Three-Dimensional Trajectory Optimization in Constrained Airspace
Except where reference is made to the work of others, the work described in this
dissertation is my own or was done in collaboration with my advisory committee.
This dissertation does not include proprietary or classified information.
Ran Dai
Certificate of Approval:
David Cicci
Professor
Aerospace Engineering
John E. Cochran Jr. Chair
Professor
Aerospace Engineering
Roy Hartfield
Professor
Aerospace Engineering
Anwar Ahmed
Associate Professor
Aerospace Engineering
George T. Flowers
Graduate School, Interim Dean
Three-Dimensional Trajectory Optimization in Constrained Airspace
Ran Dai
A Dissertation
Submitted to
the Graduate Faculty of
Auburn University
in Partial Fulfillment of the
Requirements for the
Degree of
Doctor of Philosophy
Auburn, Alabama
December 17, 2007
Three-Dimensional Trajectory Optimization in Constrained Airspace
Ran Dai
Permission is granted to Auburn University to make copies of this dissertation at its
discretion, upon the request of individuals or institutions and at
their expense. The author reserves all publication rights.
Signature of Author
Date of Graduation
iii
Vita
Ran Dai, the daughter of Degang Dai and Hongdi Qiao was born in Taizhou,
Jiangsu Province, P.R.China. She graduated from Taizhou High School in June,
1998 and entered school of Automation Science and Electrical Engineering, Beijing
University of Aeronautics and Astronautics in September, 1998. She graduated in
June, 2002 with a Bachelor of Automation Science. From January, 2004, she started
her graduate study in Auburn University in the Department of Aerospace Engineering
under the supervision of Dr. John, E. Cochran Jr. and received her master degree of
Aerospace Engineering in August, 2005.
iv
Dissertation Abstract
Three-Dimensional Trajectory Optimization in Constrained Airspace
Ran Dai
Doctor of Philosophy, December 17, 2007
(MAE, Auburn University, 2005)
(B.S., Beijing University of Aeronautics and Astronautics, 2002)
162 Typed Pages
Directed by John E. Cochran Jr.
This dissertation deals with the generation of three-dimensional optimized tra-
jectory in constrained airspace. It expands the previously used two-dimensional air-
craft model to a three-dimensional model and includes the consideration of complex
airspace constraints not included in previous trajectory optimization studies. Two
major branches of optimization methods, indirect and direct methods, are introduced
and compared. Both of the methods are applied to solve a two-dimensional minimum-
time-to-climb (MTTC) problem. The solution procedure is described in detail. Two
traditional problems, the Brachistochrone problem and Zermelo?s problem, are solved
using the direct collocation and nonlinear programming method. Because analytical
solutions to these problems are known. These solutions provide verification of the nu-
merical methods. Three discretization methods, trapezoidal, Hermite-Simpson and
v
Chebyshev Pseudospectral (CP) are introduced and applied to solve the Brachis-
tochrone problem. The solutions obtained using these discretization methods are
compared with the analytical results.
An 3-D aircraft model with six state variables and two control variables are
presented. Two primary trajectory optimization problems are considered using this
model in the dissertation. One is to assume that the aircraft climbs up from sea level
to a desired altitude in a square cross section cylinder of arbitrary height. Another
is to intercept a constant velocity, constant altitude target in minimum time starting
from sea level. Results of the optimal trajectories are compared with the results
from the proportional navigation guidance law. Field of View constraint is finally
considered in this interception problem.
The CP discretization and nonlinear programming method is shown to have
advantages over indirect methods in solving three-dimensional (3-D) trajectory opti-
mization problems with multiple controls and complex constraints.
Conclusions from both problems are presented and properties of each one are
discussed. Finally, suggestions for future research are addressed.
vi
Acknowledgments
First I would like to express my great appreciation to my advisor, Dr. John E.
Cochran Jr., for his instruction, guidance, encouragement and patience during my
graduate study in Auburn. His invaluable advice and working attitude benefit me a
lot in my research and will continue effect my future career.
Many thanks to Dr. David Cicci, Dr. Roy Hartfield, Dr. Anwar Ahmed, Dr.
Andrew Sinclair and Dr. A. Scottedward Hodel for their academic guidance, time
and assistance in my graduate courses. The knowledge and skills I studied from their
courses made good basis for my dissertation work.
I feel lucky to take classes together with Oluseyi Onawola, Dakshesh Patel, Ravi
Duggirala and Justin Martin and developed wonderful friendship with them. I also
want to thank my friends Qi Hang, Yawei Han, Hongxia Zhang, Ran Zhou, Rui
Chao, Wei Huang and Zhiyang Ding to accompany me through the time in Auburn
and always offer me general help.
Finally I would like to express my gratitude to my family, their support and faith
in me encouraged me throughout the completion of this degree .
vii
Style manual or journal used Journal of Approximation Theory (together with
the style known as ?auphd?). Bibliography follows Journal of Aircraft
Computer software used The document preparation package TEX (specifically
LATEX) together with the departmental style-file auphd.sty.
viii
Table of Contents
List of Figures xi
List of Tables xiv
Acronym xv
1 Introduction 1
1.1 Background . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1
1.2 Statement of Problem and Approach . . . . . . . . . . . . . . . . . . 3
1.3 Dissertation Outline . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
2 Literature Review 7
2.1 Methods in Solving MTTC Problem . . . . . . . . . . . . . . . . . . . 7
2.1.1 Gradient Method . . . . . . . . . . . . . . . . . . . . . . . . . 7
2.1.2 Energy State Method . . . . . . . . . . . . . . . . . . . . . . . 8
2.1.3 Singular Perturbation Method . . . . . . . . . . . . . . . . . . 9
2.1.4 Modified Sweep Method . . . . . . . . . . . . . . . . . . . . . 10
2.1.5 Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
2.2 Methods in Solving MTI Problem . . . . . . . . . . . . . . . . . . . . 12
2.3 DCNLP Application Examples . . . . . . . . . . . . . . . . . . . . . . 15
2.3.1 Aircraft Optimal Trajectory Design . . . . . . . . . . . . . . . 15
2.3.2 Space Mission Planning . . . . . . . . . . . . . . . . . . . . . 17
2.3.3 Unmanned Aerial Vehicles (UAV) Optimal Path Planning . . 18
2.3.4 Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19
3 Indirect Optimal Control Method 20
3.1 Optimal Control Problem . . . . . . . . . . . . . . . . . . . . . . . . 20
3.2 Hamiltonian Equation and Necessary Conditions . . . . . . . . . . . . 21
3.3 Path Constraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
4 The Method of Direct Collocation and Nonlinear Programming 28
4.1 Direct Collocation Method . . . . . . . . . . . . . . . . . . . . . . . . 28
4.1.1 Trapezoidal Method . . . . . . . . . . . . . . . . . . . . . . . 30
4.1.2 Hermit-Simpson Method . . . . . . . . . . . . . . . . . . . . . 32
4.1.3 Chebyshev Pseudospectral Method . . . . . . . . . . . . . . . 36
ix
4.2 Nonlinear Programming Solver . . . . . . . . . . . . . . . . . . . . . . 40
4.3 Mistake Prevention . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45
5 Using DCNLP To Solve Optimal Control Problems 47
5.1 Brachistochrone Problem . . . . . . . . . . . . . . . . . . . . . . . . . 47
5.2 Zermelo?s Problem . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53
6 Constrained Airspace Minimum Time and Fuel Flight Path 56
6.1 2-D MTTC Problem . . . . . . . . . . . . . . . . . . . . . . . . . . . 56
6.1.1 2-D Mathematical Model . . . . . . . . . . . . . . . . . . . . . 56
6.1.2 Energy State Method . . . . . . . . . . . . . . . . . . . . . . . 61
6.1.3 Gradient Method . . . . . . . . . . . . . . . . . . . . . . . . . 63
6.1.4 DCNLP Method . . . . . . . . . . . . . . . . . . . . . . . . . 65
6.2 3-D MTTC Problem . . . . . . . . . . . . . . . . . . . . . . . . . . . 69
6.2.1 3-D Mathematical Model . . . . . . . . . . . . . . . . . . . . . 69
6.2.2 Initial NLP Variable Inputs . . . . . . . . . . . . . . . . . . . 71
6.2.3 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74
6.2.4 Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76
6.3 3-D Minimum-Fuel-To-Climb Problem . . . . . . . . . . . . . . . . . 78
7 Other Constrained Optimal Trajectories 99
7.1 Three-Dimensional Proportional Navigation Guidance Law . . . . . . 99
7.2 Three-Dimensional Minimum-Time Interception Trajectory Planning 103
7.3 Three-Dimensional Minimum-Time Rendezvous Trajectory Planning . 106
7.4 View Constraints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 119
7.5 Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 120
8 Summary and Recommendations 124
Bibliography 126
A Aircraft Propulsion and Aerodynamic Data 130
B Gradient Method for 2-D MTTC problem 131
C Proportional Navigation Guidance Law Algorithm 134
D Chebyshev Pseudospectral Collocation and Nonlinear Pro-
gramming Codes in Solving Brachistochrone Problem 140
x
List of Figures
4.1 Trajectory History Discretization . . . . . . . . . . . . . . . . . . . . 29
4.2 Trapezoidal Discretization . . . . . . . . . . . . . . . . . . . . . . . . 31
5.1 Fermat?s Principle . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48
5.2 Analytical and DCNLP Results for Brachistochrone Problem . . . . . 52
5.3 Zermelo?s Problem . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54
6.1 2-D Aircraft Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57
6.2 2-D MTTC Trajectory Using Energy State Model . . . . . . . . . . . 62
6.3 2-D MTTC Trajectory Using Gradient Method . . . . . . . . . . . . 65
6.4 2-D MTTC State and Control Variables History Using Gradient Method 66
6.5 2-D MTTC Trajectory Using DCNLP Method . . . . . . . . . . . . . 67
6.6 2-D MTTC State and Control Variables History Using DCNLP Method 68
6.7 2-D MTTC Altitude and Down Range Trajectory Using DCNLP Method 69
6.8 3-D Aircraft Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . 71
6.9 Circular Helix Curve Wrapped on A Cylinder . . . . . . . . . . . . . 72
6.10 3-D MTTC Trajectory for Case 1 . . . . . . . . . . . . . . . . . . . . 81
6.11 Control and State Variables History for Case 1 . . . . . . . . . . . . . 82
6.12 3-D MTTC Trajectory for Case 2 . . . . . . . . . . . . . . . . . . . . 83
6.13 Control and State Variables History for Case 2 . . . . . . . . . . . . . 84
xi
6.14 3-D MTTC Trajectory for Case 3 . . . . . . . . . . . . . . . . . . . . 85
6.15 Control and State Variables History for Case 3 . . . . . . . . . . . . . 86
6.16 3-D MTTC Trajectory for Case 4 . . . . . . . . . . . . . . . . . . . . 87
6.17 Control and State Variables History for Case 4 . . . . . . . . . . . . . 88
6.18 3-D MTTC Trajectory for Case 5 . . . . . . . . . . . . . . . . . . . . 89
6.19 Control and State Variables History for Case 5 . . . . . . . . . . . . . 90
6.20 3-D MFTC Trajectory for Case 6 . . . . . . . . . . . . . . . . . . . . 91
6.21 Control and State Variables History for Case 6 . . . . . . . . . . . . . 92
6.22 3-D MFTC Trajectory for Case 7 . . . . . . . . . . . . . . . . . . . . 93
6.23 Control and State Variables History for Case 7 . . . . . . . . . . . . . 94
6.24 3-D MFTC Trajectory for Case 8 . . . . . . . . . . . . . . . . . . . . 95
6.25 Control and State Variables History for Case 8 . . . . . . . . . . . . . 96
6.26 3-D MFTC Trajectory for Case 9 . . . . . . . . . . . . . . . . . . . . 97
6.27 Control and State Variables History for Case 9 . . . . . . . . . . . . . 98
7.1 Engagement Geometry . . . . . . . . . . . . . . . . . . . . . . . . . . 101
7.2 3-D MTI Trajectory using DCNLP for Case 1 . . . . . . . . . . . . . 108
7.3 3-D MTI Trajectory using PNG for Case 1 . . . . . . . . . . . . . . . 108
7.4 Control Factors and Velocity History for Case 1 . . . . . . . . . . . . 109
7.5 3-D MTI Trajectory using DCNLP for Case 2 . . . . . . . . . . . . . 110
7.6 3-D MTI Trajectory using PNG for Case 2 . . . . . . . . . . . . . . . 110
7.7 Control Factors and Velocity History for Case 2 . . . . . . . . . . . . 111
7.8 3-D MTI Trajectory using DCNLP for Case 3 . . . . . . . . . . . . . 112
xii
7.9 3-D MTI Trajectory using PNG for Case 3 . . . . . . . . . . . . . . . 112
7.10 Control Factors and Velocity History for Case 3 . . . . . . . . . . . . 113
7.11 3-D MTI Trajectory using DCNLP for Case 4 . . . . . . . . . . . . . 114
7.12 3-D MTI Trajectory using PNG for Case 4 . . . . . . . . . . . . . . . 114
7.13 Control Factors and Velocity History for Case 4 . . . . . . . . . . . . 115
7.14 3-D MTR Trajectory using DCNLP for Case 5 . . . . . . . . . . . . . 116
7.15 3-D MTR Trajectory using PNG for Case 6 . . . . . . . . . . . . . . 116
7.16 Control Factors and State Variables History for Case 5 . . . . . . . . 117
7.17 Control Factors and State Variables History for Case 6 . . . . . . . . 118
7.18 Control Factors and State Variables History for Case 7 . . . . . . . . 120
7.19 3-D MTI Trajectory using DCNLP for Case 7 . . . . . . . . . . . . . 121
7.20 Control Factors and State Variables History for Case 7 . . . . . . . . 122
xiii
List of Tables
4.1 Comparison of Three Discretization Methods . . . . . . . . . . . . . . 39
5.1 Analytical and DCNLP Results for Brachistochrone Problem . . . . . 51
5.2 Comparison of Brachistochrone Results for Various Collocation Methods 53
5.3 Analytical and DCNLP Results for Zermelo?s Problem . . . . . . . . 55
6.1 Polynomial Coefficients at Different March Number . . . . . . . . . . 60
6.2 Boundary Constraints and Performance Limitations for 3-D MTTC
Problem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74
7.1 Boundary Constraints and Performance Limitations for 3-D MTI Prob-
lem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104
A.1 Thrust as a function of altitude and Mach number from Ref.[2] for
aircraft 2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 130
A.2 Lift and drag coefficients as a function of angle of attack and Mach
Number for aircraft 2. . . . . . . . . . . . . . . . . . . . . . . . . . . 130
xiv
Acronyms
Aeroassisted Orbital Transfer Vehicle AOTV
Chebyshev-Gauss-Lobatto CGL
Chebyshev Pseudospectral CP
Direct Collocation and Nonlinear Programming DCNLP
Direct Optimal Control DOC
Field of View Constraint FVC
Indirect Optimal Control IOC
Karush-Kuhn-Tucker KKT
Minimum-Time-To-Climb MTTC
Minimum-Fuel-To-Climb MFTC
Minimum-Time Interception MTI
Minimum-Time Rendezvous MTR
Nonlinear Programming Problems NLPs
optimal control problem OCP
Proportional Navigation PN
Sequential Quadratic Programming SQP
Two-Dimensional 2-D
Three-Dimensional 3-D
Two-Point-Boundary-Value-Problems TPBVPs
Unmanned Aerial Vehicles UAV
xv
Chapter 1
Introduction
1.1 Background
Over the last 50 years, the trajectory optimization problem has attracted the
interests of many researchers. Due to the development of high speed digital comput-
ers, the trajectory optimization technology has been widely applied in both military
and civil areas, including space mission planning[24], missile guidance[11], aircraft
transportation systems[17], towed-aerial-cable systems[23] and more. Some of the
researchers have focused their attention on minimum time-to-climb (MTTC) trajec-
tories. Many methods have been used to solve this kind of problem. As exam-
ples, Bryson and Denham[1] used the summarized steepest-ascent method, Calise[3]
applied singular perturbation techniques and Ardema[4] used matched asymptotic
expansions to get approximate analytical solutions to the MTTC problem. These
methods, categorized as indirect optimal control (IOC) methods, are aimed at solv-
ing only two-dimensional (2-D) MTTC problems which satisfy the initial and final
boundary conditions.
