Nonlinear flight dynamics problems analyzed using bifurcation methods

Pergamon
Plh S0376-0421 (97)00001-8
I'ro9. AerospaceSci.Vol.33,pp.539 586.1997
(~ 1997ElsevierScienceLtd
Allrightsreserved.PrintedinGreatBritain
0376-0421.,97$32.00
APPLICATION OF BIFURCATION METHODS TO NONLINEAR
FLIGHT DYNAMICS PROBLEMS
M. G. Goman, G. I. Zagainov and A. V. Khramtsovsky
Central 4erohydrodynamic Institute (TsAGI), 140160, Moscow Region. Zhukot:sky. Russia
(Received 12 December 1995; in final form 5 March 1997)
Abstract--Applications of global stability and bifurcational analysis methods are presented for different
nonlinear flight dynamics problems, such as roll-coupling, stall, spin, etc. Based on the results for different
real aircraft, F-4, F-14, F-15, High Incidence Research Model, (HIRM), the general methods developed by
many authors are. presented. The outline of basic concepts and methods from dynamical system theory are
also introduced. !:2 1997 Elsevier Science Ltd
CONTENTS
1. INTRODUCTION 539
1.1. Nonlinear aircraft problems 540
1.2. Approximate approaches 540
1.3. Continuation methods implementation 541
1.4. Bifurcation analysis and continuation technique methodology 541
1.5. The adequacy of a mathematical model 542
2. APPLICATION OF BIFURCATION METHODS TO FLIGHT DYNAMICS PROBLEMS 542
2.1. Autonomous forms of motion equations and steady state modes 542
2.1.1. Aerodynamic model for stall/spin conditions 544
2.1.2. Aircraft with control augmentation system 545
2.2. Critical flight regimes 546
2.3. Dynamics of symmetrical flight in a vertical plane 547
2.4. Nonlinear phugoid motion 547
2.5. Deep stall motion 549
2.6. The general case of the longitudinal motion 550
2.7. Roll-coupling problem 551
2.8. Dynamics at subsonic flight regime 553
2.9. Dynamics at supersonic flight regime 554
2.10. Nonlinear stall motion 558
2.11. Approximate linear criteria of stall 558
2.11.1. Controllability criteria 560
2.12. Stall dynamics analysis 561
2.13. Spin dynamics 564
2.14. Bifurcational analysis of spin motion 566
3. THEORETICAL BACKGROUND 576
4. CONCLUSIt)NS 584
ACK NOWLEDGEM ENTS 584
REFERENCES 584
1. INTRODUCTION
Nowadays there are a great number of specialized books dedicated to nonlinear dynam-
ical systems and bifurcation theory, and nobody can deny that this area of modern
mathematics is one of the most interesting and attracts specialists in numerous branches of
science and engineering.(x-t8)
Nonlinear world displays the universal features in its all manifestations. There are many
publications in such disciplines as physics, aerodynamics, biology, chemical kinetics, econ-
omics, etc. devoted to study of various nonlinear phenomena.
Nonlinear systems describing the behaviour of the objects of the real world usually are
too complex to be thoroughly studied using analytical methods. Hence the advances in

540 M.G. Gomanet al.
understanding essential features of nonlinear objects are closely coupled with the develop-
ment of numerical methods. The interest in the theory of nonlinear systems and in the tools
for computer-aided research rapidly grows. Today a researcher has at his disposal a number
of effective and reliable numerical methods intended for analysis and classification of
nonlinear phenomena. All of them were developed during the last ten/fifteen years.~19-3°~
And there is a need for more and more powerful algorithms and methods. There is also
a need for software packages eliminating the necessity for a researcher to write too many
lines of code for his computer. Partially, in flight dynamics applications the Packages
BISTAB, AUTO, ASDOBI, KRIT were used.tl~'2°'22`3xl
1.1. NONLINEAR AIRCRAFT PROBLEMS
Nonlinear aircraft dynamics problems to a great extent are encountered during high angle
of attack maneuvers and fast rotation in roll. Investigation of aircraft behavior at stall/spin
regimes and spatial maneuvers with strong roll-inertia coupling is extremely important for the
definition of the flight envelope boundary and solving the flight safety problem.
While studying these regimes, one has to take into account the nonlinear terms in the
equations of aircraft motion and nonlinear aerodynamic effects, especially for high angle-of-
attack region. Therefore, the problem to be studied is essentially nonlinear and is usually
formulated in the form of a set of ordinary nonlinear differential equations depending on
parameters. These parameters can characterize the flight conditions, mass and inertia
characteristics of aircraft. Very often the control surface deflections are considered as such
parameters, and bifurcation analysis in this case permits to study the dependency of global
dynamical behavior as the function of fixed control deflections. It is important to note that
the results of bifurcational and qualitative analysis give good basis for forecasting aircraft
dynamics in the case of varying control.
Modern aircraft are often intrinsically unstable, and stable flight is possible only due to
the control augmentation system (CAS). Both CAS mechanical elements (due to natural
constraints such rate saturation, limited deflections) and feedback algorithms introduce
additional nonlinearities. Therefore modern aircraft with flight control systems become
more nonlinear plant. Bifurcation and qualitative theory methods can also be used to
investigate the nonlinear closed-loop dynamics for verification of the implemented algo-
rithms, and therefore can be very efficient in control laws design.
One can divide all the published works devoted to aircraft nonlinear dynamics problems
into two groups. The first group contains the works with an approximate approach, and the
second one contains the works based on the continuation technique and wide implementa-
tion of results from modern theory of nonlinear dynamical systems.
1.2. APPROXIMATE APPROACHES
Stall/spin phenomena and roll-coupling instabilities were studied by many authors for
every new generation of aircraft starting from the beginning of aeronautical era. Both the
spin problem 1333"~1 and the roll-coupling problem t3~41~ were investigated using the
reduced equations of motion, simplified aerodynamic coefficients representation (as a rule,
only linear dependencies) and approximate methods for solving of these equations. The
roll-coupling instabilities were not strictly connected with the bifurcations in the equations
under consideration. For example, the critical roll rates first introduced by Phillips are very
useful for engineering estimates, but they are the result of the simplified consideration of the
problem and cannot be treated as bifurcations or stability losses. The main reason is that
the problem was analyzed by treating the roll rate (state variable) as a parameter in the
linearized pitching and yawing moment equations. For construction of the equilibrium
solutions of motion equations some artificial tricks were needed, e.g. the consideration of
the angle of attack (also the state variable) as some parameter. ~4t~Nevertheless, it was
shown, that the departures associated with roll-coupling instabilities resulted in the aircraft

Bifurcation methods in nonlinear flight dynamics 541
jumping from one stable equilibrium to another, i.e. the instability due to saddle-node
bifurcations was implicitly recognized,t38~ The wing rock phenomena at first were also
analyzed using an approximate approaches for nonlinear oscillations, without linking them
with Hopf bifurcation.~36~
1.3. CONTINUATION METHODS IMPLEMENTATION
The wide and purposeful applications of bifurcational methods and nonlinear dynamical
system theory to flight dynamics nonlinear problem at high angles of attack regimes are
now well-established.~42"44-47"50-53.60)
The common feature of all these works is the implementation of the continuation
technique for solving the nonlinear problems of an aircraft equations both for equilibrium
and oscillatory motions. Carroll and Mehra were the first who used the continuation
method in flight dynamics and connected the main types of an aircraft instabilities with
bifurcation phenomena (e.g. the occurrence of wing rock motion was connected with a Hopf
bifurcation, the examples of chaotic motion were presented, etc). Later the existence of new
types of bifurcations in aircraft dynamics were explored, e.g. origination of stable torus
manifold,~4~global bifurcation of a closed orbit related with the appearance of homoclini-
cal trajectory,~'Lg~so-called flip or period doubling and pitchfork bifurcations for closed
orbits.~So.5~ Continuation methodology and bifurcation analysis can be used for determin-
ing the recovery technique from critical regimes and for control law design for improving
dynamical behavior.~43'44"47'5x~
If the earliest works were devoted mainly to the theoretical methods and associated
numerical procedures which can be used in aircraft dynamics,~42"47~the latest publications
more often present the results of bifurcation analysis of high angle of attack dynamics of real
aircraft (F-4,~421German-French Alpha-Jet,~45~F-14,~53~F-15~sl~).
Comparisons between flight test results and predicted results obtained using bifurcation
methods and numerical calculations for a number of aircraft show a very good agreement
both in qualitative and quantitative senses,t45~These comparisons reveal the efficiency and
comprehension of bifurcational methods in flight dynamics applications. Therefore one can
say that bifurcation methodology based on the computer-aided technology is becoming
a popular and very powerful tool in the complicated nonlinear area of flight dynamics.
1.4. BIFURCATION ANALYSIS AND CONTINUATION TECHNIQUE
METHODOLOGY
Many characteristics, defining flight conditions, aircraft parameters and control surface
deflections, can vary with time slowly. In many practical cases it is reasonable to consider
such parameters as fixed ones and independent of time. This reduces the problem to a study
of nonlinear autonomous dynamical systems. The system behavior in the case of parameter
variations can be predicted using the knowledge about the specific system responses after
encountering the bifurcation conditions.
Taking into account experience of many researchers, one can formulate the following
three-step methodology scheme for the investigation of nonlinear aircraft behavior, the
scheme being based on bifurcation analysis and continuation technique:
• During the first step it is supposed that all parameters (expect for lhe state variables) are
fixed. The main goal is to search for all the possible equilibria and closed orbits and
to analyze their local stability. This study should be as thorough as possible. The
implementat:on of the continuation technique to a great extent facilitates the solving of
this problem. The global structure of the state space (or phase portrait) can be revealed
after determining the asymptotic stability regions for all discovered attractors (stable
equilibria and closed orbits). An appropriate graphic representation plays an important
role in the treating of the calculated and accumulated results.

542 M.G. Goman et al.
• During the second step the system behaviour is predicted using the information about the
evolution of the phase portrait with the parameters variation. The knowledge about the
type of encountered bifurcation and current position with respect to the stability regions
of other steady motions are helpful for the prediction of further motion of the aircraft.
The rates of parameter variations are also important for such a forecast. The faster the
parameter change, the more the difference between steady-state solution and transient
motion can be observed.
• Last, the numerical simulation is used for checking the obtained predictions and obtain-
ing transient characteristics of system dynamics for large amplitude state variable
disturbances and parameter variations.
1.5. THE ADEQUACY OF A MATHEMATICAL MODEL
The adequacy of mathematical modeling of high angle-of-attack dynamics is strictly
dependent on the adequacy of the aerodynamic model at these regimes. There is nontrivial
problem due to the very complicated nature of the separated and vortex flow in unsteady
conditions. Aerodynamics becomes extremely nonlinear and time history motion depend-
able. Qualitative and bifurcation approaches are also very fruitful when the sources and
nature of aerodynamic phenomena are considered. Special techniques were proposed to
represent the aerodynamic characteristics taking into account the dynamic effects of the
separated flow.~54'55~The bifurcation analysis of the Alpha-Jet aircraft was performed using
the dynamical unsteady aerodynamic model at high angles of attack. ~45~It is important to
include the rotary balance data into the aerodynamic model in order to obtain better
agreement with flight test results, t35'57~The bifurcation methods of analysis of the nonlinear
aircraft dynamics at high angles of attack can be efficiently used for validation of the
aerodynamic models for these complicated flight regimes.
2. APPLICATION OF BIFURCATION METHODS TO FLIGHT
DYNAMICS PROBLEMS
2.1. AUTONOMOUS FORMS OF MOTION EQUATIONS
AND STEADY STATE MODES
There are some conventional forms of equations of motion which are commonly used in
numerical and analytical studies of flight dynamics problems. If an airplane is considered as
a rigid body three different frames of references are used to write the equations of motion,
i.e. earth axeS--OEXEYEZE (inertial axes, flat approximation), body axes--OXYZ (axes
fixed in a rigid body), wind-body axes--OwXw YwZw (origin is fixed in the mass center,
OwXw is directed along the velocity vector V, and OwZw lies in the plane of the symmetry
of an airplane, t32~
The full set of motion equations can be divided into the following four groups:
• Dynamics of translational motion written in wind-body axes:*
m(/= Txw - D - mg sin Ow
Tz~ - L g
= q - (p cos ct - r sin ~) tan [3 + mV cos/3 + V----~osfl c°s 0w cos 0w (1)
Trw-C gcos0wsin/~ = p sin ct - r cos ~ + mV + v thw
*sin0w= - cos ~tcosflsin 0 + (sinflsin $ + sin ct cosp cos q~)cos 0.
cos 0wcos Sw = sin ctsin 0 + cos~tcos 0cos $.
cos0wsin Sw = cosctsin flsin 0 + (cos/3sin q~- sin ~tsin flcos$) cos 0.

