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Aircraft Control Devicesand Systems
Robert Stengel, Aircraft Flight Dynamics, MAE 331,
2010
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
Control surfaces
Control mechanisms
Flight control systems
Design for Control
Elevator/stabilator: pitch control
Rudder: yaw control
Ailerons: roll control
Trailing-edge flaps: low-angle lift control
Leading-edge flaps/slats: High-angle
lift control
Spoilers: Roll, lift, and drag control
Thrust: speed/altitude control
Critical Issues for Control Effect of control surface deflections on aircraft motions
Generation of control forces and rigid-body moments on the aircraft
Rigid-body dynamics of the aircraft
!E is an input
Command and control of the control surfaces
Displacements, forces, and hinge moments of the control mechanisms
Dynamics of control linkages
!E is a state
!!! ="
!!!E ="
Control Surface Dynamicsand Aerodynamics
Aerodynamic and
Mechanical Moments
on Control Surfaces Increasing size and speed of aircraft leads to
increased hinge moments
This leads to need for mechanical or aerodynamic
reduction of hinge moments
Need for aerodynamically balanced surfaces
Control surface hinge moment
Hsurface = CHsurface1
2!V 2Sc
CHsurface = CH !!
!! + CH! ! + CH"" + CHcommand + ...
CH !!E
: aerodynamic/mechanical damping moment
CH!E
: aerodynamic/mechanical spring moment
CH"
: floating tendency
CHcommand
: pilot or autopilot input
Hinge-moment coefficient, CH Linear model of dynamic effects
Angle of Attack andControl Surface Deflection
Horizontal tail at positiveangle of attack
Horizontal tail withelevator control surface
Horizontal tail with positiveelevator deflection
Floating and RestoringMoments on a Control Surface
Positive elevator deflection typically produces a negative (restoring) moment, H!" , on
the elevator due to aerodynamic or mechanical spring
Positive angle of attack typically produces a negative moment on the elevator
With stick free, i.e., no opposing torques, elevator floats up due to negative H#
Dynamic Model of a ControlSurface Mechanism
!!!E =H
elevator
Ielevator
=
CHelevator
1
2"V 2Sc
Ielevator
= CH !!E
!!E + CH!E
!E + CH## + C
Hcommand+ ...$%
&'
1
2"V 2Sc
Ielevator
( H!!E
!!E + H!E!E + H## + Hcommand + ...
Ielevator
= effective inertia of surface, linkages, etc.
H!!E=" H
elevatorIelevator( )
" !!; H
!E=" H
elevatorIelevator( )
"!
H#=" H
elevatorIelevator( )
"#
!!! " H!!
!! " H!! = H
## + H
command+ ...
mechanism dynamics = external forcing
Approximate control dynamics
by a 2nd-order LTI system
Bring all torques and
inertias to right side
Stability and control
derivatives of the controlmechanism
Coupling of System
Model and Control
Mechanism Dynamics
Second-order model of control-deflection dynamics
Short period approximation
!!xSP = FSP!xSP +GSP!uSP = FSP!xSP + F"ESP!x"E
! !q
! !#
$
%&&
'
())*
Mq M#
1 +L#VN
$
%
&&&
'
(
)))
!q
!#
$
%&&
'
())+
M"E 0
+L"E
VN0
$
%
&&&
'
(
)))
!"E
! !"E
$
%&&
'
())
!!x"E = F"E!x"E +G"E!u"E + FSP"E!xSP
! !"E
! !!"E
#
$%%
&
'(()
0 1
H"E H !"E
#
$%%
&
'((
!"E
! !"E
#
$%%
&
'((+
0
*H"E
#
$%%
&
'((!"Ecommand +
0 0
Hq H+
#
$%%
&
'((
!q
!+
#
$%%
&
'((
Augmented Short Period Model
Augmented dynamic equation
Augmented stability and control matrices
FSP /!E =FSP F!E
SP
FSP!E
F!E
"
#
$$
%
&
''=
Mq M( M!E 0
1 )L(VN
)L!E
VN0
0 0 0 1
Hq H( H!E H !!E
"
#
$$$$$$
%
&
''''''
!xSP ' =
!q
!"
