- and
Second-Order Modelsand Eigenvalues
Fourth-Order ModelF = G = Eigenvalue Damping Freq. (rad/s)
-0.1079 1.9011 0.0566 0 0 -1.1196 0.00883
-1 -0.1567 0 0.0958 0 0 -1.20.2501 -2.408 -1.1616 0 2.3106 0 -1.16e-01 + 1.39e+00j 8.32E-02 1.39E+00
0 0 1 0 0 0 -1.16e-01 - 1.39e+00j 8.32E-02 1.39E+00
Dutch Roll ApproximationF = G = Eigenvalue Damping Freq. (rad/s)
-0.1079 1.9011 -1.1196 -1.32e-01 + 1.38e+00j 9.55E-02 1.38E+00
-1 -0.1567 0 -1.32e-01 - 1.38e+00j 9.55E-02 1.38E+00
Roll-Spiral ApproximationF = G = Eigenvalue Damping Freq. (rad/s)
-1.1616 0 2.3106 0
1 0 0 -1.16
Comparison of Second- and Fourth-OrderInitial-Condition Responses of Business Jet
Fourth-Order Response Second-Order Response
• Speed and damping of responses is adequately portrayed by 2nd-order models
• Roll-spiral modes have little effect on yaw rate and sideslip angle responses
• Dutch roll mode has large effect on roll rate and roll angle responses
Approximate DutchRoll Mode
!!r! !"
#
$%%
&
'((=
Nr
N"
Yr
VN
)1*
+,-
./Y"
VN
#
$
%%%%
&
'
((((
!r!"
#
$%%
&
'((+
N0R
Y0R
VN
#
$
%%%
&
'
(((!0R
!DR(s) = s
2 " Nr+Y#VN
$%&
'()s + N# 1"
Yr
VN
( ) + Nr
Y#VN
*+,
-./
0nDR
= N# 1"Yr
VN
( ) + Nr
Y#VN
1DR
= " Nr+Y#VN
$%&
'()
2 N# 1"Yr
VN
( ) + Nr
Y#VN
!nDR
= N" + Nr
Y"VN
#DR
= $ Nr+Y"VN
%&'
()*
2 N" + Nr
Y"VN
• With negligible Yr
• Characteristicpolynomial,natural frequency,and damping ratio
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Response