Bifurcation methods in nonlinear flight dynamics
• Dynamics of angular motion written in principal body axes:
543
lx/~ = (It - l~)qr +
lr(l = (Iz - lx)rp + ,tl (2)
Izt: = (Ix - lr)pq + jr
• Kinematics of angular motion written for wind-body axes (the same equation are valid
for body axes without subscript W):
0w = qw cos q~w-- rw sin ~bw
qSw= Pw + (qw sin q~w + rw cos (kw)tan 0w (3)
6w = (qw sin ~bw+ rw cos Sw) sec 0w
• And kinematics of translational motion written in earth axes:
XE = V cos 0w cos ~'w
I;'E= V COS0w sin ~Ow (4)
2~E= -- V sin 0w
where altitude of flight H = - ZE. Dotted symbols denote time derivatives.
The basic idea of the investigation of flight dynamics is the concept of steady states. The
steady state of a flying vehicle is considered in a nonformal mathematical sense for
a complete set of equations. Steady state or equilibrium is considered in a more general
dynamical sense as the steadiness of all the external forces and moments, i.e. as the
aerodynamic steady state. This condition requires that the main motion parameters
(V, :¢,/L p, q, r) ~tndgravity projections (i.e. the Euler angles 0, ~b)are all constant in the body
axes frame with time. The position in space XE, YE and head angle ~Odo not affect on the
aerodynamic sleady state, and therefore can be considered separately, together with
appropriated equations. The altitude of flight H (or ZE) defines the air density and can be
added to the system of equations only in cases when the influence of density variation on
dynamics behavior is significant. If the variation of p with altitude can be neglected, the
most general case of equilibrium flight is a vertical helicoidal trajectory. These may be
climbing and gliding turns with large radius of curvature, or steady equilibrium spin modes
with relatively small radius of trajectory curvature. The spacial case of such trajectories is
the rectilinear motion.
Therefore to study flight dynamics the following autonomous system of equations can be
extracted from the full system (Eqs (1)-(4))
dx
dt F(x, ~) (5)
where x = (~, ~, p, q, r, V, 0, ~b)'e R8, ~ = (fie,64, 6,, T)' e R4 are the state vector and the
control vector. The steady state regimes, which are the vertical helicoidal trajectories, are
defined by equilibrium solutions of this system of equations.
Some additional physical assumptions, concerning the type of motion, such as the
existence of the plane of symmetry, steadiness of some state variables, etc., can be taken into
account for obtaining the approximate autonomous subsystems of equation of lower
dimension for :investigation the flight dynamics. Such subsystems can be derived for
symmetrical flight in vertical plane, for studying the roll-coupling problem, etc. (see the
following sections).

544 M. G. Goman et al.
2.1.1. Aerodynamic Model for Stall~Spin Conditions
The aerodynamic model intended for spin conditions, i.e. high angles of attack and fast
rotation, is based on the experimental data obtained from the different kinds of wind tunnel
tests--static, forced-oscillation and rotary balance. There exist a number of methods for
designing the 'combined" mathematical model of aerodynamic coefficients, which imple-
ment the experimental data in a very similar manner. 157'-~81
The rotation at high angle of attack can significantly influence the flow pattern. As
a result, the aerodynamic coefficients become nonlinear functions of the reduced rate of
rotation. That's why the aerodynamic coefficients measured in rotary balance tests are
considered as the basic or 'nondisturbed' part of the aerodynamic model for high angle-of-
attack conditions.
The disturbed motion is accompanied by the misalignment between the velocity and the
rotation vectors. The projections of the rotation vector onto the wind-body axes can
be used as parameters for describing the disturbed conical motion. The roll rate in wind-
body axes --Pw = (P cos ~ + r sin ~) cos fl + q sin fl -- defines the rate of conical rotation,
which is similar to the angular rate in the rotary balance tests. Two other projec-
tions--qw = - (p cos :~ + r sin :~)sin fl + q cos fl, rw = r cos :~- p sin :~--define both the
unsteadiness and spirality of motion: i = (qw - qw,)/cos fl,/~ = - rw + r,~. The values of
qx%, rv~, arc the wind-body angular rate in steady state spiral motion and their undimen-
sional values qw,o::/2V, rw, b/2 V are usually very small.
Assuming that the disturbances of the pure conical motion are small, the following
representation of the aerodynamic coefficients can be used:
c, = c,.. . X + c,,,,, Ty + + c,. + y:
_ qwE rwb
= G,, (~, 13,Pw, ~) + (C~,~ + C~;:"cos/~) ~p- + (G, - G;) 2 V
qw,ri: rv%b (6)
+ - Ci; 2V cos fl + Ci'; 2V
The derivatives of the aerodynamic coefficients, standing together with the reduced rates
of rotation and corresponding to the rotary flow, can be measured, by means of the
oscillatory coning technique, such as already in use at ONERA/IMFL. 15~)The terms with
qw,pand r~o can be omitted.
In many cases the data obtained from 'traditional' forced oscillations tests (recall, that
they measured in the absence of the model rotation) are used. The following approximate
transformation, which is valid for zero sideslipe, may be used:
('i~ + Ci. "" C;~
. . . . . (7)
Ci,,, - Ci~ ~- Ci,, cos ~o - C,,.. sin ~o
where the subscript f.o. denotes the data obtained in forced oscillation tests.
When the nonlinear term in (6) can be approximated by a linear function on the angular
rate, the representation (6) becomes equivalent to the aerodynamic model commonly used
for low angles of attack. The results of the static wind tunnel tests in this case can also be
incorporated into the mathematical model.
The form of the aerodynamic model (6) for stall/spin conditions is quite natural. For
example, the rotary derivatives Ci,, and C,,, do not significantly affect either the values of
kinematic parameters at the equilibrium spin or their mean values during the oscillations
with moderate amplitude. These derivatives as well as unsteady derivatives Ci~ and
C~i directly affect on the stability margin of the oscillatory spin mode. Thus they determine,
for example, the amplitude of the 'agitated' spin motion, when the equilibrium spin is
oscillatory unstable.

The rotary balance data Ci,, (~, fl, pwb/2V, fi) in the aerodynamic model allows to get
realistic values of the equilibrium spin parameters. To improve the time histories and
amplitudes of the oscillations, one can make some adjustments (if necessary) to rotary and
unsteady derivatives C~,w,Ci.... Ci~, C~i.
All the wind tunnel data are measured and tabulated in a wide state and control
parameter ranges. To facilitate the implementation of the continuation technique the
aerodynamic functions are usually smoothed by means of spline or polinomial approxima-
tion to ensure continuity and derivability conditions for the resulting nonlinear dynamic
system.
Aerodynamic asymmetry is one of the important features of aerodynamic coefficients at
high angles of attack. Asymmetry appears at zero sideslip, rate of rotation and symmetrical
aileron and rudder deflections. It may be larger than maximum aileron and rudder
efficiency. Asymmetrical roll moment can significantly influence stall behavior: asymmetri-
cal yaw moment arises at the higher angle of attack region and to a great degree defines the
spin dynamics. Due to asymmetry, the right spin modes can greatly differ from the left ones
both in the values of parameters and the character of stability.
Unsteady aerodynamic effects at high angles of attack, resulting from the separated and
vortex flow development, can significantly transform the real aerodynamic loads with
respect to their conventional representation, which was discussed above. In the frame of the
conventional approach the unsteady effects are described using linear terms with unsteady
aerodynamic derivatives. There are special regions of incidence, e.g. CL.... region, where
the conventional representation is not valid. The special approaches using the differ-
ential equations or transfer functions can improve the aerodynamic model and take into
account the nonlinear unsteady aerodynamic effects due to separated and vortex flow
dynamics}46,55~
2.1.2. Aircraft with Control Augmentation System
Modern airplanes are equipped with automatical control systems able to change its
dynamical properties radically. Aircraft with flight control systems become more nonlinear
and higher dimensional plant. In addition, a flight control system introduces further
nonlinearity and additional dynamic elements. Even in the case where an aircraft may be
represented as a linear system, there will be appreciable nonlinearity due to actuator rate
and deflection limits. The nonlinear behavior will be displayed during large amplitude
motion at large control inputs or gust disturbances.
Nonlinear stability and bifurcational analysis methods also can be implemented for
aircraft closed-loop system. The equations governing the operating of the control system
can be added to the aircraft Eq. (5):
_d~i=G x,-- ~,s (8)
dt dr'
where s = (xe, :%, x..... )' is a vector of the stick and rudder pedal deflections, and the
vector-function G is determined by the control laws and control system constraints.
Thus, for example, the followingjoint system of Eq. (9) is to be used for determining the
equilibrium regimes of the airplane:
F(x, ~) = 0
(9)
G(x, 0, & s) = 0
The elements of the vectors x and ~ are the unknowns, and vector s defines the set of
control parameters.
The specific teatures of Eqs. (5) and (8) are the limits on the values of some state variables
(i.e. the deflections of the control surfaces 6:(6 .... < 6,. <~6.....: b6~l~<6......; [6,1 ~<6....). In
nonlinear system with such state variable limits some additional equilibrium solutions can
arise with associated limit points bifurcations leading to aircraft departures.

546 M. G. Gomanet al.
2.2. CRITICAL FLIGHT REGIMES
The maneuvering capabilities of the airplane are usually limited by some boundary.
Outside this boundary lies the area of critical flight regimes, i.e.the regimes when uncontrol-
lable motions (connected with the loss of stability) develop. Pilots" actions may either
provoke or prevent the development of the instability. It depends on the task being
performed and on the pilot's skill in flying the plane in such regimes3561
A desire to enhance maneuverability inevitably results in entering the regions, where
aerodynamic characteristics are nonlinear and dynamic cross-coupling between different
forms of motion cannot be ignored. As a result, the dynamical problems become highly
nonlinear.
All the critical regimes can be divided into two groups, according to the reasons of
dangerous behavior of the vehicle.
The regimes of the first group arise when the stability margin is broken and the unstable
modes of motion begin to develop. These regimes are very different. There may be mild or
abrupt loss of stability, and the motion may be controllable or uncontrollable. Such
situations take place during stall or maneuvers with fast roll rotation.
The second group comprises the stable steady-state flight regimes with supercritical
values of parameters (especially angle of attack and rate of rotation). The roll-inertia
rotation or autorotation rolling and spin regimes belong to this group; the unusual response
to the control inputs is their characteristic feature. Departure resistence to the entering in
these critical regimes and the methods of recovery from them are the most important
questions, when the problem of flight safety is considered (see Fig. 1).
A number of different forms of the loss of stability in longitudinal and lateral/directional
motion is usually connected with aircraft stall concept.1561One may distinguish among
them:
• a sudden rise of the angle of attack ('pitch up') taking place at moderate and high angles
of attack due to pitching moment nonlinearity;
• stable self-oscillating pitching motion at high angles of attack due to the development
of the separated flow ('bucking');
• the full turn rotation in pitch ('tumbliny');
• divergent increase in the bank angle ('wing drop' or 'roll off') due to asymmetrical
aerodynamic rolling moment at high angle of attack or propelling aerodynamic
moments;
• divergent increase in sideslip angle ('nose slice" or 'yaw off') due to aperiodic instability
in yaw or asymmetric aerodynamic moments;
• oscillating motion in roll and yaw at high angles of attack ('winy rock');
• loss of stability due to the pilot's actions, for example, when the attitude stabilization or
target tracking is performed.
[~ S-l
rotation
rate
gflying l ,"
de~ Pstall
20 30 40 50 60 70 8010
Angle of attack, deg.
Fig. 1. Criticalflightregimesregions.