Meanwhile, other researchers paid more attention to minimum-time interception
(MTI) problems. Most focused on the onboard generation of trajectories varying from
1
short to long range. For example, Visser, Kelley and Cliff[11] used a singular pertur-
bation method to find an approximate optimal trajectory and then apply neighboring
optimal guidance to transfer an aircraft which has deviated from the reference trajec-
tory to the desired trajectory. Kumar, Seywald and Cliff[12] proposed finding medium
range optimal trajectories for air-to-air missiles using a three-stage guidance scheme.
These methods and some similar ones[13]-[14] are primarily applicable to small de-
viations from the planned path. For large deviations and unplanned interception,
other real-time guidance laws are applied. Proportional navigation guidance (PNG)
is very generally applied. This type guidance firstly studied by Yuan[7] was in two
dimensions and extended to three dimensions by Adler[8], Duflos[9] and Cochran[10].
PNG has the advantage of on-time realization, but its success depends on some ini-
tial conditions. Furthermore, PNG does not consider performance optimization and
reduction of final miss distance.
With the increase in computing power, the use of direct collocation and non-
linear programming (DCNLP) to convert two-point-boundary-value-problems (TP-
BVPs) into nonlinear programming problems (NLPs) the so-called direct optimal
control (DOC) methods have become feasible. Early in 1987, Hargraves and Paris[19]
applied the collocation method to solve a 2-D MTTC problem. Then, in 1993, Betts
and Huffman[20] explained the procedures of the DCNLP methods in more detail,
especially the sparse sequential quadratic programming algorithm. In 1999 Horie and
2
Conway[25] published a paper in which they described how this method was used to
solve minimum time and minimum fuel aeroassisted orbital interception problems.
Quite recently, in 2006, Geiger, Horn and Delullo[27] applied DCNLP to find un-
manned aerial vehicle trajectories that maximize viewing time.
Although the DCNLP has been widely used in aerospace trajectory optimization,
it has had limited application in 3-D MTTC and short range MTI trajectory planning.
These are the focus of this dissertation.
1.2 Statement of Problem and Approach
The operational airspace of aerospace vehicles, including airplanes and unmanned
aerial vehicles, is often restricted. These restrictions may come from local terrain con-
straints, radar coverage constraints, or collision avoidance constraints. Considering
these limitations, aerospace vehicles cannot be assumed to be free to fly anywhere in a
given airspace. Forbidden zones can be defined in 2-D or 3-D spaces. So that climbs,
descents and other maneuvers are all required to be performed in three dimensions.
One of the principal problems considered here is to find optimal trajectories of an
aircraft that is climbing from sea level to a desired altitude in an airspace, consisting
of cylindrical volume of unlimited height. The aircraft motion model used here is a
3-D, point-mass model consisting of six first-order differential system equations. It
is intuitively expected that a MTTC trajectory in a constrained airspace would be
3
similar to the horizontal projection of a 2-D MTTC trajectory with corresponding
initial and final altitudes. Hence, considerable turning will be required if the horizonal
range of the constrained airspace is shorter than the required distance in the 2-D
MTTC results.
If the IOC method is used, the traditional procedure is to write the Hamilto-
nian and then derive the necessary conditions in terms of differential equations for
Lagrangian multipliers and the maximum(minimum) of the Hamiltonian with respect
to the control variables. Because some of the initial conditions are usually unknown,
a method for determining them must be applied. For example, in the ?shooting
method?, a guess is made of to all unknown initial state variables and the Lagrangian
multipliers and the system equations are integrated forward to get final conditions,
which generally deviate from the required boundary conditions. Some adjustment of
the initial guess is then made according to the deviation in the final known boundary
conditions. The above process is repeated until the final boundary conditions are
satisfied within a specified tolerance.
When the DCNLP is applied, the trajectory is discretized into numerous seg-
ments, characterized by state and control variables as parameters. In this way, a
TPBVP is transformed into a problem of determining the parameters that satisfy the
constraints and at the same time maximize or minimize a performance index. The
constraints mentioned here can be in the form of system equations constraints, state
4
and control variables constraints, initial and final boundary constraints, and some
other forms.
In the 3-D MTTC problem, the final objective of the DCNLP method is to
find the parameterized vertical and horizontal load factors as controls and the state
variables while minimizing the performance index: the final time. A simplified model
representing a climbing helical curve similar to the final optimized MTTC trajectory
is used as the initial input of the state and the control variables.
The 3-D MTI problem is to intercept a constant-velocity, constant-altitude target
from sea level. This appears to be quite a different type of problem than the MTTC
problem. But, if the DCNLP method is applied and all the control factors and state
variables are parameterized, the interception point status will be transferred to one
of the final boundary conditions. The interception trajectory also show similarities in
terms of effect of different initial velocities on time consuming comparing with MTTC
trajectory.
1.3 Dissertation Outline
This dissertation is organized as follows. The first part of Chapter 2 contains
a review of the previous research on methods for trajectory optimization, especially
those used on 2-D MTTC problems. These are the energy state method, gradient
method, singular perturbation method and modified sweep method. Then methods
5
used in solving MTI problems will be reviewed. Finally, the DCNLP method is
considered and some previous applications discussed.
Chapter 3 deals with the traditional indirect optimal control method which con-
siders both equality and inequality constraints. Chapter 4 introduces the DCNLP
method in detail including three kinds of discretization methods, Trapezoidal, Hermit-
Simpson method and CP Method and the nonlinear programming solver. Two tra-
ditional problems, the Brachistochrone Problem and Zermelo?s Problem, are solved
using the DCNLP method to illustrate the accuracy of this method in Chapter 5.
In Chapter 6 a description is given of the 3-D aircraft model. Then the indirect
optimal method is applied to get a MTTC trajectory in constrained airspace. Follow-
ing that, the DCNLP method is applied to this same problem. Results from the two
methods are compared. Solutions to the minimum-fuel-to-climb (MFTC) problem
are also presented.
Chapter 7 starts with a derivation of the traditional PNG law. It is then applied
to 3-D intercept problem. Then, results using DCNLP method to solve this same MTI
problem are compared to the results using PNG law. Finally, field-of-view constraints
are added to the original problem and new trajectories that are consistent with the
additional constraints are discussed.
Conclusions are presented in Chapter 8. Some suggestions for future research
are also provided.
6
Chapter 2
Literature Review
The methods developed in solving trajectory optimization, especially MTTC
problems, almost became the signs of the development in the fields of optimal control.
Betts[18] summarized these methods in his survey and divided them into two major
branches: Indirect Optimal Control (IOC) and Direct Optimal Control (DOC). In
this chapter, methods in both branches are introduced and compared. The IOC
methods as applied to the MTTC and MTI problems are introduced first, followed
by the DOC methods, and then the DCNLP.
2.1 Methods in Solving MTTC Problem
2.1.1 Gradient Method
The objective of the gradient method[1], also called the method of steepest-
ascent/descent, is to find control variable changes that will cause the maximum in-
crease or decrease in the cost function. A steepest descent computation program
will start from an initial guess of the control variables and then integrate the system
differential equations to get a nominal path using this initial guess. An adjustment
7
of the control is then is determined by the numerical integration of the adjoint dif-
ferential equations when certain perturbations are added to the nominal path. This
procedure is repeated until the convergence criteria are satisfied.
The gradient method has been applied to many TPBVPs. This systematic nu-
merical procedure requires suitable initial guess and reasonable selection of the mean-
square perturbation of the control variable program which is named as ?step size?.
Under some cases, a good initial guess is not available, and this may lead to diver-
gent results. Also, when system equations include many state variables, deriving the
adjoint differential equations will be expensive and tedious. These properties limited
the application of this method.
2.1.2 Energy State Method
Normally, the 2-D aircraft point-mass model is represented by its state variables
of speed, flight path angle and altitude. At the same time, its total energy per
unit mass can be treated as the summation of potential energy and kinetic energy
expressed in terms of the variables of speed and altitude. When this energy is used as
a state variable, the minimum-time approximate model is to find the maximum time
derivative of total energy at different given energy levels. Then, the MTTC path is
the connection of the continuous summit point of excess power at constant energy.
8
This approximation method can also be applied to MFTC and minimum range glide
problems[2].
The results obtained from the energy-state approximation model are similar to
those obtained using the gradient method. But, if higher accuracy is required for the
solution, this approximate method is not suitable.
2.1.3 Singular Perturbation Method
In the system equations, some state variables have more effect on the solution
than others. Some states are important, while others can be neglected without any
obvious change to the original system. When these neglected terms are deleted from
the original equations, the solution to the approximate model is called the outer or
slow solution. Generally, the optimal control problems are in the form of TPBVPs.
It is easy to see that the simplified model cannot satisfy all of the final boundary
constraints. So in a fast time region, the state variables will vary rapidly from the
outer stage to the boundary constraints, which is called inner or fast solution. This
phenomena is similar to the solution of Navier-Stokes equations, here the names of
outer and inner solution. In the fluid mechanics problem, fast and slow layers of the
hydrodynamics solution exist. Finally, the theory of singular perturbation method is
to find the combination of simplified model and asymptotic expansion series which
represents the approximate and exact solution separately[3].
9
In the 2-D MTTC problem, the flight path angle is less important than the
altitude and altitude is less important than the energy state. So the zeroth-order
outer solution includes only the energy state variables. If higher accuracy is required,
the first-order or higher order terms will be considered. For the MTTC problem, the
first-order solution is very close to the solution obtained using the gradient method.
Compared to the gradient method, the singular perturbation method has advan-
tage of greatly reduced computation burden and avoidance of initial guess for the
adjoint variables. But its application does has some limitation. First of all, some
of the terms must have smaller effect than others. Secondly, the boundary layer
equations need to reach a stable solution which will makes the ?matching? possible.