!#E
! !#E
$
%
&&&&&
'
(
)))))
!!x
SP /"E= F
SP /"E!x
SP /"E+G
SP /"E!"E
command
State Vector
GSP /!E =
0
0
0
H!E
"
#
$$$$
%
&
''''
Roots of the AugmentedShort Period Model
Characteristic equation
!SP /"E s( ) = sIn # FSP /"E =
s # Mq( ) #M$ #M"E 0
#1 s +L$
VN( )L"E
VN0
0 0 s #1
#Hq #H$ #H"E s # H !"E( )
= 0
!SP /"E s( ) = s
2+ 2#
SP$
nSPs +$
nSP
2( ) s2 + 2#"E$n"E s +$n"E2( )
Coupling of the modes
depends on values of
M!E
,L!E
VN
,Hq, and H
"
Root locus evaluation of design parameters
Assuming high aero/mechanicaldamping of the control mechanism
Horn Balance
Inertial and aerodynamic effects
Control surface in front of hinge line
Increasing improves pitch stability,
to a point
Too much horn area
Degrades restoring moment
Increases possibility of mechanicalinstability
Increases possibility of destabilizing
coupling to short-period mode
CH ! CH"" + CH#E#E + CHpilot input
Stick-free case Control surface free to float
CH! C
H"" + C
H#E#E
Normally
CH!
< 0 : reduces short-period stability
CH"E
< 0 : required for mechanical stability
CH!
NACA TR-927, 1948
Overhang orLeading-EdgeBalance
Effect is similar tothat of horn balance
Varying gap andprotrusion intoairstream withdeflection angle
CH ! CH"" + CH# # + CHpilot input
NACA TR-927, 1948
Control Flap Carryover Effect on
Lift Produced By Total Surface
from Schlichting & Truckenbrodt
CL!E
CL"
vs.cf
x f + cf
c f x f + c f( )
All-Moving Control Surfaces
Particularly effective at supersonic speed (BoeingBomarc wing tips, North American X-15 horizontal
and vertical tails, Grumman F-14 horizontal tail)
SB.4!s aero-isoclinic wing
Sometimes used for trim only (e.g., Lockheed L-1011horizontal tail)
Hinge moment variations with flight condition
Shorts SB.4
Boeing
Bomarc
North American X-15
Grumman F-14
Lockheed L-1011
Control Surface Types
Elevator Effectiveness of high
mounting is unaffected by
wing downwash at low to
moderate angle of attack
Effectiveness of low mounting is
unaffected by wing downwash at
high angle of attack
Ailerons When one aileron goes up, the other goes down
Average hinge moment affects stick force
Compensating Ailerons Frise aileron
Asymmetric contour, with hinge line at or below loweraerodynamic surface
Reduces hinge moment
Cross-coupling effects can be adverse or favorable, e.g.
yaw rate with roll
Up travel of one > down travel of other to control yaw
effect
Abzug & Larrabee, 2002
Spoilers Spoiler reduces lift, increases drag
Speed control
Differential spoilers
Roll control
Avoid twist produced by outboardailerons on long, slender wings
free trailing edge for larger high-liftflaps
Plug-slot spoiler on P-61 BlackWidow: low control force
Hinged flap has high hinge moment
North American P-61
Abzug & Larrabee, 2002
Elevons Combined pitch and yaw control using
symmetric and asymmetric surfacedeflection
Principally used on Delta-wing configurations
Swing-wing aircraft
Grumman F-14
General Dynamics F-106Canards
Pitch control
Ahead of wing downwash
High angle of attack effectiveness
Desirable flying qualities effect(TBD)
Dassault Rafale
SAAB Gripen
Yaw Control of Tailless
Configurations Typically unstable in pitch and yaw
Dependent on flight control systemfor stability
Split ailerons or differential drag flapsproduce yawing moment
McDonnell Douglas X-36
Northrop Grumman B-2
Rudder Rudder provides yaw control
Turn coordination
Countering adverse yaw
Crosswind correction
Countering yaw due to engine loss
Strong rolling effect, particularly at high #
Only control surface whose nominalaerodynamic angle is zero
Possible nonlinear effect at low deflectionangle
Insensitivity at high supersonic speed Wedge shape, all-moving surface on North
American X-15
Martin B-57
Bell X-2
Control Tabs Balancing or geared tabs
Tab is linked to the main surface inopposition to control motion, reducingthe hinge moment with little change incontrol effect
Flying tabs
Pilot's controls affect only the tab,whose hinge moment moves thecontrol surface [BAC 1-11 deep stallflight test accident]
Linked tabs
divide pilot's input between tab andmain surface
Spring tabs
put a spring in the link to the mainsurface
BAC 1-11
Control MechanizationEffects
Control Mechanization Effects
Fabric-covered controlsurfaces (e.g., DC-3, Spitfire)subject to distortion under airloads, changing stability andcontrol characteristics
Control cable stretching
Elasticity of the airframechanges cable/pushrodgeometry
Nonlinear control effects friction
breakout forces
backlash
Douglas DC-3
Supermarine
Spitfire
Nonlinear Control Mechanism Effects
Friction
Deadzone
Control Mechanization Effects
Breakout force
Force threshold
Instabilities Due ToControl Mechanization
Aileron buzz (aero-mechanical instability; P-80)
Rudder snaking (Dutch roll/mechanical coupling; Meteor, He-162, X-1)