Bifurcationmethodsin nonlinearflightdynamics 547
There are also terms used in the flight dynamics for describing the aircraft behavior after
stall, i.e. post-stall gyration, spin and deep stall. The post-stall gyration is a transient
rotational motion of aircraft to the developed spin mode. The developed spin can possess
very different features. It can be steady with constant parameters or 'agitated',flat or steep,
erected or inverted. 'Agitated' spin mode can be with regular and 'irregular' oscillations.
Deep stall is art aircraft equilibrium flight at high critical angle of attack without rotation.
The possibiltty of the loss of stability and controllability at high roll rate (roll-coupled
problem) is also well-known. This phenomenon is especially dangerous for supersonic
aircraft with elongated ellipsoid of inertia and high level of lateral aerodynamic stability,
which bring about cross-coupling of longitudinal and lateral motions.
Roll-inertia rotation or autorotation rolling of the airplane (the trajectory being approxi-
mately horizontal and angles of incidence below stalling) is similar to the developed
spin modes. The rotation can occur despite neutral or anti-rotation aileron and rudder
deflections.
In both cases there are autorotational regimes, the difference is only in the nature of the
aerodynamic moment that supports the rotation.
All types of motion mentioned above can be connected with qualitative features of
motion equations, i.e. the different steady-state aircraft motions and bifurcations, changing
their stability conditions.
2.3. DYNAMICS OF SYMMETRICAL FLIGHT IN A VERTICAL PLANE
To demonstrate the nonlinear dynamics behavior of symmetrical flight using the bifurca-
tional and global stability analysis methods one can consider two simplified examples
arising basically from trajectory or angular pitch motion.
Taking into account that in symmetrical flight the state variables are equal to zero:
q~w = 0, ~ = 0, fl = 0, Pw = rw = 0, and 0w = 7• [- n,n], where ~' is a path angle, the
following equations of motion can be extracted from the complete system (Eqs (1)-(4)):
Tcos • CopS )
m~'=g mg ~ V2-sin)'
";' g (.T sin ~ CLpS )
= -- + V2 _ cos 7
V mg
4=q-,}
= Cm(~,q, fie)p V2S?
2Ir
(10)
)(E = VCOS~
//= Vsin7
The first two equations of this system are responsible for the so-called phugoid trajectory
motion, the second two equations are responsible for short-period rotational motion. At
relatively large velocities of flight these two subsystems approximately can be considered
separately.
2.4. NONLINEAR PHUGOID MOTION
This problem was first considered by Zhukovsky and Lanchester and later became one of
the classical examples of second-order nonlinear systems, illustrating the qualitative
methods of analysis.

Assuming that the angle of attack is a constant, and therefore the lift and drag coefficients
are also constants, the following nondimensional autonomous second order nonlinear
system can be used for investigation of phugoid motion with large amplitudes:
--cl~) = b - ay 2 - sin 7
dr
d), cos 7
dr - y Y
(11)
T CD.
where b m9 is the thrust-to-weight ratio considered as a parameter (:t <~1), a ~. ts the
2mg
drag-to-lift ratio, r = VOqt is the nondimensional time and Vo = ~]pSCLis the velocity of
horizontal flight.
The equilibrium solutions of this system at different values of parameters a and b are
~:orresponded to straightline trajectories with climb (7 > 0) or glide ('/< 0) path angle. There
is some critical boundary in parameter space exceeding which makes the equilibrium
oscillatory unstable. At some values of parameters a, b another kind of steady solution
appears. These are the stable and unstable closed orbits, enclosing the cylindrical state space
or equilibrium points. The stable closed orbits, enclosing the cylinder correspond to the
curvilinear trajectories with recurring loops, the so-called 'dead loops'. The unstable closed
orbits, enclosing the cylinder or the equilibrium points, separate the regions of attraction of
the stable equilibrium points and the stable closed orbits, enclosing the cylinder.
In Fig. 2 the bifurcation diagram is presented when the parameter b is varied for a = 0.2.
The path angle of equilibrium straightlinear flight ~ becomes positive, when b >/a. When
parameter b exceeds the critical value bcr, ,~ 0.4 stable and unstable closed orbits, enclosing
the cylinder, appear. After that the stable cycle moves up and the unstable one moves down
along velocity axes y. There is some bifurcation moment (bcr2~ 0.42), when the unstable
cycle involves the critical points y = 0, 7= + ~/2. These critical points in state space
correspond to the cusp points on the flight trajectory, where the path angle varies from
7 = ~/2 to ?, = - n/2 instantaneously. In real flight it is impossible due to inertiality of
angular pitch motion. In this case, corresponding to the well-known aircraft maneuvers
y n
rd2
-~,e2
stable closed orbits,
surroundingcylinder
/
, i "
closed orbits ~ J
stable e.quilibria~ unstable|I~I I
.------ ............ equilibria
4.2 0 0.2 0.4 0.6 0.8 Wmg
Fig. 2. Equilibrium and periodical solutions as a function of parameter b with a = 0.2.

Bifurcation methods in nonlinear flightdynamics 549
with large loss of speed (the so-called 'tail-slide' maneuvers), the full system (10) has to
considered. During the tail-slide maneuvers the extremely large variations of angle of attack
can occur up to ~ ,~ 180° and therefore the assumption about ~ = const taken above is not
valid in these points.
When b >/be,.2the unstable closed orbit enters in the internal region of the phase portrait
rt
between 7 = + ~ and encloses the equilibrium points. With further increase of parameter
b the size of the unstable cycle becomes smaller and at bifurcation point b = bCr~~ 0.55, it
disappears. This point is the undercritical Andronov-Hopf bifurcation, where the equilib-
rium point becomes also oscillatory unstable. The unstable closed orbits are important
because they separate the regions of attraction of the stable equilibria and the stable closed
orbits which enclose the cylindrical phase space.
In Fig. 3A-D four pictures of phase portrait for the system (11) are shown for the
following values of parameter b = 0.2, 0.41, 0.5, 0.7 and a = 0.2. In Fig. 4A, B two kinds of
flight trajectories are shown for b = 0.2 and 0.7, respectively.
2.5. DEEP STALL MOTION
For a number of aircraft configurations, e.g. with T-tail or taillness, there is the essentially
nonlinear dependency in pitch moment coefficient generating the stable trim at high angle
of attack even under the full nose-down elevator deflection. (561 These regimes are called
deep stall, and are dangerous for flights with high incidence. The lack of pitch-down control
moment may be critical for pitch stall and recovery from it.
a=0.2, b=0.2 a=0.2, b=0.41
2.5 2.5 stableclo~ed " .
orbit,sumrounding u~table closed
s cylinder: A orbh~mr,mding
1 " 1
0.5 0.5
0 0
-~ --~/2 0 ~/2 ~ -~ -~/2 0 ~'2
(A) (B)
a=0.2, b=0.5 a=0.2, b=0.7
3 ~ 3
stabledosedorbit ~ stable
sunroundingcylinder~. equilibrium
2 ~ 2
1 1
0--~ --~/'2 0 ~ n -~ -~r2 0 ~2 ~
(c) Y (D) Y
Fig. 3. Cylindricalphase portrait for y and ;.' state variables, a = 0.2 and A) h = 0.2, B) b = 0.41, C) b = 0.5,
D) b = 0.7.

560
420
280
140
a = 0.2o b = 0.0
560
420
280
140
i I I i
240 480 720 960
(A) L,m (B)
a=0.2, b=0.7
I I I I
240 480 720 960
L,m
Fig.4. Flighttrajectorieswith'dead' loops,A) b = 0, B) b = 0.7.
The understanding of nonlinear dynamics and the prediction of the region of attraction
generated by the critical deep stall regime are very important for stall prevention and also
for design of the recovery control from deep stall. To form the approximate equations one
can consider that the velocity of flight is not varied and the path is rectilinear:
~=q
( t~1: Cm(~)@ CmqV -[- Cm6e~e T SC
(12)
The qualitative analysis of nonlinear dynamics in this case can easily be done in the phase
plane of the state variables (~ and q) (see Fig. 5). There are three different equilibrium points
:~,, ~2, ~3, two of which (al, ~3) are stable loci and one (~2) is the saddle point. This unstable
saddle-type equilibrium point generates very informative phase trajectories W~, which are
going in it along the eigenvector, corresponding to the stable eigenvalue. Two other
trajectories W uare going out from saddle point along the eigenvector, corresponding to the
unstable eigenvector. These trajectories are attracted by stable points ~1, ~3. All these
singular trajectories can be computed by means of integrating the system (12). Initial
conditions for reconstruction of these singular trajectories are taken in a small vicinity of
the saddle point ~2 with displacements along the stable and unstable eigenvectors. The
trajectories W u are integrated in the positive time direction, and W s are integrated in the
backward time direction. Note, that the region of attraction of the deep stall regime ~3
(dashed area in Fig. 5) depends on the altitude and velocity of flight, and the magnitude of
the pitch damping derivative Cm.
2.6. THE GENERAL CASE OF THE LONGITUDINAL MOTION
As was mentioned above the interaction between the phugoid and angular modes is more
significant for flight with low velocities. In this case the large variation of angle of attack can
arise due to trajectory distortion. For example, such trajectory distortion accompanies the
tail-slide maneuvers.
The region of attraction of critical deep stall regime in the general case will be more
complicated with respect to the region, considered above in simplified manner. To take into
account the interaction between phugoid and angular modes the first four equation of (10),
which are the autonomous nonlinear system with state vector X = (V, 0, ~, q)r, have to be
considered with entire ranges of .~ and 0 variations--I- ~t,n]. To represent the stability
region in the fourth-order state space it is possible only by means of drawing its two-
dimensional cross-sections considering the disturbances only in two selected state variables.

0.5
region of attraction
of a deep-stall
W s regime o~3
Cm
Xt~ 3o 5o ctc~
0.25
0t t
Ot2
IX3
-0.25
I I I
l0 30 50
ct (deg)
Fig. 5. Phase portrait for the deep stall nonlinear dynamics.
For example, in Fig. 6 two different cross-sections of the region of attraction of deep stall
regime, i.e. stable point ~3, are shown. The disturbances in the plane of pitch angle 0 and
velocity V are considered. Two other state variables at the initial moment are the same for
all points of cross-section. In the first case (A) the angle of attack is trimmed in a lower stable
point ~t = ~tt ~ith zero pitch rate q = 0, and in the second case (B) the angle of attack is
trimmed in the critical position ~ = ct3, also with zero pitch rate q = 0. In every point of the
considered cross sections the initial path angle can be calculated using the following
formula ~ = 0 - ~t.The dashed areas on the cross sections define the initial points, starting
from when the aircraft enters in the deep stall regime, ~t3. In the second case (B) the
probability of entering into the deep stall is much greater, especially in flights with large
velocities, than in the first case (A). The number of cross sections of multidimensional
stability region can offer global information about the aircraft dynamics.
2.7. ROLL-COUPLING PROBLEM
The common form of steady state is the vertical helix trajectories which result from the
balances both in the moment and the force equations. The process of balance adjustment in
the moment equation is usually much faster than the process of balance adjustment in the
force equations; this is especially true in flights with large velocities. The specific time of the
trajectory adjustment (to the vertical helix) is adverse proportional to the magnitude of the
gravitational terms in the force equations 9/V.
The differential equations for trajectory angles Ow and ~bw
Ow = ~g (ancos qgw + a~.sin q~w - cos 0w)
g (a. sin ~bw - a, cos ~bw)tan Ow ~. flq~w---Pw +~
(13)