2.1.4 Modified Sweep Method
In the application of singular perturbation method in TPBVPs, it?s well known
that the two-time-scale solutions are composed of an outer boundary solution that
slowly approaches an equilibrium point and an inner solution rapidly changes from
the equilibrium point to the final boundary conditions. So, finding such a equilibrium
point became the key point of the problem. Rao and Mease[5] proposed a modified
sweep method to solve especially for the outer segment, also named as the infinite
horizon regulator problem.
10
In the modified sweep method, a basis vector is first chosen to identify the stable
and unstable variables in the system. Normally, the eigenvectors of the Jacobian
matrix of the phase rate vector, which is the time derivative of the state and adjoint
variables, is a good choice for this basis vector. The forward integration uses the
unstable rate coordinate as input and the backward integration uses the stable rate
coordinate as input so that the unstable behavior will be suppressed in each sweep.
The new value for the final stable rate and initial unstable rate obtained from each
loop are saved to be used in the next sweep to improve the convergency. The entire
process is repeated until relative error reaches the required level. This modified sweep
method provides a more accurate way to estimate the initial adjoint variables, but it
is used only for the zeroth-order approximation of the system. A complete solution
to the whole problem requires more work.
2.1.5 Conclusion
The above IOC methods reviewed are all designed for solving TPBVPs without
considering path constraints. For the gradient method, finding a good initial guess for
the adjoint variables is the major difficulty. Generally, the state variables and adjoint
variables are combined together so that poor first time estimation may even cause
thesevaluestoexceedthenumericalrangeofthecomputer. Althoughsometechniques
are proposed to improve the efficiency of initial guess[28], they simultaneously add
11
complexity to the problem while reducing the sensitivity of the solution. The energy
state approximation method is applicable only to a 2-D airplane model. When another
system model is used, another optimization method must be employed. Both singular
perturbation and modified sweep methods are based on the Hamiltonian and the
found necessary conditions. But these procedures do not produce unified solutions
to every optimization problem which means that although the basic theory is same,
the analytic expression must be derived according to different problems. Therefore,
when a system model is complex and includes many state variables, this task is quite
heavy.
When path constraints are considered, the Hamiltonian must be reformulated
with introduction of new Lagrange multipliers. This makes the problem more difficult
to solve, especially when the constraints are inequalities, the trajectory needs to be
divided into piecewise parts of active and inactive arcs to meet the specified path
inequality constraints.
2.2 Methods in Solving MTI Problem
Optimal methods for solving MTI problems have been used for many years.
Practical realization of ?optimality? requires the consideration of uncertainties and
disturbances like wind gusts, temperature changes, weight loss, deviation of starting
point and interception point, command error and other unexpected perturbations.
12
Most optimal methods focus on generation of onboard trajectories varying from short
to long range.
For example, Visser, Kelley and Cliff[11] use a singular perturbation method to
find the approximate optimal trajectory and then apply neighboring optimal guidance
to transfer the aircraft, which has deviated from the reference trajectory, to the
desired trajectory. This procedure is composed of two major parts. The first part
is to get the nominal path that will find the minimum time transfer trajectory from
the initial conditions to the dash point that will possess maximum velocity. The
singular perturbation method applied there has the same theory basis as mentioned
above and the zeroth order approximate model is considered here with application
of state energy model. The second part is to generate the near-optimal guidance
law to eliminate deviation effects and pull the interceptor back to the optimal path.
Then two control functions, load factor and bank angle, are linearized with respect
to the two state variables, altitude and flight path angle, respectively. The feedback
coefficients derived from the linearized model are evaluated at the reference trajectory
and multiplied by the perturbations of the two state variables. Results obtained from
the above calculation are adjusted values about the nominal control commands.
Kumar, Seywald and Cliff[12] proposed finding medium range optimal trajecto-
ries for air-to-air missiles using a three-stage guidance scheme. The first stage is the
boost phase which requires the missile reaches high altitude and desired horizontal
13
distance before the boost motor shut down. Also the nominal time optimal path is
derived with consideration of initial perturbation only. The resultant control load
factor is limited which causes the missile remains at the maximum acceleration limit
line for a finite time period. The second stage is mid-course guidance considering
both state perturbations and a maneuvering target. Transversal comparison and per-
formance augmentation methods has been applied to this longest duration phase. For
the target is capable of maneuvers, the terminal guidance law requires autonomous
tracking of the final object. Pure proportional navigation is applied in the final stage
to intercept the target. These three guidance laws applied to boost, sustain and coast
phases are more accurate than traditional single phase approximate model.
Even the three-phases guidance scheme is based on a simplified approximate
model. But under some circumstances, such an approximate model cannot adequately
represent the system and a more accurate optimization is required. Also at the final
stage when classical Proportional Navigation (PN) guidance law is applied, the time
efficiency is neglected in order to realize real-time control objective. So at least the
final phase is not time optimized which leaves space to make the trajectory further
improved.
Except for the MTI problem, the optimal guidance scheme is also applied to
other areas. For example, Corban, Calise and Flandro[15] used the combination of
singular perturbation and feedback linearization techniques to solve minimum fuel
14
transatmospheric vehicle assent. Bollino, Ross and Doman[16] used pseudospectral
methods to generate the optimal trajectory with minimal miss distance for the reentry
vehicle, then Runge-Kutta integration scheme is applied to integrate forward with
conjunction of disturbances factors to get the current state variables which is feedback
to the on-time optimal trajectory generator to find new control variables at the current
situation. Their simulation results showed that this optimal guidance scheme can
guide the vehicle to an accurate landing even under hurricane wind effect Category
5.
2.3 DCNLP Application Examples
The following are examples of DCNLP applications to aircraft optimal trajectory
design, space mission planning, and unmanned aerial vehicles optimal path planning.
2.3.1 Aircraft Optimal Trajectory Design
1. Hargraves and Paris[19] solved the 2-D MTTC problem using third-order de-
scretization and nonlinear programming. This technology, applied in 1986, is
not as mature as the method used today and some refinement is required to
obtain more accurate results, such as variable scaling, nodes selection, partial
15
derivatives computation and data smoothing. Although not as mature as to-
day?s technology, the method was well developed to successfully solve a wide
variety of problems at that time.
2. Betts and Huffman[20] described a nonlinear programming algorithm in detail in
1993, including a quadratic programming subproblem, merit function, param-
eter definition, algorithm strategy, finding a feasible point and minimization
process. They also compared characteristics of different transcription methods,
trapezoidal, hermite and Runge-Kutta, in terms of computation cost, error and
robust estimation. Maximum crossrange and MTTC problem trajectories are
simulated using the above different transcription methods.
3. Ringertz[21] solved the minimum fuel turn problem using this DCNLP method.
He applied six-degree-of-freedom dynamics model of the aircraft with control of
lift and bank angle. The aircraft was expected to make a heading angle turn
of 180? and keep the same altitude, velocity and flight path angle as initial
conditions while maximizing the final fuel mass.
4. Norsell[22] found a multistage aircraft long distance optimal trajectory consid-
ering radar coverage constraints. In this problem, the 3-D aircraft model is
simplified under different additional conditions, such as constant altitude, con-
stant indicated airspeed or constant march number. The performance to be
16
minimized is the radar cross section beam which will reduce the possibility of
being detected when passing the radar station.
5. Williams, Sgarioto and Trivailo[23] applied a DCNLP method to aerial-towed
cable system path planning. The cable tip was attached to a aircraft with
wings and control surfaces and the interaction between the cable and aircraft
was connected by the tension force of the cable tip. Aerodynamic drag and wind
forces are also considered in the aircraft model. The task was designated to let
the cable tip pass specified points in 3-D airspace while optimizing combined
aircraft performance.
2.3.2 Space Mission Planning
1. Herman and Spencer[24] used higher-order collocation to solve a wide variety
of orbital transfer problems including from low Earth orbit to geosynchronous
Earth orbit, medium Earth orbit, and high Earth orbit while minimizing fuel
consumption. The numerical results of the optimized trajectories were com-
pared to the analytical solution which used burn-coast-burn structured assump-
tion and the comparison showed the results using the two methods are very
close.
17
2. Horie and Conway[25] developed the minimum-time and minimum-fuel optimal
trajectories for an aeroassisted orbital transfer vehicle (AOTV) to intercept a
target located on a lower circular orbit. Because the low density of the atmo-
sphere at high altitude will cause the large reduction of the aerodynamic force
and cause the disappearance of the control variables, the maximum altitude
constraints are considered. Another constraint introduced is one on the aero-
dynamic heating that comes from the atmospheric flight. The results indicate
that the AOTV has more advantages on time consumption than ballistic flight
in interception problem.
3. Betts[26] published his results about space shuttle optimal orbit transfer in the
presence of uncertainty. The transfer task was accomplished by two successive
stages with different impulsive velocity increments and the performance, pay-
load, was maximized. A solid rocket motor was ignited to perform the designed
nominal task and a liquid propellant reaction control system was used to com-
pensate for the off-nominal part. The major uncertainty considered here came
from the specific impulse predicted by weight history.
2.3.3 Unmanned Aerial Vehicles (UAV) Optimal Path Planning
Although DCNLP has wide application in aircraft and space mission planning,
it has limited use in UAVs. So in 2006 Geiger, Horn, DeLullo and Long[27] published
18
their studies on UAVs focusing on maximizing the viewing time for a camera fixed on
the UAV. A single UAV may perform tasks like rendezvousing a slow or fast target or
surveil a stationary target while maximizing the sensor coverage of the object. When
two UAVs are involved, the optimized trajectory may provides full time coverage
of the stationary target. But this method is still not fast enough to be applied in
realtime operation, faster algorithm is required for onboard calculation.
2.3.4 Conclusion
Withtheadvantagesoffastconvergence, avoidanceofadjointvariableestimation,
and the capacity of including complex boundary conditions, the DCNLP has been
successfully applied in wide areas of aircraft and spacecraft research. But this method
has limited application in 3-D trajectory optimization in constrained airspace. In
order to simplify the problem, the previous studies focused on 2-D model considering
constraints about control variables only. The following discussion will make problems
like MTTC and MFTC more complete and accurate with the same properties showed
in 2-D model. The optimality verification will also be conducted to prove the accuracy
of this method.
19
Chapter 3
Indirect Optimal Control Method
3.1 Optimal Control Problem
The objective of an optimal control problem (OCP) is to find the history of
the control variable(s) that will maximize or minimize a given performance index
while satisfying the system constraints. The system constraints considered herein
include first-order, ordinary differential equations subject to initial and final boundary
conditions and some additional constraints on the states and controls. The differential
equations are written here as
?x = f(t,x,u) (3.1)
where x is an n?1 vector of states, f is an n?1 vector of functions, t is the time
and u is an m?1 vector of controls. With prescribed initial conditions
x(0) = x0 (3.2)
and prescribed final boundary conditions
?(tf,xf) = 0 (3.3)
20
where ? is a p?1 vector of functions, t0 is the initial time and the terminal time tf
is free. The scalar performance index is expressed as
J = ?(tf,xf) +
integraldisplay tf
t0
L(t,x,u)dt (3.4)
where ? and L are scalars.
The basic idea of IOC is to find the control u that will make the gradient of
the objective function be zero. So it is necessary to compute the Jacobian gradient
of J to decide which solution will drive it to zero. When the system constraints are
considered, the Hamiltonian equation is introduced as well as the Euler-Lagrange
equations.
3.2 Hamiltonian Equation and Necessary Conditions
The traditional, or indirect, optimal control method is to formulate the Hamil-
tonian function with Lagrange multipliers, also called adjoint variables vector ?j(t)
H = L+?Tf (3.5)
21
where ? = [?1,?2,...,?n]T. Now, the optimal states x(t), controls u(t) and Lagrange
multipliers ?(t) must satisfy the Euler-Lagrange equations
?x = ?H??T = f(t,x,u) (3.6)
??T = ??H
?x = ?
?L
?x ??
T?f
?x (3.7)
The optimized performance is obtained by choosing the control variables, such that
0 = ?H?u = ?L?u +?T?f?u (3.8)
Then, the solution to the indirect OCP is found by solving Eqn.3.6-Eqn.3.8
subject to the constraints Eqn.3.2 and Eqn.3.3. If there are n state variables xj and
m control variables u, from Eqn.3.8, one can express control variables as functions of
the n state variables xj and n adjoint variables ?j. By inserting these new forms of
u into Eqn.3.6 and Eqn.3.7, the problem is solved by integrating the 2n first-order
derivative functions of x and ? while satisfying the specified boundary conditions
on x. If we also know the initial conditions of ?, the problem could be finished by
solving the differential equations. Unfortunately, the initial conditions on the ?j are
not known. Moreover, it is often difficult to find their values.
22
The initial and final state and adjoint variables satisfy boundary constraints that
have the following forms:
Initial Conditions
?xj(t0)?j(t0) = 0 (3.9)
Final Conditions
?(tf) = (???x +?T???x)
T
t=tf
(3.10)
[???t +?T???t + (???x +?T???x)f +L]
t=tf
= 0 (3.11)
?(tf,xf) = 0 (3.12)
where ? is a p?1 new introduced Lagrange multiplier vector.