Aeroelastic coupling (B-47, Boeing 707 yaw dampers)
Rudder Snaking Control-free dynamics
Nominally symmetric control position
Internal friction
Aerodynamic imbalance
Coupling of mechanical motion withDutch roll mode (see Flight Dynamics)
Solutions Trailing-edge bevel
Flat-sided surfaces
Fully powered controls
Douglas DC-2
Roll/Spiral Limit Cycle
Due to Aileron
Imbalance
Control Surface
Buzz
North American FJ-4
At transonic speed,normal shocks may
occur on control surface
With deflection, shocks
move differentially
Possibility of self-
sustained nonlinear
oscillation (limit cycle)
ARC R&M 3364
Solutions
Splitter-plate rudder
fixes shock location for
small deflections
Blunt trailing edge
Fully powered controls
with actuators at the
surfaces
Rudder Lock Rudder deflected to stops at high
sideslip; aircraft trims at high $
3 necessary ingredients
Low directional stability at highsideslip due to stalling of fin
High (positive) hinge moment-due-to-sideslip at high sideslip(e.g., B-26)
Negative rudder yawing moment
Problematical if rudder isunpowered and requires highfoot-pedal force (rudder floatof large WWII aircraft)
Solutions
Increase high-sideslip directionalstability by adding a dorsal fin(e.g., B-737-100 (before), B-737-400 (after))
Hydraulically powered rudder
Martin B-26
Boeing 737-100
Boeing 737-400
Control Systems
Downsprings and Bobweights Adjustment of trim and stick-
force sensitivity to airspeed*
Downspring
Long mechanical spring withlow spring constant
Exerts a trailing-edge down
moment on the elevator
Bobweight
Weight on control column thataffects feel or basic stability
Mechanical stabilityaugmentation
Beechcraft B-18
* See pp. 541-545, Section 5.5, Flight Dynamics
Mechanical and Augmented
Control Systems
Mechanical system Push rods, bellcranks, cables, pulleys
Power boost Pilot's input augmented by hydraulic servo that
lowers manual force
Fully powered (irreversible) system No direct mechanical path from pilot to
controls
Mechanical linkages from cockpit controls toservo actuators
Mechanical, Power-Boosted Systems
Grumman A-6
McDonnell Douglas F-15
Boeing 767 Elevator Control System
Abzug & Larrabee, 2002
Classical Lateral Control Logic for
a Fighter Aircraft (c.1970)
MIL-DTL-9490E, Flight Control Systems - Design, Installation and Test
of Piloted Aircraft, General Specification for, 22 April 2008
Superseded for new designs on same date by
SAE-AS94900http://www.sae.org/servlets/works/documentHome.do?comtID=TEAA6A3&docID=AS94900&inputPage=dOcDeTaIlS
Effect of Scalar Feedback Control
on Roots of the System
!y(s) = H (s)!u(s) =kn(s)
d(s)!u(s) =
kn(s)
d(s)K!"(s)
Block diagram algebra
Aircraft Transfer Function (open-loop): H (s) =kn(s)
d(s)
Control System Gain :K
= Kkn(s)
d(s)!yc (s) " !y(s)[ ] = K
kn(s)
d(s)!yc (s) " K
kn(s)
d(s)!y(s)
Closed-Loop Transfer Function
!y(s) = Kkn(s)
d(s)!yc (s) " K
kn(s)
d(s)!y(s)
!y(s) + Kkn(s)
d(s)!y(s) = K
kn(s)
d(s)!yc (s)
!y(s)
!yc (s)=
Kkn(s)
d(s)
1+ Kkn(s)
d(s)
"
#$
%
&'
Roots of the Closed-Loop System
Closed-loop roots are solutions to
!closedloop
(s) = d(s) + Kkn(s) = 0
or Kkn(s)
d(s)= !1
!y(s)
!yc (s)=
Kkn(s)
d(s)
1+ Kkn(s)
d(s)
"
#$
%
&'
=Kkn(s)
d(s) + Kkn(s)[ ]=
Kkn(s)
!closedloop
s( )
Root Locus Analysis of Pitch Rate Feedback toElevator (2nd-Order Approximation)
Kkn(s)
d(s)= K
!q(s)
!"E(s)= K
kq s # zq( )s2+ 2$SP%nSP s +%nSP
2= #1
Number of roots = 2
Number of zeros = 1
Destinations of roots (for k = "):
1 root goes to the zero of n(s)
1 root goes to infinite radius from the origin
Angles of asymptotes, %, for theroots going to "
K -> +": 180 deg
K -> ": 0 deg
Root Locus Analysis of PitchRate Feedback to Elevator(2nd-Order Approximation)
Kkn(s)
d(s)= K
!q(s)
!"E(s)= K
kq s # zq( )s2+ 2$SP%nSP s +%nSP
2= #1
Center of gravity : doesn!t matter
Locus on real axis K > 0: Segment to the left of the zero
K < 0: Segment to the right of the zero
The feedback effect is analogousto changing Mq
Root Locus Analysis of AngularFeedback to Elevator (4th-Order Model)*
Flight Path Angle Pitch Rate
Pitch Angle Angle of Attack
* p. 524, Flight Dynamics
Direct Lift andPropulsion Control
Direct-Lift Control-ApproachPower Compensation
F-8 Crusader
Variable-incidence wing toprovide better pilot visibility
Flight path control at lowapproach speeds requires throttleuse, could not be made with pitchcontrol alone
Engine response time is slow
Flight test of direct lift control(DLC), using the ailerons as flaps
Approach power compensationfor A-7 Corsair II and direct liftcontrol were studied at Princetonusing our Variable-ResponseResearch Aircraft
Princeton VRA
Vought A-7
Vought F-8
Direct-Lift/Drag Control
Direct-lift control on S-3A Viking
Implemented with spoilers
Rigged up during landing toallow lift.