2O 40
n V
(A)
I ! ! I
60 80 100 120 V, m/s
V
0
////,
K
(B)
Fig. 6. Cross-sections of asymptotic stability region for deep stall regime z~3.
L - Tzw C- Trw
where a. - - - , as - - - are normal and side load factors; can be used for the
mg mg
estimation of the averaged variation of the trajectory angle during fast rotation. The
averaging of these equations during the period of one turn T = 2n/~ yields:
g
0w = - ~ cos 0w (14)
This equation can be taken in quadrature. The time when the average trajectory angle will
be changed, e.g. from 0 to - 45 ° is T4s ~ 0.88 V/9. The influence of the gravitational terms
in the equations of motion will be negligible, if the number of turns during this time will be
large, i.e. n = T45/T ~ 0.14, ~V/g ~> 1.
The role coupling problem at fast roll rotation is usually considered without taking into
account the gravitational terms. No less important when considering this problem is that
velocity of flight is fixed and the first equation in (1) is omitted.
Therefore only the last two equations from (1) without gravitational terms and three
moment equations from (2) can be left for consideration:
0~= q - (p cos ~ - r sin ~) tan 13+ z
/~ = p sin ~- r cos ~ + y
l) = - ixqr + l (15)

Bifurcationmethodsin nonlinearflightdynamics 553
gl = i2rp + m
= -- i3pq + n
I, -- Iy, i2 I. - Ix, i3 Iy - I~where il . . . . ~ = - - are nondimensional inertia coefficients,
I~ Ir I~
l = ~/lx, m = ,4[/I r, n = .A~/lz, y= - C/mV, z = - L/mV. The reduced normal and side
forces z,y, pitch, roll and yaw moments m,! and n in Eq. (15) can be represented as
a function of state variables and control parameters both in linear or nonlinear forms:
z =z(~) + zq(~)q +z~,(a)fe+ -..
Y =Yo + Yo(~)fl+ y,(~)r + y~,(~)O, + ...
m = m(~) + mq(~t)q + m~,(ot)~ + Am(~, fl.... )
l = la(~)fl + lp(~)p + l,(~)r (16)
+ lo,(cO~,+ lo,(~)6. + Al(a, fl.... )
n=no(a)fl+np(a)p+n,(a)r
+ n6,(ot)6, + n6o(~t)fi, + An(a, fl, ... )
Equation (15) completed by (16) form the closed nonlinear autonomous system of the fifth
order with the state vector x = (a, fl, p, q, r) r. The nonlinearities in these equations can be
divided into three groups: kinematic, inertia moments and aerodynamics terms. All these
terms can lead to the cross-coupling between the longitudinal and lateral modes of motion.
An attempt to take into account the influence of the gravitational terms in the fifth order
equation (15) was made in Ref. 39.
The main feature of the problem considered is that even in the case of linear representa-
tion of aerodynamic coefficients the existence of multiple stable steady-state solutions, e.g.
equilibrium and periodic, is possible. The bifurcational analysis of all the possible steady-
state solutions and their local and global stability analysis can show the genesis of stability
loss and explain in many cases the very strange aircraft behavior.
The validity of the system (15), (16) is confined in time. Therefore, in the cases of weakness
or lack of stability of the steady states the conclusions resulting from the consideration of
the asymptotic stability in Lyapunov sense, when t ~ ~, can be wrong. A similar problem
can arise for short-term control inputs. The prediction of the bifurcation analysis will be
more consistent when the considered steady states have the sufficient margin of asymptotic
stability. In any case the numerical simulation of aircraft motion using the complete set of
equations has to be used for final verification of the bifurcational analysis results.
To demonstrate the possibilities of bifurcational and global stability analysis when
roll-coupling problem is studied, two different examples will be considered. The first one is
taken from Ref. 32 (pp. 447-451) and corresponds to a small maneuverable single-enginejet
airplane at flight with zero altitude H = 0 and subsonic velocity V = 500 fps. The second
one corresponds to hypothetical swept-wing fighter at flight with altitude H = 20000 m and
supersonic velocity V = 1800 fps. The types of equilibrium solutions in these cases are
different due to various contributions of damping terms in motion equations.
2.8. DYNAMICS AT SUBSONIC FLIGHT REGIME
In the first case (H = 0, V = 500 fps) all the equilibrium solutions form the single
continuous surface. The equilibrium surface in this case possesses the canonical form of
a singularity in the mapping of the equilibrium surface on the plane of control parameters
6e, 6,, which are called 'cusp catastrophe' and 'butterfly catastrophe'. 2 Although the

554
autorotationrolling
regimes
M. G. Goman et al.
rollrate
equilibriumsurface
bifiarcational
diagram
~a
Fig. 7. Surfaceof equilibriumroll rates. Bifurcationaldiagramin the plane16,.6~).
catastrophe theory was developed for the gradient type of dynamical systems, the for-
mulated singularities are, nonetheless, also apparent in the autonomous dynamical systems.
In Fig. 7 the common view of the calculated equilibrium surface is shown along with the
bifurcational diagram below it in the plane of elevator and aileron deflections. The number
of equilibrium solutions, shown in the figure, varies with control inputs. There is a critical
region with auto-rotational rolling regimes at pitch-down elevator and near-neutral aileron
deflections ('butterfly catastrophe'). At pitch-up-elevator deflections there are two 'cusp
catastrophes', which lead to the hysteresis type behavior under the roll control. The
quantified dependencies of equilibrium roll rate p on the aileron for a number of elevator
deflections, corresponding to different initial values of normal factor (az = - 4 - 3), are
presented in Fig. 8b. Another set of such dependencies on elevator deflections for 6, - var is
presented in Fig. 8a (solid lines--stable solutions, dashed lines--divergent solutions, dash-
dotted lines--oscillatory unstable solution). The approximate Phillips' critical roll rates
P~ = ~/-~-~, P~ = "v'~, obtained without taking into account the damping terms,
nevertheless define the roll-rate regions where the roll-coupling effect is more significant.
These critical roll-rate values are shown in Fig. 8b by horizontal dashed lines. The oblique
dashed line defines the controllability in roll mode without taking into account the dihedral
roll moment, arising due to the inertia coupling of longitudinal and lateral motions.
2.9. DYNAMICS AT SUPERSONIC FLIGHT REGIME
In the second case (H = 20000 m, V = 1800 fps) the critical roll rates really exist and are
very close to Phillips' approximate values. As a result the equilibrium solutions are divided

4
o :
~2"Io
0
-'-2
0
-4
-6
4
o
=~--~--S~--S >l 0_ '-
-4
I
' ' --6
-10 -5 0 5 10
(A) Elevator deflection, deg (B)
-20 0 20
Aileron deflection, deg
Fig. 8. Equilibrium roll rates for different elevator and aileron deflection (M < 1). A) t5e - var, •° = 0,
B) 6a - var, a: = - 0.4 + 3.0. See text for explanation.
1o
8
6
4
2
-4
-6
-8
-10
-30
I I I I I
.... ~-.,-. :- ~.: .-.....~.-- - - ; ~ . . . ~ . - ~ . . p ~ . . _
...... ~---=~-=~ .... : .... i .... ~ .... =_.~ ;.... :.z. ~.!:-..~- ....
I I .... I I I
-20 - I 0 0 10 20 30
Aileron deflection, deg
Fig. 9. Equilibrium roll rates as a function of aileron deflection for initial trimmed flight with az = - 2 + 5
(M > 1). See text for explanation.
by this critical line into different unconnected families, which are shown in Fig. (9). The
dependencies of equilibrium roll rate on aileron deflection are presented for different
elevator deflections, corresponding to different values of normal factor parameter
az = -2- 5. As in the previous case the type of line defines the code of equilibrium
stability (solid lines--stable solutions, dashed lines~divergent solutions, dash-dotted
lines--oscillatory unstable solutions).
There are two different subcritical types of curves starting from the zero point. The first
ones for positive az > 0 possess the 'loss of controllability' -- the increase of control
moment does not proportionately increase the roll rate. This effect is due to the arising of
opposite roll moments from sideslip, resulting from roll coupling. The second ones for zero
and negative a, ~<0 possess the departure point, where stable and divergent equilibria
disappear.

There is also another family of equilibrium curves both between the critical rates p,, pa
and in the outer regions. Some of them, corresponding to a: = 1, 0, - 1, - 2, generate the
stable autorotational roiling regimes at zero aileron deflection. The autorotational rolling
solutions for a: = 0, - 1, -2 exist at all aileron deflections, but at 'pro-roll' aileron
deflections (p > 0, 6a > 0) the autorotational equilibrium solutions become oscillatory
unstable (dash-dotted lines) and after the Hopf bifurcation point on each curve the family of
stable closed orbits arises.
A similar oscillatory instability of equilibrium solutions appear at subcritical equilibrium
curves at large aileron deflections. The Hopf bifurcation points in this case also give birth to
the families of stable closed orbits. To illustrate this in Fig. 10A the different branches of the
equilibrium curves and the families of closed orbits placed on the subcritical and supercriti-
cal equilibrium branches for case with 6~ = - 5.4: (a: = 1.0) are presented.
The envelope curves defining the maximum and minimum values of roll rate in oscilla-
tory solutions at 6a > 0 pass through the Hopf bifurcation points H1 and Hz. The
oscillatory solutions originating in the point Hz are stable at all values of aileron deflec-
tions. But the oscillatory solutions originating in the point H1 on subcritical equilibrium
branch become unstable at 6= ~ 18.9. After this the cascade of flip or period-doubling
bifurcation leads to the appearance of chaotic motion.
The bifurcation tree for the closed orbits originating in point H~ was obtained by
numerical Poincar6 mapping with the cross section plane fl = 0 for different aileron
deflections. It clearly shows the chaotic motion appearance (see Fig. 10B). The amplitudes
of the state variables in the stable oscillation regimes can be seen in bottom of Fig. 10C, D,
where the (~, [/)- and (p, fl)-projections of the closed orbits for different aileron deflections
are presented.
The roll-maneuver dynamics is highly dependent on the magnitude of control input. The
examples of transient dynamics, obtained for different step-like aileron inputs by means of
10
0
® 5In
d 0
~ -5
-10
~,~ closed orbits
-20 0 20
(A) Aileron deflection, deg
10
2
°
-10
18 20 22 24
(B) Aileron deflection, deg
80
60"o
o 40¢g
¢o
,e,-
o 20
o)
c 0
,<
-20
10
(.1
• 5
'1o
$ o
-10
-40 -20 0 20 -40 -20 0 20
(C) Sideslip angle, deg (D) Sideslip angle, deg
Fig. 10. Periodical and chaotic behavior at fast-roll maneuver. See text for explanation.