From the initial conditions, it can be seen that either the initial conditionxj(t0) is
given or?j(t0) = 0. This means that the initial conditions of either the state variables
or the adjoint variables are known. The final boundary conditions must satisfy all
three constraints Eqn.3.10, Eqn.3.11 and Eqn.3.12. Eqn.3.11 is a general case with
the state variables specified at an unspecified terminal time. When the final time is
specified, this equation is not needed. In either case, the information provided at the
final point is not sufficient to solve for the final state variables, Lagrange multipliers
and final time. If all these variables were known at the final time , it would be easy
23
to integrate them backward to get the initial conditions, but they are not. Finally,
the backward integration method still turned out to be unsuccessful.
Considering the importance and difficultly of getting the initial values of the ?j
in solving optimal problems, many researchers have proposed methods to find them.
Since the adjoint variables often have no easily discernable physical meaning, we
cannot determine them from common sense. Furthermore, they sometimes tend to
be very sensitive to changes in the adjoint variables. That is, a small change can cause
great difference in the calculated answer. All this adds up to make obtaining initial
values of adjoint variables very difficult. This difficulty provided the motivation for
finding ways to avoid using the adjoint variables.
3.3 Path Constraints
In the above discussion, we did not consider the extra constraints, which are func-
tions of state and control variables. Some are types of equality constraints including
state and control variables
C(x,u,t) = 0 (3.13)
where C is a q ? 1 vector. When these constraints are considered in the OCP,
the Hamiltonian function may be reformulated by introducing additional Lagrange
24
multipliers ?
H = L+?Tf +?TC (3.14)
as well as new adjoint equations
??T = ??L
?x ??
T?f
?x ??
T?C
?x (3.15)
and optimality conditions
0 = ?L?u +?T?f?u +?T?C?u (3.16)
The path constraints C here reduces the number of independent control variables to
m?q. IfL, f andC are not explicit functions of time, then the Hamiltonian equation
will be a constant value over the whole time interval[29, 30]. The control variables
can be decided from the Pontryagin?s minimum principle which is formulated as
Theorem 3.3.1 [30]The optimal control u?(t) must make the Hamiltonian an abso-
lute minimum with respect to the control at every point of the minimal path x?(t) and
stated as
H(t,x?,u?,??) ?H(t,x?,u,??) t? [t0,tf] (3.17)
25
Other kinds of constraints are inequality constraints
C(x,u,t) ? 0 (3.18)
where
?
??
???
????
= 0, C < 0
> 0, C = 0
(3.19)
In this case, the adjoint equations are
??T = ??H
?x =
?
????
???
?
??L?x ??T ?f?x, C < 0
??L?x ??T ?f?x ??T ?C?x, C = 0
(3.20)
and the optimality conditions are
?H
?u ?
?L
?u +?
T?f
?u +?
T?C
?u = 0 (3.21)
In solving such inequality constraints problems, it is common practice to divide the
solution space into active(C = 0) and inactive(C < 0) arcs. In the inactive arcs,
the solution is assumed to satisfy the additional constraints and necessary condition
Eqn.3.21 is solved as discussed above. In the active pieces, Eqn.3.18 and Eqn.3.20
must be combined together to solve u(t) and ?(t). The classification of active and
inactive arcs may start by deciding the junction point between these two types. If
26
the control u(t) at the junction point is discontinuous, it is called a corner. But if ?,
?H
?u and H are continuous, it follows that u(t) and ?(t) will be continuous across the
junction point. Generally, making a priori estimate of the arc sequences is quite dif-
ficult. Without prior knowledge of the number of the constrained and unconstrained
subarcs, it is very difficult to fix the accurate junction point and each arc boundary
conditions.
Then, these constraints all add to the complexity of the problem and make finding
the initial values of the adjoint variables even more difficult. Also when the system
equations include many state variables, deriving the adjoint equations will seem to be
a nightmare. These are some of the reasons why the DCNLP method introduced in
Chapter 4 is so attractive as an efficient way to solve complex optimization problems.
27
Chapter 4
The Method of Direct Collocation and Nonlinear Programming
In this chapter we introduce the idea of direct collocation and describe procedures
for the trapezoidal, Hermite-Simpson and CP discretization methods. The nonlinear
programming solver is also introduced along with its detailed algorithm. The impor-
tance of the application of direct collocation method in setting up the NLP solver
will also be illustrated.
4.1 Direct Collocation Method
A complete trajectory of an aerospace vehicle is composed of numerous points
representing its continuous coordinates as they change from initial time to final time.
Actually, if we pick up some discrete characteristic points on this trajectory, they
can be used to reproduce an approximate trajectory that is almost as good, for some
purpose, as the continuous trajectory. The basic idea of Direct Collocation (DC) is
to discretize the continuous solution to a problem represented by state and control
variables by using linear interpolation to satisfy the differential equations. In this
way, an OCP is transformed into a nonlinear programming problem (NLPP). Since
the solution to the OCP is in terms of infinitely many values of state and control
28
1
2
3
0
i
1
i
n
1
n
1
2
3
0
i
1
i
n
n
i
i
1
i
Figure 4.1: Trajectory History Discretization
variables, the DC is an approximation. However, this is also the case for any solution
obtained via numerical integration.
To explain the DC method, we present a simple example illustrated in Fig. 4.2.
The continuous trajectory, T, is divided into n segments at n+1 different time points
such that t0 = t1
radicalBig
V2(xf2 +yf2)?(pyf ?qxf)2 (5.17)
54
Otherwise, the positive sign will be used here. To obtain this optimal solution, the
control variable ? is
? = arctan[yf ?qtmx
f ?ptm
] (5.18)
The results from the DCNLP are compared with this analytical results in table.5.3
under same initial conditionx0 = 0,y0 = 0, but different final conditions and different
velocity of the water.
Table 5.3: Analytical and DCNLP Results for Zermelo?s Problem
final position water velocity analytical solution DCNLP solution
(m, m) (m/sec, m/sec) (sec, rad) (sec, rad)
xf = 100, yf = 100 p = 1, q = 3 tm = 11.111, ? = 0.644 tm = 11.111, ? = 0.644
xf = 50, yf = 80 p = 1, q = 3 tm = 7.231, ? = 0.938 tm = 7.231, ? = 0.938
xf = 50, yf = 80 p = 3, q = 1 tm = 7.712, ? = 1.215 tm = 7.712, ? = 1.215
The analytical solution and the DCNLP results were identical to three significant
figures. These results again verify the accuracy of the DCNLP method in solving at
least some optimal control problems.
55
Chapter 6
Constrained Airspace Minimum Time and Fuel Flight Path
This chapter starts with an introduction of a two-dimensional aircraft model and
the solution of MTTC trajectories using energy state method, gradient method and
DCNLP method. The horizontal projection of this 2-D MTTC trajectory is calculated
and the necessities of considering space constraints are illustrated. Following this the
model is expanded to three dimensions and its MTTC and MFTC trajectories under
different constrained conditions are listed.
6.1 2-D MTTC Problem
6.1.1 2-D Mathematical Model
The state equations for a two-dimensional, point-mass aircraft model that is
commonly used to formulate minimum time problems are[3]
?V = [(T(M,h)?D(M,h,L))/W ?sin?]g
?? = (g/V)[L?cos?]
?h = V sin?
?x = V cos?
(6.1)
56
Here, V is the flight speed, ? is the flight path angle, h is the aircraft?s altitude, and
Horizontal Distance
Altitude
W
r
L
r
?
Horizontal
T
r
D
r
Center of
Aircraft mass
Figure 6.1: 2-D Aircraft Model
x is the horizontal coordinate of the aircraft. Also, T is the magnitude of the thrust,
which is assumed to be aligned with the velocity, D is the drag; W is the weight of
the aircraft, L is the lift; and M is the flight Mach number. The load factor n is the
control variable and is defined as
n = L/W (6.2)
57
The lift and drag are modeled using
L = 12?V2SCL? (6.3)
and
D = 12?V2S[CD0 +?CL??2] (6.4)
respectively. By setting L = nW, we may replace Eqn.6.4 by
D = 12?V2SCD0 +? 2n
2W2
CL??V2S (6.5)
The lift curve slope CL? and zero lift drag coefficient CD0, and the drag due to lift
factor ?, together with the aircraft?s weight and wing area S were used to obtain the
numerical results given in this paper. The atmospheric density, ?, is derived from
the 1976 U.S. Standard Atmosphere[43]. The thrust magnitude is given by a table in
Ref.[2] and reproduced in Appendix A.
One of the conditions identified for special attention in Chapter 4.3 is to avoid
using linear interpolation of tabular data, thereby maintaining continuity and differ-
entiability of the system equations and the objective function. Unfortunately, because
of the physical constraints, analytical expressions for thrust, lift and drag coefficients
are generally unavailable, because propulsion and aerodynamic data from experiments
58
are usually stored in tabular forms. No matter what caused the format of these data,
for computational efficiency, it is necessary to devise an alterative way to express these
data. Therefore, an approximate model is developed and is used here to reproduce
the thrust and aerodynamic properties. At a fixed March number, the thrust can be
expressed as a polynomial function T(h) of degree n that fits the data T(hi), hi with
least square error.
T(h) = pnhn +pn?1hn?1 +...+p1h+p0 (6.6)
Here, pi, i = 0,...,n is the polynomial coefficients of the thrust function. If a series
of such polynomial functions is given for different Mach numbers, then these pi can
be written as a function of Mach number pi(M). The polynomial function of degree
m is used again to approximately fit the variation of these coefficients with Mach
number
pi(M) = qmMn +qm?1Mn?1 +...+q1M +q0 (6.7)
In the thrust table.A.1, consider only the portion for which 0 ? h ? 80k ft and
0 ?M ? 2. Using a second-order polynomial to fit the function T(h), the coefficients
at different March number are listed in table.6.1. From the table, the coefficients data
is more irregular than the thrust function of altitude, so a higher degree of polynomial
approximation function, degree four, is used here to approximate the coefficients as
59
Table 6.1: Polynomial Coefficients at Different March Number
M 0 0.4 0.8 1.2 1.6 2.0
p2 0.0045 0.0041 0.0038 0.0040 0.0019 -0.0033
p1 -0.6523 -0.6101 -0.6085 -0.6830 -0.5367 -0.3551
p0 23.5739 22.6811 24.7042 30.1086 32.0151 32.0435
functions of Mach number. They are in the form of as
p2(M) = 0.0014M4 ?0.0070M3 + 0.0091M2 ?0.0039M + 0.0045 (6.8)
p1(M) = ?0.1221M4 + 0.6723M3 ?1.0093M2 + 0.4581M ?0.6556 (6.9)
p0(M) = 0.9393M4 ?10.2110M3 + 24.4391M2 ?11.3770M + 23.6367 (6.10)
Finally, the thrust functions variation with altitude and Mach number is expressed
as
T(h,M) = (0.0014M4 ?0.0070M3 + 0.0091M2 ?0.0039M + 0.0045)h2+
(?0.1221M4 + 0.6723M3 ?1.0093M2 + 0.4581M ?0.6556)h+
(0.9393M4 ?10.2110M3 + 24.4391M2 ?11.3770M + 23.6367)
(6.11)
Using this same method, the lift and drag coefficients as a function of Mach number
can also be expressed as polynomial functions of degree 4.
CD0(M) = 0.0010M4 ?0.0113M3 + 0.0257M2 ?0.0130M + 0.0066 (6.12)
CL?(M) = 0.0122M4 ?0.0568M3 ?0.1500M2 + 0.2847M + 2.2397 (6.13)
60
6.1.2 Energy State Method
In Eqn.6.1, the state variable V can be replaced mathematically by the total
energy per unit mass, which is defined as
E = 12V2 +gh (6.14)
and then
?E = V ?V +g?h (6.15)
Using ?V and ?h expressions from Eqn.6.1 to rewrite ?E, we find
?E = V(T ?D)/m (6.16)
Assume LsimilarequalW, then the time integration can be expressed as
tf ?t0 =
integraldisplay m
V(T ?D)dE (6.17)
In order to minimize time interval tf ?t0, it is obvious to maximize the denominator
of the right part of Eqn.6.17 at a given energy value E. We already know that
thrust T and drag D is a function about V and E, then the objective is to find the
velocity V that will maximize excess power, V(T(V,E) ?D(V,E)), at a fixed E.
61
Figure 6.2: 2-D MTTC Trajectory Using Energy State Model
Assume the aircraft is going to takeoff from sea level to an altitude of 80k ft and the
contours of constant E are plotted on a h?V plane. According to the above theory,
these numerous points of the maximum V(T(V,E)?D(V,E)) at different contours
constitute the approximate MTTC trajectory which is shown in Fig. 6.2.
In Fig. 6.2, the contours of both the constant energy and constant excess power
are illustrated. The dash line is the MTTC energy path which consists of an acceler-
ation of flight at sea level and a climb after the aircraft gains the speed over march
number. Then the gradient method is applied to this problem to see the difference
from the trajectory obtained above.