Speed brakes on T-45A Goshawkmake up for slow spool-up timeof jet engine
BAE Hawk's speed brake movedto sides for carrier landing
Idle speed increased from 55% to78% to allow more effectivemodulation via speed brakes
Lockheed S-3A
Boeing T-45
Next Time:Flight Testing
Supplementary
Material
Trailing-EdgeBevel Balance
See discussion in
Abzug and Larrabee
CH ! CH"" + CH# # + CHpilot input
Aft Flap vs. All-Moving
Control Surface
Carryover effect
Aft-flap deflection can be almost as effective as
full surface deflection at subsonic speeds
Negligible at supersonic speed
Aft flap
Mass and inertia lower, reducing likelihood of
mechanical instability
Aerodynamic hinge moment is lower
Can be mounted on structurally rigid main
surface
The Unpowered F4D Rudder Rudder not a problem under normal flight conditions
Single-engine, delta-wing aircraft requiring small rudder inputs
Not a factor for upright spin Rudder was ineffectual, shielded from flow by the large delta wing
However, in an inverted spin rudder effectiveness was high
floating tendency deflected rudder in a pro-spin direction
300 lb of pedal force to neutralize the rudder
Fortunately, the test aircraft had a spin chute
Powered Flight Control Systems
Early powered systems had a single
powered channel, with mechanical
backup
Pilot-initiated reversion to"conventional" manual controls
Flying qualities with manual controloften unacceptable
Reversion typically could not be
undone
Gearing change between control stickand control to produce acceptable pilot
load
Flying qualities changed during a high-stress event
Hydraulic system failure was common
Redundancy was needed
Alternative to eject in military aircraft
A4D
A3D
B-47
Advanced Control Systems
Artificial-feel system
Restores control forces to those of an
"honest" airplane
"q-feel" modifies force gradient
Variation with trim stabilizer angle
Bobweight responds to gravity and tonormal acceleration
Fly-by-wire/light system
Minimal mechanical runs
Command input and feedback signalsdrive servo actuators
Fully powered systems
Move from hydraulic to electric power
Boeing 777 Fly-By-Wire Control System Control-Configured Vehicles Command/stability augmentation
Lateral-directional response
Bank without turn
Turn without bank
Yaw without lateral translation
Lateral translation without yaw
Velocity-axis roll (i.e., bank)
Longitudinal response
Pitch without heave
Heave without pitch
Normal load factor
Pitch-command/attitude-hold
Flight path angle
USAF F-15 IFCS
Princeton Variable-Response Research AircraftUSAF AFTI/F-16
Root Locus Analysis of AngularFeedback to Thrust (4th-Order Model)
Flight Path Angle Pitch Rate
Pitch Angle Angle of Attack
United Flight 232, DC-10
Sioux City, IA, 1989 Uncontained engine failure damaged all three flight control
hydraulic systems (http://en.wikipedia.org/wiki/United_Airlines_Flight_232)
296 people onboard
185 survived due to pilot!sdifferential control of engines
http://www.youtube.com/watch?v=vTXr1QR3rbQ
Propulsion Controlled Aircraft Proposed backup attitude control in event of flight control system failure
Differential throttling of engines to produce control moments
Requires feedback control for satisfactory flying qualities
NASA MD-11 PCA Flight Test
NASA F-15 PCA Flight Test
Proposed retrofit to McDonnell-Douglas (Boeing) C-17
Recommended