• 60
-o 40
•g 20
0 0
< -20
0 20 40 60
•1o 0 "o 0
-so ~ -so
0 20 40 60
I ' ' 1
~o ~o
~-5 ~-5
557
0 20 40 60
~.2o • ~ 20
" 0 d 0
a o
-20 ' ' ' -20
0 20 40 60
(A) Time, $ec
"o 40
<" 20
0 0
< -20 . . . . . . . . . . . . . . . . . . . .
0 20 40 60
i , , ,
0 20 40 60
0 20 40 60
0 20 40 60
(B) Time, sec
Fig. 11. Numerical simulation of roll-maneuver dynamics. See text for explanation.
numerical simulation of(15), (16), are presented in Fig. 11A, B. At (5 = 8° (see Fig. 11A) the
roll rate p varies aperiodicaily to the constant value p ~ - 2.0, corresponding to the stable
equilibrium (see Fig. 10A), the angle of attack :~ and sideslip/~ in this case are practically
constants. After applying the control input with (5 = 18° the periodic oscillations in all the
state variables are established. Their amplitudes also correspond to that calculated before
closed orbits (see'.Fig. 10A). Applying the larger control input (5 = 23° (see Fig. 11B) leads to
the jump at t > 10 s from one side rotation to another, with large amplitude periodic
oscillations in all state variables. The jump from the chaotic motion region, which became
unstable, is possible both to the stable oscillations at high angles of attack and positive roll
rate and to the stable equilibrium point with very fast negative roll rotation p ~ - 8 1/s.
The 91obal stability analysis or investigation of asymptotic stability regions is very
important in the case when there are multiple locally-stable equilibrium points and closed
orbits. The reconstruction of these stability regions gives the global information about
phase space of considered systems and determines the critical disturbances of the state
variables leading to the loss of stability. As already noted, the more convenient way to
reconstruct the multidimensional stability region is the calculation of its twodimensional
cross-sections. The disturbances only in two selected state variables under fixed other ones
are considered. It is natural, for example, that a cross section would pass through the locally
stable equilibrium point under consideration.
At zero aileron and rudder deflections (sa = (5, = 0 and (5~= - 5.4° (az = 1.0) for super-
sonic flight regime (see Fig. 12D) there are five equilibrium points, three of them--'A', 'B',
'C'--are locally stable and two--'D', 'E'--are locally unstable. Three different cross-sections
of the asymptolic stability regions of 'A', 'B', 'C' equilibrium points are presented in
Fig. 12A, B, C. All the cross sections pass through the equilibrium point 'A'. The first
cross-section is placed in the plane of angle of attack ~ and sideslip fl (Fig. 12A), the second
one is placed in Ihe plane of roll rate p and sideslip/~ (Fig. 12B), and the third one is placed
in the plane of yaw rate r and sideslip/~ (Fig. 12C). The region of attraction for equilibrium
'A' is marked wi!h lighter points. One can see, that the more critical level of perturbation is
in yaw rate - r¢r ~ + 0.4 l/s. The boundary of stability region far from equilibrium point
'A' has thin structure with multiple folds.
The size of as)mptotic stability region of equilibrium 'A' becomes significantly less when
the aircraft is tr',mmed at zero or negative angle of attack. To illustrate this feature the
similar cross-sections of the stability regions for (5~=- 0.5° (a= = 0) and (se = 4.4°
(az = - 1.0) are presented in Fig. 13A, B, C and in Fig. 14A, B, C.

40
0
-lo 30
u 20
m 10
o
< -10
u 4
"~ 2
0
~-4
-20
(A)
1O
O
0.5
~ -0.5
-1
0
Sideslip, deg
-6
20 -20 0 20 40
(B) Sideslip, deg
10
¢,3
o-5
-10I
-20 0 20 40 -20 0 20
(C) Sideslip, deg (D) Aileron deflection, deg
Fig. 12. Asymptotic stability region for flight regime with a. = 1.0 (point "A'), M > I. See text for explanation.
2.10. NONLINEAR STALL MOTION
Aircraft high-angle-of-attack excursions are connected with a degradation of lateral
aerodynamic characteristics. The aerodynamic moments become markedly nonlinear func-
tions on angle of attack, sideslip and rotation rate about the velocity vector. They may
significantly change their values resulting in the stability and controllability loss of the
lateral motion.
2.11. APPROXIMATE LINEAR CRITERIA OF STALL
The character of the disturbed lateral motion and its stability under small perturbations
may be evaluated using linearized equations of motion, assuming that the equilibrium angle
of attack is considered as a parameter. To do so, it is necessary to get the dependencies of the
lateral aerodynamic derivatives on angle of attack using data from static, forced oscillation,
rotary balance wind tunnel tests.
Considering the simplified equations for the fast rotational modes of the lateral motion
without taking into account spiral mode and assuming that V = const, ~ = const, C = 0
give the third order characteristic equation:
with coefficients
~.3 + A2}2 + At;'. + At) = 0
A2~
A1 ~_
A 0 =
-- Clp ' ,, + Cn =
tx ", .. i x
-i, c,,cos~-%sin~Iz =-_a~l~
Iz
--" (.:noClp~ -- Cl ('np ~ = -- ¢'7r~ ~
txtz ixiz
(17)

4O
0
•~ 30
o 20
m
t::
== 10
"6
--~ 0
t'-
< -10
O
0.5"10
2
g 0
2
• -0.5
>-
-1
-20 0 20
(A) Sideslip, dog
4
O
2
=-o
2_ 2
0
n'_ 4
-6
(B)
10
O
s
'lO
=- 0
2
o -5
fie
-10
-20 0 20 40
-20 0 20 40
Sideslip, deg
.....~ ' 7 ' _ .....
.. : , ~
i
-20 0 20
(C) Sideslip, dog (D) Aileron deflection, deg
Fig. 13. Asymptotic stability region for flight regime with a= = 0.0 (point 'A'), M > 1. See text for explanation.
40
0
'1o 3O
o 20¢1
== 10
0
• 0
C
-10
1O
O
0.5"ID
d 0
~ -0.5
m
-1
-20 0 20
(A) Sideslip, deg
-20 0 20
(C) Sideslip, deg
~ 4
"~ 2
¢- 0
~-2
0
~-4
-6
10
0
0
~ -S
-10
40
,,~~h'):]t;
-20 0 20
(B) Sideslip, deg
40
5 ...... . . ~ B . ..................:...--......
-20 0 20
(D) Aileron deflection, dog
Fig. 14. Asymptotic stability region for flight regime with a: = - 1.0 (point 'A'), M > 1. See text for explanation.

where ix, iz, are the reduced inertia moments, # is the relative density of aircraft, subscript
f.o. denotes forced oscillations, c~,, c,,~ are derivatives, obtained in the rotary balanced
tests. Since p >> 1, the terms with products of rotational derivatives in the expression for A1
were neglected.
The algebraic stability conditions result in the simple stability criteria
tr/~< 0 ap < 0 a,.~< 0 (18)
R=a,,,+tr~tra>0ortra<---
trti
Breaking of some of these criteria indicates the loss of lateral stability or stall occurring (see
Fig. 15A, B, C). There are various types of the loss in stability. If only one condition a~, is
broken then a real eigenvalue becomes positive (Fig. 15A).These critical points correspond to
the pitchfork bifurcations. For example, in the supercritical case two stable regimes with right
and left rotation appear, while the initial symmetrical regime becomes aperiodically unstable.
If a condition R (Routh criterion) is broken then a complex pair of eigenvalues becomes
unstable (Fig. 15B--Andronov-Hopf bifurcation). This leads to oscillatory unstable
motion provided the frequency of lateral oscillations t-)o ~ ,vf- trp does not become very
small. This is well known case of "wing rock' motion.
The change of sign of ~r~(note the equivalence of a~ with - c,~o,°)first can lead to the
oscillatory instability accompanied with decrease in ~Ooand breakdown of complex pair
into two real roots moving along the real axis. After that a branching point or pitchfork
bifurcation (tr,, = 0) may also occur (Fig. 15C). Taking into account the typical depend-
encies of the aerodynamic derivatives %, c,,,, c~,~at high incidence, such a behavior may take
place in the narrow range of the angle of attack A~ - 1: + T. One may say that this type of
instability is similar to the case when ~r,o> 0.
The criteria tro < 0 and a~ < 0 govern the dynamic stability in yaw. The breaking of the
third criterion try,< 0 indicates the loss of stability in roll or the appearance of the
aerodynamic'autorotation. The departure in yaw is typical of aircraft with swept-like wings
with low aspect ratio. In the great extent the vortex breakdown phenomenon is responsible
for the nonlinear aerodynamics at high angles of attack in this case. Instability in roll occurs
for aircraft with unswept high aspect ratio wings. In the last case the reason for stall is
aerodynamic autorotation of the wing due to asymmetrical flow separation at high angles of
attack (when CL ~ CL..~).
Taking into account the spiral mode and all the omitted aerodynamic effects, the change
in the quantitative results of the linear stability analysis is usually insignificant.
2.11.1. Controllability Criteria
The conventional lateral controls, i.e. ailerons and rudder, at high angles of attack can
hardly be classified as being pure lateral or directional, since aerodynamic moments c~,,and
----O "r Ii
Re
/
ImP.
" R =0
---41--o
Rek
i
I
_ _ l
I
L,n~ i
i
R=O
(A) (B) (C)
Fig. 15. Different kinds of stall through the lateral/directional modes. See text for explanation.

c.~, (i = a, r) are of the same order. That is why the modes of the controlled lateral motion
are closely coupled.
The steady-state value of the rate of rotation about the velocity vector Pw under small
deflections of aileron or rudder 6 can be evaluated from the balance of yaw and roll
aerodynamic moments. This results in the simplified relationship
Pw -~ - 6 (19)
pwb
where/Sw = ~7' tr~ = (c,,cl~ - cz,c,,).
If the perturbed roll motion is stable (ao, < 0), then there will be a 'direct' reaction of the
airplane in the response to the control input provided the condition tr6 < 0 is satisfied.
Nondimensional parameter tr6 is similar to the LCDP (Lateral Control Departure Para-
meter): a6 = LCDP * cz. ~561
This criterion is also coupled with the stall prediction problem, since the reversed
response to the aileron deflection, a~o,changes its sign and may bring about instability in the
closed-loop dynamics.
2.12. STALL DYNAMICS ANALYSIS
The linear analysis determines only the critical values of the angle of attack (or elevator
deflection), when the loss of stability occurs. But the linear theory is unable to predict the
motion after instability occurs. To investigate the motion development due to instability the
analysis of nonlinear equations of motion is needed. The analysis of the equilibrium and
periodic solutions along with their local stability characteristics provides very important
information for nonlinear dynamics prediction.
To illustrate the nonlinear phenomena which are encountered at high angles of attack,
the aerodynamic model for hypothetical fighter-type aircraft (normal scheme with low-
aspect ratio wing and extended strakes) was used. The main aerodynamic characteristics of
lateral/directional motion--c~: c,: c~,w,c.~- defined with spline functions of angle of
attack are shown in Fig. 16A, B. The roll and yaw moment coefficients become nonlinear
functions of sideslip angle and reduced rate of rotation at the high angle of attack region,
where the aerodynamic derivatives change their signs.
The main goal of the aerodynamic design of real aircraft is to provide the stability of
aircraft motion at high angle of attack. But to eliminate bifurcation phenomena completely
is not so easy. The considered hypothetical aerodynamic model gives birth to various
nonlinear phenomena, which are also typical for many real aircraft.
The dependencies ao,, R = a,~ + a#tr# and a~o,presented in Fig. 16C, D, show that Hopf
bifurcations Ht and H2, two pitchfork bifurcations P~ and P2 and the reverse response C1
in aileron control can be expected.
All calculations in this section were performed for hypothetical aircraft using the
fifth order equations of motion (15). The root loci of the linearized equations for the
symmetrical branch of equilibria (the aerodynamic asymmetry is absent and 6o = 0, 6, = 0)
qualitatively confirm the bifurcations predicted by the approximate linear criteria (see
Fig. 16C, D).
Note, that the loss of stability can be with breaking of symmetry and without it. For
example, the Hopf bifurcation, leading to the wing-rock regime from the former stable
symmetrical equilibrium, remains valid for the asymmetry of motion. The pitchfork bifurca-
tions both for equilibrium and oscillatory regime break the symmetry of flight in the vertical
plane. Below some examples of such phenomena are presented.
Autorotational regimes appear in the pitchfork points P~ and P2 on the symmetrical
equilibrium branch, corresponding to the aircraft trim without rotation. The greater the
level of instability between the points P~ and P2, the more prolonged will be the autorota-
tional branches in the positive elevator deflection range. The size and location of the

r-
.d
.'3
-5
X 10-~
..... :~ : / /
0
(A)
20 40 60
0.5
o
"u
-0.5
•-L~
. . . . . . . . . . . . . . . . i . . . . . . . .
0
(B)
20 40 60
=
X 10-~
'7,,
_ -." HI ,PI p2i, "1"12
I
20 40 60
Angle of Attack (deg)
X lO-~'
-5
0 0 20 40
(C) (D) Angle of Attack (deg)
60
Fig. 16. Lateral/directional aerodynamic characteristics. Approximate stall criteria. See text for explanation.
2
', "~ autorotatlonal
1 equilibria
-2
-40 -20 0 20
o-1
rr
-2
'. •f •
I ~- autorotat~onal1" ~-
' ,~ ~-f
H I
I
%.
"- . closed orbits i-
/
i i ~, i
-40 -20 0 20
e~a~ a~mt~ a~ ~vator ae~oa a~
Fig. 17. Autorotational equilibrium and periodic solutions for different elevator deflections•
autorotational branches depend also on the character of the nonlinear aerodynamic terms
Ac~(~,r,/~w) and Ac,(~, r,/Sw).
The equilibrium solutions were computed using (15} and are shown in Fig. 17. They
reveal, along with the symmetrical and autorotational branches generated in the pitchfork
points Pt and P2, some additional autorotational branches, which correspond to the
inertia-coupling regimes (see previous section).
When the autorotational asymmetrical equilibrium branches are stable, the jump-like
departure (undercritical bifurcation) or smooth entering (supercritical bifurcation) to these
regimes will occur at pitchfork points Pt and P2. In the presence of symmetry, there is an
equal probability that aircraft will depart to the left or the right rotation. The side on which
the aircraft departures, depends on the existence of small perturbations when the control
input passes its critical value. In the case under consideration the autorotational branches
are oscillatory unstable. The high angle of attack aircraft behavior is usually ruled by the
sequence of bifurcations.