62
6.1.3 Gradient Method
The gradient method for this 2-D MTTC problem requires to derive the adjoint
equations for the system first. With introduction a 3?1 adjoint vector [?V,??,?h]T,
the Hamiltonian equation is written as
H = 1 +?V [(T ?D)/W ?sin?]g+??[(g/V)(n?cos?)] +?hV sin? (6.18)
and the derivative of the adjoint equations are
??V = ??H
?V = ??V (
?T
?V ?
?D
?V )
g
W +??
g
V2(n?cos?)??h sin? (6.19)
??? = ??H
?? = ?Vgcos????
g
W sin???hV cos? (6.20)
??h = ??H
?h = ??V
g
W
?T
?h (6.21)
with the initial and final boundary conditions
V(0) = V0, h(tf) = hf
?(0) = ?0, ??(tf) = [0 0 0]prime
h(0) = h0, ??(tf) = [0 0 1]prime
(6.22)
Here ?? and ?? are Lagrange multipliers, their definition and the detailed procedure
of the gradient method are given in Appendix B. In using the gradient method, it
63
starts from an initial guess of control variables and final time to get the nominal path
and then gets the corrections of the control variables according to the procedure in
Appendix B. This process is repeated until the convergence condition is reached. As
an example, in the 2-D MTTC problem, assume that the aircraft starts at at sea level
with initial and final conditions,
V(0) = 558.2ft/sec, h(0) = 0
?(0) = 0, h(tf) = 80kft
(6.23)
The vehicle?s MTTC trajectory is shown in Fig. 6.3 and the time histories of the load
factor, velocity, flight path angle and altitude are shown in Fig. 6.4 separately. The
minimum time for the aircraft fly from sea level to an altitude of 80k ft is approx-
imately 180 sec. Compared to the energy state method, the results found from the
gradient method is more accurate and the state and control variable histories reflect
the real values instead the approximate values used in previous method. But, this
method?s successful implement depends on an adequate guess of the initial control
variables. This means that a large deviation of the initial guess will cause diver-
gence. Furthermore, the whole gradient calculation procedure is a burdensome task.
However, these difficulties can be avoided in the DCNLP method.
64
0 0.5 1 1.5 2
0
1
2
3
4
5
6
7
8
x 10
4
March Number
Altitude (ft)
Figure 6.3: 2-D MTTC Trajectory Using Gradient Method
6.1.4 DCNLP Method
The DCNLP method that is applied in this sub-section to solve a 2-D MTTC
problem uses CP discretization method with 18 collocation points. The objective is
to minimize tf subject to the equality constraints
2
tfDV = [(T ?D)/W ?sin?]g
2
tfD? = (g/V)[n?cos?]
2
tfDh = V sin?
(6.24)
65
0 20 40 60 80 100 120 140 160 180
0
2
4
time (sec)
n
0 20 40 60 80 100 120 140 160 180
0
1000
2000
v (ft/sec)
0 20 40 60 80 100 120 140 160 180
-2
0
2
gamma
0 20 40 60 80 100 120 140 160 180
0
5
x 10
4
Altitude (ft)
Figure 6.4: 2-D MTTC State and Control Variables History Using Gradient Method
Here V, ?, h, T, D and n represent the solution vectors to this problem at the
collocation points. The initial and final conditions are the same as in the gradient
method examples. The minimum time obtained here is 178.25 sec with boundary
constraints and system equality constraints well satisfied. The MTTC trajectory
using this method is shown in Fig. 6.5 and the time history of load factor, velocity,
flight path angle and altitude are shown in Fig. 6.6 separately.
The results using DCNLP method are very close to the results obtained using
the gradient method, but the initial guesses for all the collocation points is more
arbitrary. Also, the convergence of the solution is much more rapid.
66
Considering the complexity of 3-D aircraft model which will includes more state
variables and control variables, it will be more difficult to obtain the adjoint derivative
equations and guess adequate nominal control variables. So the DCNLP method
which will avoid deriving adjoint equations and guessing initial values is superior
than other two methods in solving the following 3-D problems.
0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2
0
1
2
3
4
5
6
7
8
x 10
4
March Number
Altitude (ft)
Figure 6.5: 2-D MTTC Trajectory Using DCNLP Method
In Fig. 6.7, the trajectory of an ?Altitude-Down Range? plot is illustrated. From
this figure, we see that the footprint required for an aircraft with modest performance
to climb from sea level to the desired altitude of 80kft, the down range distance will be
more than half of its altitude. This is because most of the time the aircraft?s flight path
67
0 20 40 60 80 100 120 140 160 180
0
5
10
15
time(sec)
n
0 20 40 60 80 100 120 140 160 180
500
1000
1500
2000
v (ft/sec)
0 20 40 60 80 100 120 140 160 180
0
1
2
gamma
0 20 40 60 80 100 120 140 160 180
0
2
4
6
8
x 10
4
Altitude (ft)
Figure 6.6: 2-D MTTC State and Control Variables History Using DCNLP Method
angle is less than pi/3 which makes the horizontal projection of the trajectory longer
than half of its magnitude of the altitude. While for some cases, such long horizontal
distance is not available because of local terrains constraints, so it?s necessary for the
aircraft to make some turns to avoid the collision. Thus, the 3-D aircraft model is
considered and some maneuvers are conducted based on this new model.
68
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5
x 10
4
0
1
2
3
4
5
6
7
8
x 10
4
Down Range (ft)
Altitude (ft)
Figure 6.7: 2-D MTTC Altitude and Down Range Trajectory Using DCNLP Method
6.2 3-D MTTC Problem
6.2.1 3-D Mathematical Model
Unlike the 2-D aircraft model, which focuses on the control in the longitudinal
plane only, the 3-D aircraft model also considers the control in lateral plane. So the
previously used control variable is replaced with two controls, a vertical load factor,
nV , and a horizontal load factor, nh. The heading angle, ?, and cross range, y, as
additional state variables are included in the system functions. Then, the system
69
equations are also expanded to six first-order differential equations,
?V = [(T(M,h)?D(M,h,n))/W ?sin?]g
?? = (g/V)[nV ?cos?]
?? = (g/V)[nh/cos?]
?h = V sin?
?x = V cos?cos?
?y = V cos?sin?
(6.25)
The resultant load factor n is given by
n =
radicalBig
n2V +n2h (6.26)
and the aircraft bank angle, ?, is given by
? = tan?1(nh/nV ) (6.27)
Together with other variables, ? is illustrated in Fig. 6.8. The other properties, such
as thrust, aerodynamic data and air density are treated the same as in the 2-D model.
70
W
r
L
r
?
T
r
D
r
Center of Aircraft
mass
?
?
Figure 6.8: 3-D Aircraft Model
6.2.2 Initial NLP Variable Inputs
Although in most cases the choice of initial input of the NLP variables are arbi-
trary, a good guess generally improves the rate of convergence and avoids singularities
of the Jacobian matrix. Assuming that the aircraft is climbing inside a cylinder with
square projection in x-y plane, in order to allow for a large enough horizontal flight
distance, the final trajectory will make good use of the constrained airspace as much
71
as possible. An initial guess of the path will probably be similar to a helix curve
wrapped on the a right-circular cylinder of radius R inscribed in the square cylinder,
as shown in Fig. 6.9.
?
?
Figure 6.9: Circular Helix Curve Wrapped on A Cylinder
72
If the aircraft flies the helix curve with a constant speed and inclination angle.
Then, the system equations for the initial trajectory can be simplified to
?V = 0 =?V = VI
?? = 0 =?? = ?I
?? = VI cos?/R
?h = V sin? =?h = VItsin?
x = Rsin?
y = Rcos?
(6.28)
where VI and ?I are the pre-assumed constant velocity and constant flight path angle
of the aircraft. Normally, they are assumed to be the take-off velocity and flight path
angle separately. If the time intervals are chosen as Chebyshev points, then, all the
initial NLP variables can be fixed according to Eqn.6.28.
73
6.2.3 Results
The system equation constraints using the CP discretization method in this 3-D
MTTC problem are formulated as
2
tfDV = [(T ?D)/W ?sin?]g
2
tfD? = (g/V)[nV ?cos?]
2
tfD? = (g/V)[nh/cos?]
2
tfDh = V sin?
2
tfDx = V cos?cos?
2
tfDy = V cos?sin?
(6.29)
with the boundary constraints and performance limitations listed in table.6.2
Table 6.2: Boundary Constraints and Performance Limitations for 3-D MTTC Prob-
lem
Constraints Values
Initial Coordinate Constraints x = x0, y = y0, h = 0
Initial V, ? and ? Constraints V = V0, ? = ?0, ? = ?0
Final Altitude h = hf
Airspace Constraints xmin ?x?xmax, ymin ?y ?ymax
Maximum and Minimum Vertical Load Factor nV max = 10, nV min = ?10
Maximum and Minimum Horizontal Load Factor nhmax = 10, nhmin = ?10
The system equation constraints are treated as nonlinear constraints in the NLP
solver and the boundary constraints and performance limitations are set as NLP
74
variable constraints in the solver. The objective function is the time interval between
initial time t0 and final time tf, which is free.
Case 1: Set the airspace constraints on x and y as ?10,000 ft. The airplane is
required to fly from sea level to an altitude of hf = 30,000ft with an initial flight
speed of Mach = 0.5, an initial flight path angle of ?0 = 12.6?, and an initial heading
angle of ?0 = 0. The minimum time obtained for this case is 89.76 sec. The three-
dimensional trajectory is shown in Fig. 6.10 and the time histories of vertical and
horizontal load factors nV and nh, velocity V, flight path angle ? and heading angle
? are shown in Fig. 6.11.
Case 2: Using the same initial conditions as in Case 1, but with the constraints on
both x and y of ?7,500 ft, the aircraft cannot reach 30,000 ft within a time similar
to that in case 1. However, when the final altitude was reduced to 25,000 ft, the
three-dimensional MTTC trajectory shown in Fig. 6.12 with a similar climb time of
91.49 sec was obtained. The corresponding state and control variable time histories
are shown in Fig. 6.13.
Case 3: When the final 30,000 ft altitude was reached under the same initial
condition of Case 1 and with the constraints on both x and y of ?7,500 ft, the
minimum time calculated here is 125.24 sec. The trajectory, corresponding state and
control variable time histories are shown in Fig. 6.14 and Fig. 6.15 respectively.
75
Case 4: When the final flight speed constraint ofM = 0.8 is added to Case 1, the
climb time is increased to 156.01 sec with other initial and final conditions unchanged.
The trajectory, corresponding state and control variable time histories are shown in
Fig. 6.16 and Fig. 6.17, respectively.
Case 5: When the same final flight speed constraint M = 0.8 is added to Case
3 which has the smaller airspace horizontal projection, the climb time under this
condition is 175.28 sec. The trajectory, corresponding state and control variable time
histories are shown in Fig. 6.18 and Fig. 6.19 respectively.
6.2.4 Conclusions
Using the CP discretization method and the NLP solver, the 3-D MTTC problem
has been solved under different assumption conditions. The results show that the
performance index, the time to climb, has been determined accurately and system
equality constraints, boundary constraints and control constraints were satisfied.
In all the five cases, the final trajectory is similar to a helical curve wrapped on
a cylinder, which indicates that the initial guess of a helical trajectory is reasonable.
Cases 1 and 2 have similar climb times, but in Case 1, a higher altitude was reached.
The only difference between the cases is that Case 2 had a smaller constrained space
in terms of the x?y base. As expected, smaller airspace volumes result in more time
for the aircraft to climb to a desired altitude.
76
Case 3 had the same airspace constraints as Case 2, but 20% higher final altitude.
The extra altitude in Case 2 cost 36.9% more in terms of climb time. These data
indicate that the climb time is not proportional to the altitude. It can also be said
that the time distribution is different at different altitude.
The optimal trajectory in each case can be treated as a process of ?climb-dive-
climb?. The aircraft stays at sea level initially to gain speed and then makes a fast
climb. In some cases, the aircraft?s speed decreases to close to zero before the aircraft
reaches its final altitude. To complete the climb task, it?s necessary for the aircraft
to make a dive to gain speed and then climb again. Cases 3 to 5 illustrate this
characteristic. When higher altitudes are required, more dives and climbs may be
performed. That?s why the trajectory is a repeat process of ?climb-dive-climb?. This
property can also be shown in the 2-D MTTC problem, when higher final altitudes
are required.
Another characteristic property of the trajectory is the final speed. If there is no
constraint on the final speed, then the aircraft speed will normally decrease close to
stall speed. This is not realistic. When a constraint on the final flight speed is added,
the stall point can be avoided subject to the penalty of more climb time. This can
be seen by comparing climb times for Cases 1 and 4 or Cases 3 and 5.
77
6.3 3-D Minimum-Fuel-To-Climb Problem
In the 3-D MTTC problem, the aircraft weight is treated as a constant. Actually,
the consumption of fuel during this time interval makes the weight of the aircraft a
function of time. The change of the weight is omitted in above discussion because
the climb task is completed in a few minutes, which makes this change of weight
negligible. If this weight change is modeled, then the additional equation,
?m = ?f = T/cg (6.30)
where c = 2800 sec and f is the thrust specific fuel consumption, is included and the
mass is an additional state variable to the pervious 6?1 state variable vector. And
the previous system equations keep the same except the weight is substituted as a
function of time
W(t) = m(t)g (6.31)
The objective function now is to maximize the final weight
J = ?Wf (6.32)
Some cases are tested with same boundary constraints and performance limitations
set as in previous MTTC problem.