Bifurcation mcthods in nonlinear flight dynamics 563
Winy Rock motion usually arises in the Hopf bifurcation point, which in our case is
preceded to the first pitchfork point. The Hopf bifurcation points Ht (at ~ ~ 2C) and H2
(at :t ~, 46°) generate the families of the stable closed orbits, which defines the wide range of
the wing rock motion.
In Fig. 17 the amplitudes of the asymmetrical closed orbits, calculated by means of
continuation, starting from point Hi, are presented by dark lines. Really the oscillatory
behavior is much complicated. In Fig. 18 the Poincar~ mapping, defined as the crossing
points of the plane p = 0, is presented for :t and/3 as a function of elevator deflection. The
crossing points at every given elevator deflection are taken after a very long time of
adjustment (t = 200s). Therefore, all the calculated results correspond to the stable solu-
tions. There are some ranges with symmetric periodic solutions with a single and doubled
period. There is a narrow range (fie~ - IC) with chaotic motion. This regime arises from
the symmetrical periodic solutions after a sequence of bifurcations. First the symmetrical
stable closed orbits (see Fig. 19A, B) becomes unstable due to pitchfork bifurcation. After
that, both the unstable symmetrical and stable asymmetrical are subjected to the sequence
of flip bifurcations (see Fig. 19C, D) and as a result a strange attractor appears in accord-
ance with Feigenbaum scheme/~°~ To visualize this 'strange' attractor, four projections of
phase trajectory (p, fl),(p, ~t),(r, fl) and (r, :t), obtained by means of plotting only unconnec-
ted sequential points (20 dots per second), are presented in Fig. 20. Before chaotic behavior
simulation the long time adjustment was fulfilled (t = 2000 s).
Poincar~ mapping reveals also the wide regions with stable asymmetrical closed orbits
(see also Ref. 50). In such cases the aircraft will enter in the oscillatory wing-rock regime
applied on spiral motion.
The more complicated case of the periodic solutions is presented in Fig. 21C, D for
6~ =- 20.5 ~. There are three stable closed orbits (one of them symmetrical and two
asymmetrical) and two unstable closed orbits. In such a case, the final kind of motion will
depend on the initial conditions of state variables, i.e. on the prehistory of control.
40
30
20
(A)
I I I I I I I
-22 -20 -18 -16 - 14 -12 -10
fie (deg)
10
5
0
-5
-10
-15
chaoticmotion
-- ~ ] doubling
, , , ,, , , ,
-22 -20 - 8 - 6 -14 - 2 -I0
(B) 6~.(deg)
Fig. 18. Poincar6 mapping of p = 0 plane for different elevator deflections. See text for explanation.

0.5
-0.5
0.5
-0.5
i i i J i
-10 -5 0 5 10
fl (deg)
i i 0 i i i i i i
-15 15 -3 -20 -10 0 10 20 30
(A) (B) ~t(deg)
0.5
-0.5
i i i i i i i
-10 -5 0 5 10
fl (deg)
0.5
-0.5 '
i i i i
-15 15 -30 -20 -10 0 10 20 30
(C) (D) x(deg)
Fig. 19. Closed orbit projections for different elevator deflections. See text for explanation.
In Fig. 21A, B the family of stable symmetrical closed orbits at 6e =- 40 +- 30.0°,
generated in the Hopf point H2, are presented.
2.13. SPIN DYNAMICS
Analysis of the spin dynamics is carried out using the same general approach as in the
analysis of the roll-coupling problem or the stall instabilities, but considering the eighth
order autonomous system of motion, Eq. (5). The equilibrium solutions of these equations
for the state vector x =. (ct,fl, p, q, r, V, 0, ~b)'~ R8 define the motion along the vertical spiral
trajectories. At high angles of attack and a fast rate of rotation such spirals have small
radius and correspond to the spin modes. Atmospheric density in such a study is usually
held constant, corresponding to considered altitude. The thrust effects only the quantities of
steady-state velocity and radius of spin, but not the qualitative nature of the results,
therefore in many cases it is taken to equal zero.
The searching for all the possible steady states at the given control input
= (fie,6a, ~,, T)' ~ R4 and investigation of their local stability characteristics and regions
of attraction are necessary for construction of the global phase portrait. The solutions of the
approximate spin equations, defining the balances in all forces and moments,C57~can be used
as the initial values for equilibria of the complete nonlinear system (5). This is the better way
to search for the multiple equilibrium points at the given control input. The continuation of
each of these points can give the entire set of solutions.
The bifurcational analysis reveals the qualitative changes of the global phase portrait of
equations of motion. These changes are closely connected with the stability loss and
different kinds of dangerous aircraft behavior. But the real transient dynamics to a great
extent will depend on the character of the control input history in time. Therefore the

1.2
0.8
0.4
o
~- -0.4
-0.8
-1.2
1.2
D.8
3.4
3.4
0.8
1.2
' 'o ' ' ' 'o ' 'o ' ' '-15 --1 -5 0 5 1 15 2 30
# (deg) ~ (deg)
0.75 15
0.50 ~ 10
0.25 5
o
:,---0.25 "~- -5
-0.50 - 10
-0.75 -15
' 0 ' ' ' l'0 ' 2' '-15 --I -5 0 5 15 0 30
(A) # (deg) (B) :~(deg)
Fig. 20. Chaotic behavior due to the 'strange' attractor at 6~ = - 16.5°. See text for explanation.
2.0
1.0
0
-I.0
-2.0
(A)
I I I I I
-20 -10 0 10 20
13(deg)
"8
,....
2.0
1.0
0
-I.0
-2.0
(B)
I
10
((! I
20 30
(deg)
I
40
0.5
o
e.. -0.5
- 1.0
symmetrical
stableorbit
stableorbits
1.0
0.5
0
-0.5
-I.0
1 I I I I I I I I
-10 -5 0 5 10 10 20 30 40
(C) 13(deg) (D) ot (deg)
Fig. 21. Family of closed orbits at fie = - 40 + 20.0° (A, B). Three-stable periodical motion case at fie = - 20.5~
(C, D).

numerical simulation needs to be implemented to check the bifurcation analysis results. It is
especially important for multidimensional spin dynamics due to the large differences in the
time scales of translational and rotational modes of motion.
2.14. BIFURCATIONAL ANALYSIS OF SPIN MOTION
Hereafter some examples of application of continuation technique and bifurcation
analysis for spin dynamics analysis will be presented.
Spiral instability of lateral/directional motion considering the eighth order equations of
motion is qualitatively similar to the rotational instability in roll, when one positive
eigenvalue appears. This case also has the pitchfork bifurcation points giving birth to the
bothside mirror-symmetrical solutions. The symmetrical equilibrium conditions with zero
bank angle are unstable, while the two steady states, representing the right and left spirals
are stable. The main difference is that spiral instability is very weak and asymmetrical stable
spirals have a very small rate of rotation and sideslip, but very large variations in the bank
and pitch angles. If the symmetrical trim position becomes, for example, due to insufficient
dihedral effect, unstable, the apparent right and left spiral solutions with 0 ~- 60:,
~ + 40° are stableJ46'52.53~Bank-angle development due to spiral instability is very slow
and could be easily controlled by a pilot.
Steady state spin modes or equilibrium solutions of motion equations at high angles of
attack, computed along with their local stability characteristics for entire ranges of control
surface deflections, are the basic information for the global dynamics analysis.
In Fig. 22 the general example of equilibrium surface obtained by the continuation
method is presented. These results are taken from Ref. 43 and corresponds to the F-4
aircraft. It is clearly seen that aircraft at given control surface deflections have multiple
equilibrium solutions. The bifurcation diagram shows the number of equilibria.
In Ref. 53 the F-14 dynamics have been studied by determining the steady states of the
eighth order equations of motion and seeking bifurcations. In Fig. 23 the steady states of
F-14 as a function of aileron deflection for an elevator deflection of - 10~and zero rudder
deflection are presented. The multiple steady states exist for the entire range of aileron
deflections.
For example, at zero aileron deflection there are five steady states, three of them are
stable. Expect for the stable trim at low angle of attack without any rotation, aircraft can
enter to the left or right flat spin mode at ~ ~ 80~ with large steady-state yaw rate.
The steep spin modes in the intermediate angle of attack range are all unstable. The stable
branch of aircraft equilibria at low angle of attack exists only in the limited range of aileron
deflections, which is bounded by two saddle-node bifurcation points. Exceeding these
critical aileron deflections will be followed by jumps or departures to the flat spin modes.
The results for critical control deflections, predicted by steady-state analysis, and similar
results obtained in numerical simulation of aircraft manuever, can be different. This
difference usually arises due to the transient effects and the limited sizes of the domains of
attraction for different stable steady states at the considered fixed control deflections.
The calculated steady-state spin modes, as shown above, can be used for departure
prediction and spin recovery analysis. For example, the F-14 steady-state flat spin modes
exist for all aileron deflections; there are some ranges of aileron deflections, where these
spin modes are oscillatory unstable. But in these ranges the steady oscillatory spin motion
is developed and aircraft remains at high angles of attack. Therefore, only by means of
aileron control the spin recovery of the F-14 is very difficult; the rudder control also
cannot improve recovery conditions due to the lack of rudder efficiency at high angle of
attack.
The bifurcation and stability analysis of High Incidence Research Model (HIRM)
dynamics at high angles of attack are presented in Refs 59-61. All calculations in Ref. 59
were performed using a smoothed mathematical model for HI RM aerodynamic character-
istics~62~and KRIT Package, specialized for flight dynamics problems,lall

Bifurcationmethods in nonlinear flightdynamics 567
E(
Fig. 22. F-4 equilibriumsurfaceand bifurcationdiagram.
In Fig. 24 the dependencies for equilibrium c~,/~,0, ~, p, q, r and V on elevator deflection
from Ref. 59 are presented. The results obtained (Fig. 24) show that at zero canard, aileron,
rudder and differential canard deflections, without aerodynamic asymmetry and with zero
thrust there are some critical branches of solutions with nonzero rotation. Three of these
branches (I, 2, 3) in the (~, 6~) plane correspond to spin-like regimes. The branch (4)
corresponds to the roll-inertia rotation regimes with small positive and negative :t. If at the
spin regimes the velocities of descent are small - V ~ 80 - 90 m/s, in the roll-inertia
regimes the velocities are significantly larger --- V = 200 + 300 m/s. Yaw rates in spin
regimes can vary from r ~. 20:/sec for branch 1 to r .~ 60'/sec for branches 2, 3. Bank angle
and sideslip in spin modes are relatively small, but in the case of the roll-inertia rotation
regimes these angles can be essentially greater. Stable equilibria are plotted as solid lines;
divergent (or statically unstable) and oscillatory unstable equilibria using dashed and
dash~lotted lines (original computer-generated pictures are color ones).
The critical equilibria on branches 1, 2, 3 in the most cases are oscillatory unstable. There
are such regimes on the basic symmetric branch A in the range :¢ .~ 20 + 38c at fie ~,
- 12: + - 16~. This symmetric equilibria with zero rotation and zero sideslip generate the
oscillatory motion which leads to wing-rock regime.
It is interesting to note that the oscillatory unstable spin equilibria on branch 2 with
:t ,,~ 60~ at 6~/> -- 10: are the possible source of so-called "agitated' spin modes. Branches
I and 4 are generated in the pitchfork bifurcation points on basic symmetric branch A.
In Fig. 25 the dependencies of equilibrium solutions :~,r, 0 and 4) (for different elevator
deflections - 20, - 30~) on aileron deflection are presented. The first two variables (~ and
r) are the most informative spin characteristics. The equilibrium spin modes demonstrate
the nonlinear dependence on aileron dcflection and have some limit points, which are
potential points of departures and hysteresis-type behavior.
In Fig. 26 the bifurcation diagram for equilibrium solutions is presented in the plane of
control parameters (6o, 6~). Only the solutions with :~ > 20~are considered. The crossing of
the each bifurcation boundary leads to changing of the number of equilibrium points. The

80 i i i
40 ....... ........
o~ 0 • _
~- - 40 " ";~--r~........
- 80 I "<:""'"'1"'~ I
- 40 - 20 0 20 40
6,~ (deg)
16 , , ,
,~ 8 -
-~ o~ -_.
~ -8
-16 i I t
-40 -20 0 20 40
6~ (deg)
16 , , ,
0 ..... "Z'. ""'"--- ~ 4::........
-8 / S
I I '" I
-16
-40 -20 0 20 40
~,, (deg)
180 , , , /
150 " ......
120
90 ~2~Z ..... -~i
-40 -20 0 20 40
6~ (deg)
300
150
0
- 150
-300
- 40
i i ! i t
i-20 0 20 40
6,,(deg)
30
0
-30
-60
-90
-40
I ; i
- 20 0 20 40
6~ (deg)
~5
100
80
60
40
20
- 40
i ! i
~-..~ ~
t ...... ~ ;~.... I
-20 0 20 40
6o(deg)
50 ' ' t
0
-50 I
-40 -20 0 20 40
fi~ (deg)
Fig. 23. F-14 equilibrium solutions of eighth order equations of motion.
difference always equals two. Such results give to the researcher the valuable knowledge
about critical values of control deflections and possible departures due to the disappearance
of stable equilibrium.
Limit cycles or oscillatory spin modes analysis is also very important for understanding of
the global dynamics behavior. Agitated oscillatory spin dynamics is often encountered in
real flight. Such behavior can be connected with the steady oscillatory spin modes, which
are described by the limit cycles or stable closed orbits in the state space of the considered
system of equations of motions.
The origin of a closed orbit in many cases is due to the Hopf bifurcation encountering.
Therefore the starting point and the plane of closed orbit is well known. The oscillatory
unstable equilibrium solutions usually co-exist with stable and unstable periodic solutions
or closed orbits. The computation of the closed orbits and its local stability analysis at
different values of parameters can be performed also using continuation technique in
a similar way as the equilibrium analysis, but with a single difference---computation is more
time-consuming.
In Ref. 50 entire families of equilibria and limit cycle solutions as control surface
deflections and control system parameters have been analysed for the F-15 fighter aircraft.
For continuation and stability analysis of equilibrium and periodical solutions (limit cycles)
of the equations of motion the general-purpose software package AUTO was used.

0
=-
I
Ill>.
100
50
0
-50
-100
-60
-,~ ..... ~-- ~,~... ~,.
_~.~ .'~..../!- 4.
-40 -20 0 20
50
"~ 0
c
-50
ill.
-100 ' ' '
40 -60 -40 -20 0 20 40
•v 50
0
O
-50r-
b=o. ---':-

"x~." ~•4~ " "
. ~ .. t.1
• . . . . .
, = "1
-60 -40 -20 0 20
ta 5o
'ID
o
n- -50 f ' *~.
40 -60 -40 -20 0 20 40
0
"0
o
n,.
100
50
0
-50
-100
-60
i
.
.....,.,k.-,'-.~-- • ".<.
.~ ..... /. ....__~" "-~
--,--> ,~ -.~ ....% /
......~,.... / ' - ~
I
I,
-40 -20 0 20
Q....
2O
ID)
0
•u 10
t-
O
-10
u) _20i
-60
i
40 -40 40
• • I
.~, ._ f: ...-z.:
"<; ;..:..:1;..:..; ......
/
-20 0 20
t.-
n
40
30
20
10
0
-10
-60
f°
• f
/
.--~'~ ""~' '--__:'~" : .-Z.• ~
I
-40 -20 0 20
Elevator deflection, deg
250
200
E 150
>-
100
• l'Vi
I ll!'N~ '.
.,._~.~_...._ - ~_.._- .X_=...~.~..
. . .....:,. .... r ~ .~.~
50
40 -60 -40 -20 0 20 40
Fig. 24. HIRM equilibrium solutions ofeighth order equations for different elevator deflections and 6a = 0, 6c = 0,
6, = 0, no asymmetry, no thrust, H = 5 km.
Particular attention was given to periodic wing-rock motion both with and without control
system.
In Fig. 27 the bifurcation diagrams for eighth order equations, showing the equilibrium
and periodic solutions for symmetric stabilator variation and rudder variation, are pres-
ented• The possible stable steady state and limit cycle values of angle of attack are marked
by solid lines and black circles, the unstable ones with dotted lines and white cycles. The
circles represent the maximum values taken by angle of attack during the period of steady
oscillations. Branches of equilibria at high angle of attack, representing spiral or spin
motions, are mainly unstable. Small regions of stable fiat-spin modes are bounded by Hopf
bifurcation points, which give rise to oscillatory spin modes. The right and left spin modes
are different due to aerodynamic asymmetry in yaw at high angle of attack (see Fig. 27B).
Note that some apparently-disconnected branches of equilibrium solutions in this case were
detected by means of exceeding real physical limits of the control surface deflections.

I00
0
-100
-2b
.., __..-. ~'~
i
0
fi~(deg)
20
0
- 30
- 60
"------:--:::~...... ~ , . .. ~"

i i
-20 0
?i (deg)
20
75 "- : "
• -._...
50
25
I
-2'0 0
6. (deg)
,.... -- ~,
I
20
40
-~ 0
-40
(A)
.....- "
6 2'0
~. (deg)
100
0
" -100
O
L. "7".:-..-
I I
-20 0
.==•o•
•'== ..= "=-'" •
.4
I
20
6o(deg)
0
"~ -30
-60
:.::::r i ~ ~:::2~
i,..,..., f i °----'*
-2'0 6 2b
6,, (deg)
75
"~ 50
/5
25
...,,,~
.--. ,-
~"" ".... ----4 • -----. " .... "'~
40
o
-40
/'""-........
-2b 2b -io
6.(deg) (B) 3o(deg )
Fig. 25. HIRM equilibrium solutions of eighth order equations of motion for different aileron deflections and
3, = 0, 3, = 0, 3, = - 30, - 20": respectively for A) and B) cases, no asymmetry, no thrust, H = 5 km.
Wing-rock motion onset in the angle of attack range ~ ~ 20 ÷ 30", which is seen from the
bifurcation diagram (Fig. 27), agrees well flight test data.~Sm Similar to the equilibrium
pitchfork bifurcation of symmetrical solution the symmetrical periodic solution can also
bifurcate in the pitchfork manner. After such bifurcation of stable limit cycle two asymmet-
rical limit cycles (mirror-images of one other) appear, while the symmetrical periodic
solution becomes unstable. The example of such bifurcation, arising on the wing-rock
branch of periodic solutions, is presented in Fig. 28.
Extended analysis of oscillatory spin modes was performed in Ref. 59 for High Incidence
Research Model (HIRM). In Fig. 29 the equilibrium solutions for ~, r, 0 and ~bat 6c = - 20°
as functions of elevator deflection at zero lateral control and zero asymmetry, taken from
Ref. 59, are presented. The first family of the closed orbits is located 'around' the oscillatory
unstable equilibrium points corresponding to the branch 2. The maximum and minimum
values of the state parameters for periodic solutions are marked by vertical lines. Thick lines
denote stable closed orbits, thin lines the unstable ones. The closed orbits are stable only

20
-20
-40
"'1 13 [
/ 9 7 //1 7 1 9 ~x~
5 :'~
..... ~ ~; -..z.,.,.. 7 "~"'-~" ~ -- .....
I I I
-20 0 20
~. (deg)
Fig. 26. Bifurcational diagram of HIRM equilibria in the plane of elevator and aileron deflections. Eighth order
equations, no asymmetry, no thrust, ~t > 20°, 6c = 0, 6, = 0, H = 5 km.
100
80
60
40
20
0
-20
(A)
-,m...... ~pin,s
'-'"377'.'~] ..........................................................
WingRock~
/
-30 -25 -20 -15 -I0 -5 0 5 10 15 20
• o • .* .k
.......".:::.,~. ..."
80 - "w': ".................
Spi,~ f
60-
40 " WingRock ............
.:,,..; ....... ---"..:::.,..,~::::............. ..
• , 4,,, -. :
20- .......... _ 1 ..........
6, =-10.19"
0
-70
(B) ~,
-50 -30 -10 10 30 50 70
Fig. 27. F-15 equilibrium and periodic solutions of eighth order equations of motion; A) symmetric stabilator
variation, 6, = 0; B~ rudder variation, fie = - 12.2°. Bifurcation diagram for equilibrium and periodic solutions.
36 30 •
32
28
24
20
16
-1.5
-'28.1
-1.0 -0.5 0.0 0.5 1.0 1.5 6
(A) P (a)
28 - 21.2 22 2
~ 26 • 20.2 "~s 22.1
18.9 ~
24. 18.1
17.6
Flong= 17.2
22
-6 -4 -2 0 2 4
Fig. 28. F-15 wing-rock motion. Periodic orbit projections for varying longitudinal parameter; A) orbits for
open-loop system, B) orbits for closed-loop system.
when their amplitudes are relatively small at fie ~ - 2° - 2° (see Fig. 30A). At 6e > 2° the
closed orbits become unstable (see Fig. 30B) after the flip or period doubling bifurcation
arises. As one can see in Fig. 29, the amplitudes in pitch (Act, A0) of single-period closed
orbits are much larger than the amplitudes in the lateral motion (Ar, Ap, A~, Aft).