78
Case 6: Use the same initial and final conditions and airspace constraints in Case
1 with unspecified climb time. The MFTC trajectory spent similar time, 93.13 sec,
comparing to Case 1. The trajectory, corresponding state and control variable time
histories are shown in Fig. 6.20 and Fig. 6.21 respectively. The trajectory in Case
6 is similar to the trajectory in Case 1 and the mass of the aircraft changes in a
small scale which can be negligible. In order to obtain the objective of minimum fuel
consumption in the climb procedure, it is expected that the aircraft can reach the
final altitude in the least time to consume as little as possible fuel. From this point
of view, it is reasonable that the MTTC trajectory and MFTC trajectory is close to
each other.
Case 7: Use the same initial and final conditions, airspace constraints as in Case
3 with unspecified climb time. The MFTC trajectory also spent similar time 140.54
sec comparing to Case 3. The trajectory, corresponding state and control variable
time histories are shown in Fig. 6.22 and Fig. 6.23 respectively. When the airspace
constraints became smaller, it will cost more fuel to climb to the same altitude, the
same conclusion as the MTTC trajectories in different airspace constraints.
Case 8: When the same final flight speed constraint M = 0.8 is added to case
6 with final time still unspecified, the climb time under this condition is 175.28 sec.
The trajectory, corresponding state and control variable time histories are shown in
Fig. 6.24 and Fig. 6.25 respectively. Also as in corresponding MTTC conclusion,
79
final speed constraints cost more time and fuel consumption comparing to none final
speed constraints.
Case 9: All the above MFTC cases consider only unspecified final time, the NLP
solver will choose the optimized time to get the minimum fuel results. When the final
time is specified as 100 sec in Case 6, the consumption of fuel is more than previous
result. The MFTC trajectory, corresponding state and control variable time histories
are shown in Fig. 6.26 and Fig. 6.27 respectively. Any specified time larger than the
time got from the unspecified final time result will cost more fuel consumption.
80
-1
-0.5
0
0.5
1
-1
-0.5
0
0.5
1
0
0.5
1
1.5
2
2.5
3
X(10000ft)
Y(10000ft)
Altitude(10000ft)
Figure 6.10: 3-D MTTC Trajectory for Case 1
81
0 10 20 30 40 50 60 70 80 90
-1
0
1
2
3
4
nv an
d
n
h
nv
nh
0 10 20 30 40 50 60 70 80 90
0
0.05
0.1
v (
10000f
t
/
sec)
0 10 20 30 40 50 60 70 80 90
0
0.5
1
1.5
2
f
l
i
ght
pat
h angl
e (
r
a
d)
0 10 20 30 40 50 60 70 80 90
0
2
4
6
8
time(sec)
head
i
ng a
ngl
e
(
r
ad)
Figure 6.11: Control and State Variables History for Case 1
82
-0.5
0
0.5
-0.5
0
0.5
0
0.5
1
1.5
2
2.5
X(10000ft)
Y(10000ft)
Altitude(10000ft)
Figure 6.12: 3-D MTTC Trajectory for Case 2
83
0 20 40 60 80
-2
0
2
4
nv
and nh
nv
nh
0 20 40 60 80
0
0.05
0.1
v (
1
0000f
t
/
s
ec)
0 20 40 60 80
-0.5
0
0.5
1
1.5
2
f
l
i
g
ht
pat
h
angl
e (
r
a
d)
0 20 40 60 80
0
2
4
6
8
time(sec)
headi
ng angl
e
(
r
a
d)
Figure 6.13: Control and State Variables History for Case 2
84
-0.5
0
0.5
-0.5
0
0.5
0
0.5
1
1.5
2
2.5
3
X(10000ft)
Y(10000ft)
Altitude(10000ft)
Figure 6.14: 3-D MTTC Trajectory for Case 3
85
0 20 40 60 80 100 120
-2
0
2
4
6
nv
and nh
nv
nh
0 20 40 60 80 100 120
0.04
0.06
0.08
0.1
0.12
v (
1
0
0
0
0
f
t
/
se
c)
0 20 40 60 80 100 120
-1
0
1
2
f
light
pat
h
angle (
r
ad)
0 20 40 60 80 100 120
0
5
10
15
time(sec)
headi
n
g angl
e (
r
ad)
Figure 6.15: Control and State Variables History for Case 3
86
-1
-0.5
0
0.5
1
-1
-0.5
0
0.5
1
0
0.5
1
1.5
2
2.5
3
X(10000ft)
Y(10000ft)
Altitude(10000ft)
Figure 6.16: 3-D MTTC Trajectory for Case 4
87
0 50 100 150
-2
0
2
4
nv
and nh
nv
nh
0 50 100 150
0
0.05
0.1
v (
1
0
0
0
0
f
t
/
se
c)
0 50 100 150
-2
-1
0
1
2
f
l
i
g
ht
pat
h
angl
e (
r
a
d)
0 50 100 150
0
2
4
6
8
time(sec)
headi
ng angl
e (
r
ad)
Figure 6.17: Control and State Variables History for Case 4
88
-0.5
0
0.5
-0.5
0
0.5
0
0.5
1
1.5
2
2.5
3
X(10000ft)
Y(10000ft)
Altitude(10000ft)
Figure 6.18: 3-D MTTC Trajectory for Case 5
89
0 20 40 60 80 100 120 140 160 180
-5
0
5
10
nv
and nh
nv
nh
0 20 40 60 80 100 120 140 160 180
0
0.05
0.1
v
(1
0
0
0
0
ft/s
e
c
)
0 20 40 60 80 100 120 140 160 180
-2
-1
0
1
2
f
l
i
g
ht
pat
h
angl
e (
r
ad)
0 20 40 60 80 100 120 140 160 180
0
5
10
time(sec)
headi
ng angl
e (
r
a
d)
Figure 6.19: Control and State Variables History for Case 5
90
-1
-0.5
0
0.5
1
-1
-0.5
0
0.5
1
0
0.5
1
1.5
2
2.5
3
X(1000ft)
Y(1000ft)
Altitude(1000ft)
Figure 6.20: 3-D MFTC Trajectory for Case 6
91
0 20 40 60 80 100
-2
0
2
4
nv
and nh
nv
nh
0 20 40 60 80 100
0
0.05
0.1
v (
10000f
t
/
s
ec
)
0 20 40 60 80 100
-0.5
0
0.5
1
1.5
fl
i
ght path angl
e (
r
ad
)
0 20 40 60 80 100
0
2
4
6
headi
ng angl
e (
r
ad)
0 20 40 60 80 100
3.35
3.4
3.45
time(sec)
wei
ght
(
1
0000l
b)
Figure 6.21: Control and State Variables History for Case 6
92
-0.5
0
0.5
-0.5
0
0.5
0
0.5
1
1.5
2
2.5
3
X(10000ft)
Y(10000ft)
Altitude(10000ft)
Figure 6.22: 3-D MFTC Trajectory for Case 7
93
0 50 100 150
-2
0
2
4
6
nv and nh
nv
nh
0 50 100 150
0
0.05
0.1
v
(
1
00
00
f
t
/
s
ec)
0 50 100 150
-2
-1
0
1
2
f
l
i
ght
pat
h angl
e (
r
ad)
0 50 100 150
0
2
4
6
8
headi
ng
a
ngl
e
(
r
ad)
0 50 100 150
3.3
3.35
3.4
3.45
time(sec)
w
e
i
g
ht
(
1000
0l
b)
Figure 6.23: Control and State Variables History for Case 7
94
-1
-0.5
0
0.5
1
-1
-0.5
0
0.5
1
0
1
2
3
4
X(10000ft)
Y(10000ft)
Altitude(10000ft)
Figure 6.24: 3-D MFTC Trajectory for Case 8
95
0 20 40 60 80 100 120 140 160
-2
0
2
4
6
nv and nh
nv
nh
0 20 40 60 80 100 120 140 160
0
0.05
0.1
v (
1
000
0f
t
/
sec)
0 20 40 60 80 100 120 140 160
-2
-1
0
1
2
f
l
i
g
h
t
pa
t
h
an
g
l
e (
r
a
d)
0 20 40 60 80 100 120 140 160
0
2
4
6
8
headi
ng an
gl
e
(
r
ad)
0 20 40 60 80 100 120 140 160
3.3
3.35
3.4
3.45
time(sec)
w
e
i
ght
(
100
00l
b)
Figure 6.25: Control and State Variables History for Case 8
96
-1
-0.5
0
0.5
1
-1
-0.5
0
0.5
1
0
0.5
1
1.5
2
2.5
3
X(10000ft)
Y(10000ft)
Altitude(10000ft)
Figure 6.26: 3-D MFTC Trajectory for Case 9
97
0 20 40 60 80 100
-2
0
2
4
nv an
d nh
nv
nh
0 20 40 60 80 100
0
0.05
0.1
v (
1
0
0
0
0
f
t
/
se
c)
0 20 40 60 80 100
-1
0
1
2
f
l
i
g
ht
p
a
t
h
an
gl
e (
r
a
d
)
0 20 40 60 80 100
0
5
10
headi
ng angl
e (
r
ad)
0 20 40 60 80 100
3.35
3.4
3.45
3.5
time(sec)
w
e
i
g
ht
(
1
000
0l
b
)
Figure 6.27: Control and State Variables History for Case 9
98
Chapter 7
Other Constrained Optimal Trajectories
This chapter will focus on solving problems about optimal trajectories for an
aircraft intercepting or rendezvousing a constant velocity target. Results for several
different initial interceptor and target velocities are presented. A proportional navi-
gation type guidance law has been applied to the same interception problem to obtain
results that are compared to the time optimal trajectories. Then Field-of-view[35, 36]
limitations are introduced and added to the DCNLP problem and results for this more
constrained case are compared with original problem results.
7.1 Three-Dimensional Proportional Navigation Guidance Law
The most commonly used guidance laws are forms of Proportional Navigation
Guidance (PNG). It is therefore of interest to compare the trajectories generated
using DCNLP with corresponding PNG trajectories. Here, we use the idea of a three-
dimensional PNG law formulated in terms of the relative geometry. The commanded
acceleration from this law is perpendicular to the instantaneous interceptor to target
line-of-sight (LOS) and proportional to the LOS angular velocity magnitude (?LOS
rate?) times the magnitude of the closing velocity[6]. Mathematically, this can be
99
expressed as[9, 10]:
??A
I = ?K ?RTI
??R
TI
|RTI| ?
??? (7.1)
In Eqn.7.7, K is the constant effective navigation ratio (to be chosen greater than
2), ??? is the LOS angular velocity which is provided by the interceptor radar, and
??R
TI =
??R
T ?
??R
I is the relative position vector of the target with respect to the
interceptor and can be easily measured by the Doppler radar. The desired acceleration
command ??AI is perpendicular to both the LOS and the LOS angular velocity. This
relationship is represented geometrically in Fig. 7.1.
In Fig. 7.1, the line connecting the interceptor and target is known as LOS, the
LOS composed an angle velocity ? with respect to the fixed reference frame. For
simulation purposes, the LOS angular velocity may be computed using:
??? = ??RTI ???V TI
|RTI|2 (7.2)
where ??V TI = ??V T ? ??V I is the relative velocity of the target with respect to the
interceptor. The other component of acceleration ?RTI is found from
?RTI = RTIxVTIxx +RTIyVTIy +RTIzVTIz
|RTI| (7.3)
100
T
R
r
I
R
r
TI
R
r
T
V
r
I
V
r
Figure 7.1: Engagement Geometry
where RTIx, RTIy and RTIz and the X, Y and Z components respectively of RTI and
VTIx, VTIy and VTIz are X, Y and Z component of VTI.
For the purpose of interception, is was expected that the distance between the
interceptor and target RTI is as small as possible and the identical value is zero.
The final closest distance between them is known as miss distance. From the theory
of maximum and minimum, when a maximum or minimum point of a function is
101
obtained, the derivative of this function at this point will be zero and sign of this
derivative function will change before and after this summit point. So that at the
closest point, the derivative of the distance between interceptor and target, VTI, is
zero and the sign of this value will change after this point is reached. So the ending
point of this maneuver is when the VTI changes its sign. This is also called necessary
conditions for capture. If this condition can not be satisfied in expected intercept time,
it indicates that the interceptor can not perform this task under the current initial
condition and desired K value. Then it is necessary to adjust the initial condition of
the interceptor or change the K value to meet the capture condition.
This 3-D PNG law is applied in the problem of intercepting a constant-velocity,
constant-altitude target under different flight conditions. So the target velocity is
treated as a pre-defined value during the whole process as well as its initial coordi-
nates. The target?s corresponding position change can also be easily derived accord-
ing to its initial condition. The interceptor?s initial velocity and coordinates are also
known. Thereafter, the components of its acceleration ?, ??RTI and ?RTI are all known
at this launch point. So that we can integrate the interceptor velocity and 3-D coor-
dinates values forward using trapezoidal integration method to get its corresponding
velocity and coordinates in the next step time point. And this procedure is repeated
until the closing velocity VTI changes its sign.
102
Some Cases are simulated using this PNG law in the following section with
different launch velocity and target velocity. Trajectory results, state variables and
control variables history are compared with the corresponding simulation results using
DCNLP method with the objective of minimizing interception time. Some cases also
include the situation when the capture condition can not be satisfied in the desired
interception time to show the advantage of DCNLP method in solving such kind of
problems. The PNG algorithm code is included in Appendix C.
7.2 Three-Dimensional Minimum-Time Interception Trajectory Planning
The ideal interception condition is that the interceptor and target has the same
coordinate, so we have the following relationship
hI = hT, xI = xT, yI = yT (7.4)
here, the subscripts T and I are used to denote the interceptor and the target, re-
spectively. Eqn.7.4 can be treated as final boundary conditions of the MTI problem.
Because the target is flying under constant velocity. Assuming the y axis is parallel
to the velocity direction of the target, then the final position of the target can be
103
expressed as
hTf = hT
xTf = xT
yTf = yT0 +VT(tf ?t0)
(7.5)
where hT, xT and VT are the constant altitude, x-position and speed of the target, re-
spectively. Also,yT0 andyTf are the initial and final y-position of target, respectively.
tf is the final time and t0 is the initial time. Its performance limitations for Aircraft
2 as well as the initial conditions and final interception conditions are included in
table.7.1.
Table 7.1: Boundary Constraints and Performance Limitations for 3-D MTI Problem
Constraints Values
Initial Coordinate Constraints xI0 = x0, yI0 = y0, hI0 = 0
Initial V, ? and ? Constraints V = V0, ? = ?0, ? = ?0
Final Altitude and x-position of interception point hIf = hT, xIf = xT
Final y position of interception point yIf = yT0 +VT(tf ?t0)
Maximum and Minimum Vertical Load Factor nV max = 2, nV min = ?2
Maximum and Minimum Horizontal Load Factor nhmax = 2, nhmin = ?2
In table.7.1, hIf, xIf and yIf are the final altitude, x-position and y-position of
the interceptor, respectively. The state and control variables constraints in table.7.1
together with the nonlinear constraints listed in Eqn.6.29 and objective function
J = tf ?t0 formulated the MTI problem. The DCNLP solution will optimize the
interception time and at the same time satisfy both the maneuvered system equations
and the interception conditions.
104
Case 1: set the initial conditions on the interceptor state as: xI0 = 0, yI0 =
?10000 ft, hI0 = 0, ?0 = 12.6?, MI0 = 0.4; set the initial and constant conditions
on the target as: xT = 10000 ft, yT0 = ?10000 ft, hT = 10000 ft and MT = 0.5. In
Case 1, the minimum time for the interceptor to reach the target is 47.82 sec and
the miss distance at the terminal time is very close to zero. The 3-D trajectory and
relative properties for this case are shown in Fig. 7.2 and Fig. 7.4. The trajectory
and velocity history of Case 1 using PNG laws with K = 5 is shown in Fig. 7.3 and
Fig. 7.4. Interception time using this method is 418.473 sec and the miss distance is
0.0257 ft. It?s obvious to see that the DCNLP planning trajectory will intercept the
target with much less time and high accuracy.
Case 2: For the interceptor, the same initial conditions as Case 1 are used, except
we set MI0 = 0.5 and MT = 0.4. The minimum time for the interceptor to reach
the target is 36.52 sec and the miss distance at the terminal time is also close to
zero. The 3-D trajectory and relative properties of this case is shown in Fig. 7.5 and
Fig. 7.7. The trajectory and flight speed history for Case 2 using the PNG laws with
K = 5 are shown in Fig. 7.6 and Fig. 7.7. The interception time using this method
is 272.025 sec and the miss distance is 0.0026 ft. The time is so large because the
interception speed does not increase as much as in the MTI solution.
Case 3: For the interceptor, the same initial conditions as Case 1 are used, except
we set MT = 0.6. The minimum time for the interceptor to reach the target is 51.45
105
sec and the miss distance at the terminal time is also close to zero. The 3-D trajectory
and relative properties of this case is shown in Fig. 7.8 and Fig. 7.10. While the
PNG law does not work on this case. The closing velocity will change its sign in the
distance of 12492.1 ft which is far away from the desired miss distance. The trajectory
and flight speed history for Case 3 using the PNG laws with K = 5 are shown in Fig.
7.9 and Fig. 7.10.
Case 4: For the interceptor, the same initial conditions as Case 2 are used, except
we set MI = 0.6. The minimum time for the interceptor to reach the target is 30.43
sec and the miss distance at the terminal time is also close to zero. The 3-D trajectory
and relative properties of this case is shown in Fig. 7.11 and Fig. 7.13. While the
PNG law does not work on this case. The closing velocity will change its sign in the
distance of 13368.7 ft which is far away from the desired miss distance. The trajectory
and flight speed history for Case 4 using the PNG laws with K = 5 are shown in Fig.
7.12 and Fig. 7.13.
7.3 Three-Dimensional Minimum-Time Rendezvous Trajectory Planning
The 3-D minimum-time rendezvous (MTR) problem is to catch up the target, also
in constant-altitude and constant-velocity, with same velocity and desired distance.
So this kind of problem is similar to the 3-D MTI problem except that MTR problem
requires the interceptor has same final velocity as the target?s velocity in addition to
106
the final coordinates constraints. They are formulated as
hI = hT xI ?xT = d
yI = yT vIf = vT
?If = 0 ?If = ?pi/2
(7.6)
where d is the final x-position distance between the interceptor and the target, vIf,
?If and ?If are the final velocity, flight path angle and heading angle of the in-
terceptor, respectively. Two cases are simulated using different launch velocity and
target velocity for the MTR problem with same performance limitations illustrated
in table.7.1.
Case 5: Use the same initial conditions of interceptor and target as Case 1.
The final distance in x-position is d = 50 ft. The minimum time for the interceptor
to catch up the target is 38.84 sec using DCNLP method. The 3-D trajectory and
relative properties of this case is shown in Fig. 7.14 and Fig. 7.16.
Case 6: Use the same initial conditions of interceptor and target as Case 2.
The final distance in x-position is also 50 ft. The minimum time for the interceptor
to catch up the target is 57.90 sec using DCNLP method. The 3-D trajectory and
relative properties of this case is shown in Fig. 7.15 and Fig. 7.17.
107
0
0.2
0.4
0.6
0.8
1
-1
0
1
2
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.2: 3-D MTI Trajectory using DCNLP for Case 1
0
0.2
0.4
0.6
0.8
1
0
5
10
15
20
25
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.3: 3-D MTI Trajectory using PNG for Case 1
108
0 5 10 15 20 25 30 35 40 45 50
0.5
1
1.5
2
ve
r
t
i
c
a
l
l
o
ad
f
a
ct
o
r
0 5 10 15 20 25 30 35 40 45 50
-1.5
-1
-0.5
0
h
o
r
i
zo
nt
a
l
l
oad
f
a
c
t
or
0 5 10 15 20 25 30 35 40 45 50
400
500
600
700
800
900
f
l
i
ght
spe
ed (
f
t
/
s
e
c
)
DCNLP
0 50 100 150 200 250 300 350 400 450
440
460
480
500
520
540
560
f
l
i
ght
s
peed
(
f
t
/
s
e
c
)
time (sec)
PNG
Figure 7.4: Control Factors and Velocity History for Case 1
109
0
0.2
0.4
0.6
0.8
1
-1
-0.5
0
0.5
1
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.5: 3-D MTI Trajectory using DCNLP for Case 2
0
0.2
0.4
0.6
0.8
1
0
2
4
6
8
10
12
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.6: 3-D MTI Trajectory using PNG for Case 2
110
0 5 10 15 20 25 30 35 40
0.5
1
1.5
2
ver
t
i
c
al
l
oad
f
a
ct
or
0 5 10 15 20 25 30 35 40
-2
-1.5
-1
-0.5
0
ho
r
i
zo
n
t
a
l
l
o
ad
f
a
ct
or
0 5 10 15 20 25 30 35 40
550
600
650
700
750
f
l
i
g
h
t
sp
eed (
f
t
/
sec)
DCNLP
0 50 100 150 200 250 300
400
420
440
460
480
500
520
540
560
f
l
i
ght
s
peed
(
f
t
/
s
e
c
)
time (sec)
PNG
Figure 7.7: Control Factors and Velocity History for Case 2
111
0
0.2
0.4
0.6
0.8
1
-1
0
1
2
3
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.8: 3-D MTI Trajectory using DCNLP for Case 3
0
0.2
0.4
0.6
0.8
1
-1
-0.5
0
0.5
1
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.9: 3-D MTI Trajectory using PNG for Case 3
112
0 10 20 30 40 50 60
0
0.5
1
1.5
2
v
e
rti
c
a
l
lo
a
d
fa
c
t
o
r
0 10 20 30 40 50 60
-2
-1
0
1
2
hor
i
z
ont
al
l
oad f
a
ct
or
0 10 20 30 40 50 60
400
600
800
1000
f
l
i
ght
speed (
ft/s
e
c
) DCNPL
0 5 10 15 20 25 30
600
620
640
660
680
time (sec)
flig
h
t
s
p
e
e
d
(ft/s
e
c
)
PNG
Figure 7.10: Control Factors and Velocity History for Case 3
113
0
0.2
0.4
0.6
0.8
1
-1
-0.5
0
0.5
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.11: 3-D MTI Trajectory using DCNLP for Case 4
0
0.2
0.4
0.6
0.8
1
-1
-0.5
0
0.5
1
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.12: 3-D MTI Trajectory using PNG for Case 4
114
0 5 10 15 20 25 30 35
0.5
1
1.5
2
ver
t
i
c
al
l
oad
f
a
ct
or
0 5 10 15 20 25 30 35
-2
-1.5
-1
-0.5
0
hor
i
z
o
n
t
a
l
l
o
ad
f
a
ct
or
0 5 10 15 20 25 30 35
600
650
700
750
f
l
i
ght
spe
ed (
f
t
/
sec) DCNLP
0 5 10 15 20 25
440
450
460
470
480
490
time (sec)
fli
g
h
t
s
p
e
e
d
(ft/s
e
c
)
PNG
Figure 7.13: Control Factors and Velocity History for Case 4
115
0
0.2
0.4
0.6
0.8
1
-1
-0.5
0
0.5
1
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.14: 3-D MTR Trajectory using DCNLP for Case 5
0
0.2
0.4
0.6
0.8
1
-1
0
1
2
3
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.15: 3-D MTR Trajectory using PNG for Case 6
116
0 5 10 15 20 25 30 35 40
-2
-1
0
1
2
ve
r
t
i
c
al
l
oad f
a
ct
or
0 5 10 15 20 25 30 35 40
-2
-1
0
1
2
hor
i
z
ont
al
l
oad
f
a
ct
or
0 5 10 15 20 25 30 35 40
0
0.5
1
f
l
i
ght
pat
h
angl
e(
r
ad)
0 5 10 15 20 25 30 35 40
-2
-1.5
-1
-0.5
0
head
i
ng ang
l
e
(
r
ad)
0 5 10 15 20 25 30 35 40
400
500
600
700
800
time (sec)
f
l
i
ght
speed (
f
t
/
sec)
Figure 7.16: Control Factors and State Variables History for Case 5
117
0 10 20 30 40 50 60
-2
-1
0
1
2
ve
r
t
i
c
al
l
oad f
a
ct
o
r
0 10 20 30 40 50 60
-2
-1
0
1
2
hor
i
z
ont
al
l
oad f
a
ct
o
r
0 10 20 30 40 50 60
-0.5
0
0.5
1
f
l
i
ght
pat
h angl
e(
r
ad)
0 10 20 30 40 50 60
-2
-1.5
-1
-0.5
0
headi
ng ang
l
e
(
r
ad
)
0 10 20 30 40 50 60
-2
-1.5
-1
-0.5
0
headi
ng ang
l
e
(
r
ad
)
Figure 7.17: Control Factors and State Variables History for Case 6
118
7.4 View Constraints
In tactical flight, the interceptor seeks the target using the homing eye to provide
the necessary measurement required for the implemented guidance law. So the success
of the interception task depends largely on accuracy of the measurement. Some typical
missile uses gyroscopes and a antenna mounted on the gimbals of the interceptor and
the rotation of these gimbals is in a limited angle. So that the interceptor?s look angle
is also limited. These missiles? velocity direction is aligned with the imaginary LOS
and their looking angle is measured from this imaginary LOS.