0
"lO
=-
Itl
>-
100
50
o
-50
-100
-60
,/
.... ~?-£ - ;
-40 -20 0 20
~ 2O
m
-o 0
d-~ -20
~, -40
-60
a. -80
-60 -40 -20 0 20
~- 50
¢u 0
0
~ -50
e-
inchA--~~os~d orb~
branch A . ..- ~ : :--: : -- - -=-
_. "-..
-60 -40 -20 0 20
Elevator deflection deg
200
-~ lOO
m-
o~ oc
o -100
n-
-200
-60
-.>
..>
-40 -20 0 20
Fig. 29. HIRM equilibrium and periodic solutions of eighth order equations of motion for different elevator
deflections and fir = - 20'~, go = 0, ?i, = 0, no asymmetry, no thrust, H = 5 km.
Figure 31 shows the closed orbits originating at the symmetrical branch of HIRM
equilibrium without rotation when 6~ ~ - 27L At first, oscillatory solutions are stable, then
the Hopf bifurcation occurs (a complex pair of multipliers crosses the unit circle), after that
pitchfork bifurcation leads to the appearance of two asymmetrical closed orbits.
The following shows some examples of complicated chaotic behavior. When any simple
stable attractor, equilibrium or periodic, is absent at the given control setting the chaotic
dynamics can be realized due to the appearance of a strange attractor. In Fig. 32 the
examples of such behavior are presented (these results are taken from Ref. 50).
The first case (Fig. 32B) indicates the presence of toroidal attractor (to reveal the toroidai
structure of the phase space in the bottom figure only the discrete data points along the
trajectory are plotted). Note that amplitude state variables in this case are very large
- :<,~ 40° + 90~, r ~ 1.0 + 3.0 rad/sec. It is clear that toroidal structure is based on the
transitions from the oscillatory fiat-spin mode to the highly oscillatory steep-spin mode,
which are both unstable. The time of such transitions is much more than one period of
angular oscillations.
The second example (Fig. 32A) shows a global attractor with the sequence of transitions
between spin mode and wing-rock regime, which arises just beyond the region of stable
wing rock.
In Fig. 33 the example of complicated oscillatory behavior in the fiat-spin mode, taken
from Ref. 47, is presented. The transient dynamics possess two time scales of parameter
variations--short and long periods. The first one is about 3 s, the second is much greater,
about 150 s. Short period oscillation can increase and decrease the amplitude (see Fig. 33A).
Therefore the short time dynamics depends on the initial conditions of motion. The phase
trajectory projections on (p,/~), (r,/~) planes (Fig. 33B) and Poincar+ mapping (Fig. 33C)
clearly reveal the qualitative nature of such dynamics. There exists the toroidal stable
manifold of trajectories, and phase trajectories from its domain of attraction go to this
invariant manifold. Inside the toroidal attractor there is an oscillatory unstable closed orbit
(Fig. 33C).
Departure characteristics quant!/ication can be done using the results of equilibria and
closed orbits computation along with their stability analysis.
The first type of departures is connected with the bifurcations, leading to the disappear-
ance of stable equilibria (e.g. saddle-node and undercritical pitchfork bifurcations). There

1.5
~m
= 0.50
r~
0.g
~" O.8
0.7
(A)
1.5
"10
o 0.s
re"
-2 0
Sideslip, deg
-2 0
Sideslip, deg
0.4
0.2
0
-0.2
-0,4
2 30
9.
1
2 3O
O.G
0.4
0.2
0
-0.2
-0.4
O
40 50 60
Q vs AOA, deg
40 50 60
AOA, deg
2
1
_E
-I
-2
-2
6
5
3
2
-2
2
1
-1
-2
-2
Flip bifurcatio,n
0
Real
IIIWXNItXNNxxxXXN~I
ItU11 "~
?Plip bitur~tion "
0 2 4 6 8
Elevator, deg
o
o
-2 0 2 40 60 80 0
Sideslip, deg Q vs AOA, deg Real
, 1 t
,,4 5 . . . . . ' x ~ ' ~ . . . . . . .
;0, /
>. 0.8~ ;3 -t
-2
07 t -3 2
-2 0 2 40 60 80 10 15 20
(B) Sideslip, deg AOA, deg Elevator, deg
Fig. 30. HIRM clo:~ed orbit projections with multiplies for different elevator deflections and 6, = - 20, rio = 0,
6, = 0, no asymmetry, no thrust, H = 5 km; A) stable orbits, B) unstable orbits after flip bifurcation.
may be also the bifurcation points, where the stable closed orbits disappear. In these cases
the transient motion transfers an aircraft to another stable equilibrium or oscillatory
motion (or to a more complex attractor). When there are some different attractors, the
motion behavior following the bifurcation encountering depends on the initial location in
the state space of the bifurcated solution. The behavior is determined by the attractor, in
which region of attraction the bifurcated point is placed.
The character of aircraft dynamics when parameters cross the bifurcation boundary
depends on the rate of parameter variation. When parameters vary slowly, the aircraft
dynamics can be predicted more correctly, because it is close to the equilibrium values. In
the case of fast variation of parameters the difference between the transient motion and
steady-state solutions will be more considerable. Nevertheless, to quantify the departure
characteristics one can use the bifurcation diagrams in the parameter planes, which define
critical conditions of stable solutions disappearing. The influence of a parameter variation

u~
"0
¢-
12.
0.05
-0.05
O
"10
Q.
(/3
.'g_
(/)
10
5
0
-5
-10
1
-o 0.6
o -0.5CC
-1
32 34 36 0 -5 0 5 10
Angle of attack, deg Sideslip, deg
0.4
¢o
0.2
0
-0.2
>-
-0.4
-1032 34 36 -5 0 5
Angle of attack, deg Sideslip, deg
10
Fig. 31. H1RM closed orbit projectionsfor differentelevator deflectionsand 6, = - 20~, 6, = 0, 6, = 0, no
asymmetry,no thrust, H = 5 km. Pitchforkbifurcation.
rate on the character of 'jumps' or departures can be evaluated only by means of numerical
simulation of aircraft motion.
Figure 34 shows the example of HIRM transient dynamics under the elevator variation
with dfJdt ,~- 2.0~/sec. Aircraft motion starts from 'agitated' spin mode (branch 2)
:~~ 60°, r ~ 50°/sec. The oscillatory motion at the bifurcation point fie ~ - 4: is followed
by a 'jump' to the equilibria on the symmetrical branch A without rotation. This 'jump' is
the result of the disappearance of equilibria and closed orbits on the critical branch 2, but
due to parameter variation it is significantly delayed with respect to the equilibrium
analysis. Numerical simulation of the 'agitated' spin modes confirms the amplitudes of the
calculated closed orbits around the branch 2. The further variation of elevator results in
the movement of state point along the symmetric branch A up to the next departure at
6e ,~ - 23°.
The second type of departures is connected with the loss of stability under the application
of finite (not small) perturbations. In many cases the stable equilibria or closed orbits
have the limited asymptotic stability regions; and there exist critical values of state
variable disturbances, the exceeding of which leads to loss of stability and transfer or
'jump' to another attractor, which can be realized at the same value of parameters. The
nearer the equilibrium to bifurcation point, the smaller the size of its asymptotic stability
region. To quantify the departure resistance in such cases, the characteristic size of stability
region with respect to the existing level of perturbations may be used as a measure of
quantification.
Closed-loop dynamics analysis for control-law design purposes is a very important
application of bifurcation analysis. Feedback control laws can significantly change the
closed-loop dynamics and eliminate bifurcation phenomena, leading to stability loss and
departures. For example, the wing rock and spiral divergence instabilities could be control-
led with a simple feedback control system. In Ref. 51 such results are presented. Roll rate
and sideslip feedbacks to the ailerons can be used to supplement the aerodynamic deriva-
tives, which are responsible for instability, i.e. damping in roll or dihedral derivative. In such
a way the instability regions of steady states can be eliminated. Continuation methods make
it possible to determine the effects of the control system by means of the closed-loop
dynamics of a nonlinear system stability analysis.

1oo
80
6o
40
20
0
4.0
2.0-
l,
-0.0 -
-2•0
0
100
80
6O
40
20
0
-2.0
t
J
, 1 i
200 400 600
t
100.
90'
80'
70.
60'
50'
40.
30'
20'
80O 0 2OO
4.0 '
40O
t
600
575
800
i
200 400 600 800
t
• • .~.~.~....
• . . . . ~
;,: :'
"."~.. ~.. ~r:' " ':'
:..; ~. ,/-:" ...:.
'.. ". -'" "'~d" • . .."
~ '.'..~= "- o ~- ....
°
3.0 '
,- 2.0'
1•0 '
0.0
0
100
i
200
i i
400 6O0
t
90
80
70
60
50
40
30
20
0.0
800
.°°., °,..,.•
• . •.... . o°,~,
.:.;. :. -. :,.:-;..-...._.,
•~:~ .,:.~4~..:~...:, ,-
• t.ol. • , o.
~' O; ". •
i | | '
-1.0 0.0 1.0 2.0 3.0 4.0 1.0 2.0 3.0
r r
(A) (B)
4.0
Fig. 32. Chaotic attractors in spin dynamics of the F-15. See text for explanation•
In Ref. 50 it has been shown that the geometrical nature of the nonlinear results can be
used for high angle of attack control-law design applications, for example, for wing rock
suppression, departure elimination, spin avoidance/recovery. Simplified lateral-directional
feedback strategies for the F-15 aircraft were investigated using continuation and bifurca-
tion analysis of the closed orbit solutions. These strategies comprised the conventional roll
damping augmentation by roll control and a roll rate 'crossfeed' to rudder, scheduled by
angle of attack. Some adjustable gain parameters were searched for to obtain the better high
angle of attack behavior, i.e. to reduce as much as possible the amplitudes or eliminate the
wing rock altogether•
The bifurcation analysis results can be used for design of control surface interconnec-
tion.(43"45"46"53)For example, the saddle-node points at low angles of attack equilibrium
branch are responsible for entering spin (Fig. 23). Therefore, the boundaries in the control

- 0.0
.~ -0.05
-0.1
~ tad -0.!
Poincar~ mapping
oscillatory
unstable
closed orbit
(c)
-0.05
saddle ,omt
unstable acus
Stable invariant toroidal manifold
0.05 ..... ~............. : :..... :..... !.... ""
1.5
.~ 1.4
~ 1.3
~ 1.2
~ 1.1
1.0
-i.1 -1.00
-i.3 -1.25
~ -1.4
-1.5 ~, -I.50
0 40 80 120
(A) Time. sec
160
-I.I
-1.3
-1.4
r, I/s
"~ 1.50
1.25
o
~0
~. 1.00
~ 0.6
" 0.4
-- 0.1
~ 0.0
-0.1 -0.05 0 0.05
(n) Sideslip angle, tad
oscillatory
unstable
closed orbit
Fig. 33. Tordoial attractor in the 'agitated" flat spin mode; A) Transient time histories, B) phase trajectory
projections, C) Poincar~ mapping. See text for more detailed explanation.
plane (6e,6a, 6,), formed by these critical limit points, can be used for spin prevention. These
boundaries can be taken into account in the control surface interconnection for putting
limits on the control surface deflections to prevent jump phenomena.
3. THEORETICAL BACKGROUND
Hereafter some concepts and methods of the dynamical systems theory, which are more
often used for flight dynamics applications, are presented. ~3-5'27'10~
The nonlinear dynamical system, depending on parameters, will be considered in the
following form:
dx- - =
dt F(x,c), x~R",cEMcR m (20)

¢=
'10
=-
all
]=
¢l
>.
100 - -
50
0
-50
-100 - -
-60 -40 -20 0 20 40
0 -- -
nl
-so
-100
-60 -40 -20 0 20 40
0
-v 50
~ 0
o
o
-50e-
branch 2
branch A
-60 -40 -20 0 20
100
5O
"o
" 0r-
eO
~- -50
-100
40 -60 -40 -20 0 20 40
100
} so
$ 0
~ -50
-100 L
-60
_ -
-~i' / /
, ~ ,
-40 -20 0 20
k
2O
ii0Io-10
-20
40 -60 -40 -20 0 20 40
et
o
"o
°
m
J~
0.
40
30
20
10
0
-10
-60
$ /
<P
** / t
iI" oo • • o-o- ~
-40 -20 0 20
250
200
E 150
100
I
L
I -
50
40 -60 -40 -20 0 20 40
d3~
Fig. 34. Transient H1RM dynamics under elevator variation with ~- ~ - 2.0~/sec in comparison with equilib-
rium solutions (3, = - 20', 30 = 3, = 0, H = 5 km).
where F is a smooth vector function. The vector field F defines a map R"÷" ---,R'. The
system (20) satisfies the conditions of the existence and uniqueness of solutions x(t, Xo) with
initial condition x(0, Xo) = Xo. The solution curve ~o,(Xo)= x(t, Xo) is called the trajectory
and together wtth other trajectories forms the flow of the dynamical system. The set of all
trajectories constitute the phase portrait of the dynamical system.
The concept of the critical elements of a nonlinear autonomous dynamical system is the
basic one. The critical elements are invariant manifolds of trajectories, the dimensions of
which vary from 0 up to n - 1. The simplest ones are well-known equilibrium points or
zeroes of a vector field and closed orbits, which define the periodic solutions. They may be
attractive for other trajectories, if they are stable. There are also more complex attractors in
high order dynamical system (n/> 3). For example, invariant toroidal manifolds of trajecto-
ries and 'strange' attractors (see Fig. 35).

Nonlinear flight dynamics problems analyzed using bifurcation methods

Nonlinear flight dynamics problems analyzed using bifurcation methods

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