When a field of view constraint (FVC) is considered for the interceptor the tar-
get is expected to remain within the volume specified during the engagement. In this
paper, we address only the interceptor velocity view limitation. In Fig. 7.18, the
dotted line is the interceptors trajectory. At any point the angle between the inter-
ceptors velocity vector and LOS, ?LOS, is expected to be smaller than the constraint
cone angle, ?c. In other words, the target is required to stay between the two points
intercepted with the target trajectory and cone bottom. This additional condition
can be treated as a nonlinear constraint and stated as:
?LOS = arctan(
??V ???R
TI
|??V |?|??RTI| ??c) (7.7)
119
T
LOS
c
?
LOS
?
V
r
I
Figure 7.18: Control Factors and State Variables History for Case 7
One case is simulated considering this FVC. Case 7: Set the initial flight path
angle of the interceptor as ?I0 = 0.5 with other conditions the same as Case 1. When
we added the FVC condition to the same problem and let ?c = 30?, the minimum
time increased to 42.43 sec. The new trajectory is shown in Fig. 7.19 with relative
properties shown in Fig. 7.20.
7.5 Conclusion
Three-dimensional minimum time intercept trajectory planning has been consid-
ered using direct collocation and nonlinear programming based on CP discretization.
The results show that the performance indices, the interception times, are much less
120
0
0.2
0.4
0.6
0.8
1
-1
-0.5
0
0.5
1
0
0.2
0.4
0.6
0.8
1
X (10000ft)
Y (10000ft)
Altitude (10000ft)
Figure 7.19: 3-D MTI Trajectory using DCNLP for Case 7
than those for trajectories generated by an idealized Proportional Navigation Guid-
ance law, and the miss distances are smaller.
The interception time will of course depend on the engagement conditions, such
as the interceptor?s starting velocity and the target?s speed, since to reach the desired
target altitude, the interceptor needs to make a climb after it gets enough speed.
Hence, the larger the magnitude of the starting velocity, the shorter the interception
time. This property is similar to the 3-D MTTC problem. On the other side, a larger
magnitude constant velocity of target directed away from the interceptor will require a
larger interception time. The DCNLP method showed advantage in short interception
time, low miss distance as well as wide range of initial conditions constraints for
121
0 5 10 15 20 25 30 35 40 45
0
0.5
1
1.5
ve
r
t
i
c
al
l
oad f
a
ct
or
0 5 10 15 20 25 30 35 40 45
0
0.5
1
1.5
hor
i
z
ont
al
l
oad
f
a
ct
or
0 5 10 15 20 25 30 35 40 45
0.35
0.4
0.45
0.5
f
l
i
g
h
t
pa
t
h
an
gl
e
(
r
a
d)
0 5 10 15 20 25 30 35 40 45
0
0.5
1
1.5
2
he
adi
ng angl
e(
r
a
d
)
0 5 10 15 20 25 30 35 40 45
550
600
650
700
time (sec)
f
l
i
ght
spe
ed (
f
t
/
sec)
Figure 7.20: Control Factors and State Variables History for Case 7
122
interception. While PNG law will not work out for any condition of interception
which is illustrated in Case 3 and 4.
The rendezvous results cost more time than the corresponding MTI trajectory
to compensate the time used to satisfy the additional final constraints, those are the
speed, flight path angle and heading angle constraints.
When field-of-view constraints are added to the original problem, more time is
required to intercept because the interceptor cannot turn as rapidly. Although the
direct collocation and nonlinear programming method has many advantages in solving
this kind of optimal control problem, the computation time is large enough that even
with fast computers it is not suitable for real-time applications. When the target?s
velocity is variable and trajectories need to be updated frequently according to radar
tracking data, even more computation time will be needed. Then PNG law is superior
considering on-board application.
123
Chapter 8
Summary and Recommendations
Inthisdissertation, theproblemsoffindingsolutionstothree-dimensionalminimum-
time-to-climb (MTTC), minimum-fuel-to-climb (MFTC), minimum-time-interception
(MTI) and minimum-time-rendezvous (MTR) problems in constrained airspace have
been considered. The direct optimization method, direct collocation and nonlinear
programming (DCNLP), has advantages in terms of (1) the avoidance of guessing ini-
tial adjoint variables, (2) fast convergence and (3) the capacity of including complex
boundary conditions. Hence, it is considered to be the best method for solving these
minimum time problems. The Chebyshev Pseudospectral discretization method was
chosen as the primary collocation method due to its high accuracy and fast conver-
gence in interpolating between the collocation points.
The MTTC and MFTC trajectories are similar to each other under same bound-
ary conditions if the climb task is performed in a relatively short period time. Simi-
larities also exist between the effect of engagement conditions on the performance of
MTI and MTR trajectories. There are also common points between the MTTC and
MTI trajectories in the effect of start up velocity on the final time. Accurate results
and properties of the optimal trajectories are useful for the studies of the approximate
124
or analytical solutions for these problems. They also provide valuable guidance to
pilots regarding maneuver an aircraft how to achieve high performance.
Although the DCNLP method has many advantages in solving this kind of op-
timal control problems, the computation time is large enough that even with fast
computers it is probably not suitable for real-time applications. It follows that the
trajectories calculated here are flight planning trajectories which can be used as nomi-
nal for on-board guidance methods. This is specially true regarding the MTI problem,
when the target?s velocity is variable and trajectories need to be updated frequently
according to radar tracking data, more computation time will be needed. Future
research should focus on the development of simplified computation algorithms that
can be applied on-board.
125
Bibliography
[1] Bryson, A. E. and Denham, W. F., ?A Steepest-Ascent Method for Solving
Optimum Programming Programs,? Journal of Applied Mechanics, Vol. 29, pp.
247-257, June 1962.
[2] Bryson, A. E. and Desai , M. N. and Hoffman, W. C., ?Energy-State Approxima-
tion in Performance Optimization of Supersonic Aircraft,? Journal of Aircraft,
Vol. 6, No. 6, pp. 481-488, 1969.
[3] Calise, A. J., ?Extended Energy Management for Flight Performance Optimiza-
tion,? AIAA Journal, Vol. 15, No. 3, pp. 314-321, March 1977.
[4] Ardema, M. D., ?Solution of the Minimum-Time-to-Climb Problem by Matched
Asymptotic Expansions,? AIAA Journal, Vol. 14, No. 7, pp.843-850, 1976.
[5] Rao, A. V. and Mease, K. D., ?Minimum Time-to-Climb Trajectories Using a
Modified Sweep Method,? AIAA-95-3263-CP, 1995.
[6] Zarchan, P., ?Tactical and Strategic Missile Guidance,? AIAA Tactical Missile
Series, AIAA, Chap 2, pp. 11-29, 1997.
[7] Yuan, C.L., ?Homing and Navigation Courses of Automatic Target-Seeking De-
vices,? , Journal of Applied Physics, Vol 19, pp. 1122-1128, 1948.
[8] Adler, F. P., ?Missile Guidance by Three-Dimensional Proportional Navigation,?
Journal of Applied Physics, Vol. 27, No. 5, pp. 500-507, 1956.
[9] Duflos, E., Penel, P. and Vanheeghe, P., ?3D Guidance Law Modeling,? IEEE
Transactions on Aerospace and Electronic Systems, Vol. 35, No.1, pp. 72-83,
1999.
[10] Cochran, J. E. and No, T. S. and Thaxton, D. G., ?Analytical Solutions to a
Guidance Problem,? Journal of Guidance, Control, and Dynamics, Vol. 14, No.
1, pp. 117-122, 1991.
[11] Visser, H. G., Kelley, H. J. and Cliff, E. M., ?Energy Management of Three-
Dimensional Minimum-Time Intercept,? Journal of Guidance, Control, and Dy-
namics, Vol. 10, No. 6, pp. 574-580, 1987.
126
[12] Kumar, R. K., Seywald, H., and Cliff, E. M., ?Near-Optimal Three-Dimensional
Air-to-Air Missile Guidance Against Maneuvering Target,? Journal of Guidance,
Control, and Dynamics, Vol. 18, No. 3, pp. 457-464, 1995.
[13] Weston, A., Cliff, G. and Kelley, H., ?Onboard Near-Optimal Climb-Dash En-
ergy Management,? Journal of Guidance, Control, and Dynamics, Vol. 8, No. 3,
pp. 320-324, 1985.
[14] Bilimoria, K. and Cliff, E. M, ?Singular Trajectories in Airplane Cruise-Dash
Optimizaiton,? Journal of Guidance, Control, and Dynamics, Vol. 12, No. 3, pp.
304-310, 1989.
[15] Corban, J. E., Calise, A. J. and Flandro, G. A, ?Trajectory Optimization and
Guidance Law Development for Transatmospheric Vehicles,? Control and Appli-
cations, Proceedings. ICCON ?89. IEEE International Conference, pp. 460-465,
1989.
[16] Bollino, K. P., Ross, M. I. and Doman, D. D., ?Optimal Nonlinear Feedback
Guidance for Reentry Vehicles,? AIAA Guidance, Navigation and Control Con-
ference and Exhibit, AIAA 2006-6074, 2006.
[17] Newman, B., Britcher, C., Kassaye, Y., Krizansky, M. and Acheson, M., ?Trajec-
tory Management Concepts for Future Small Aircraft Transportation Systems,?
Journal of Aircraft, Vol. 43, No. 6, pp. 1643-1654, 2006.
[18] Betts, J. T., ?Survey of Numerical Methods for Trajectory Optimization,? Jour-
nal of Guidance, Control, and Dynamics, Vol. 21, No. 2, pp. 193-207, 1998.
[19] Hargraves, C. R. and Paris, S. W., ?Direct Trajectory Optimization Using Non-
linear Programming and Collocation,? Journal of Guidance, Control and Dy-
namics, Vol. 10, No. 4, pp 338-342, 1987.
[20] Betts, J. T. and Huffman, W. P., ?Path-Constrained Trajectory Optimization
UsingSparseSequentialQuadraticProgramming,? JournalofGuidance, Control,
and Dynamics, Vol. 16, No. 1, pp. 59-68, 1993.
[21] Ringertz, U., ?Optimal Trajectory for a Minimum Fuel Turn,? Journal of Air-
craft, Vol. 37, No. 5, pp. 932-934, 2000.
[22] Norsell M., ?Multistage Trajectory Optimization with Radar-Range Con-
straints,? Journal of Aircraft, Vol. 42, No. 4, pp. 849-857, 2005.
127
[23] Williams, P., Sgarioto, D. and Trivailo, P. M., ?Constrained Path-Planning for
an Aerial-Towed Cable System,? EUCASS 2005, 3.01.03, 2005.
[24] Herman, A. L. and Spencer, D. B., ?Optimal, Low-Thrust Earth-Orbit Transfers
Using Higher-Order Collocation Methods,? Journal of Guidance, Control, and
Dynamics, Vol. 25, No. 1, pp. 40-47, 2002.
[25] Horie, K. and Conway, B. A., ?Optimal, Aeroassisted Orbital Interception,?
Journal of Guidance, Control, and Dynamics, Vol. 22, No. 5, pp. 625-631, 1999.
[26] Betts, J. T., ?Trajectory Optimization in the Process of Uncertainty,? The Jour-
nal of the Astronautical Sciences , Vol. 54, No. 2, pp. 227-243, 2007.
[27] Geiger, B. R., Horn, J. F. DeLullo, A. M. AND Long, L. N., ?Optimal Path Plan-
ning of UAVs Using Direct Collocation with Nonlinear Programming,? AIAA
Guidance, Naviagtion, and Control Conference and Exhibit, AIAA 2006-6199,
2006.
[28] Yeo, B. P., ?An Extension in Optimal Initial Choice of Multipliers in the Quasi-
linearization Method,? International Journal of Control, Vol. 24, No. 5, pp. 593-
608, 1976.
[29] Bryson, A. E. and Ho, Y. C., ?Applied Optimal Control; Optimization, Estima-
tion, and Control,? Waltham, Mass., Ginn and Co, 1969.
[30] Hull, D. G., ?Optimal Control Theroy for Applications,? Mechanical Engineering
Series, Springer.
[31] Herman, A. L. and Spencer, D. B., ?Optimal, Low-Thrust Earth-Orbit Transfers
Using Higer-Order Collocation Methods,? Journal of Guidance, Control, and
Dynamics, Vol. 25, No. 1, pp. 40-47, 2002.
[32] Fahroo, M. and Ross, I. M., ?Direct Trajectory Optimization by a Chebychev
Pseudospectral Method,? Journal of Guidance, Control, and Dynamics, Vol. 25,
No. 1, pp. 160-166, 2002.
[33] Trefethen, L. N., Spectral Methods in MATLAB, Society for Industrial and Ap-
plied Mathematics, Philadelphia, 2000.
[34] Pietz, J. A., ?Pseudospectral Collocation Methods for the Direct Transcription of
Optimal Control Problems,? M.A. Thesis, Dept. of Computational and Applied
Mathematics, Rice University, Houston, TX, 2003.
128
[35] Xin, M., Balakrishnan, S.N. and Ohlmeyer, E.J.,?Guidance Law Design for Mis-
siles with Reduced Seeker Field-of-View,? AIAA Guidance, Navigation, and Con-
trol Conference and Exhibit, Keystone, Colorado, 2006.
[36] Manchester, I.R., Savkii, A.V. and Faruqi, F.A.x, ?Optical-flow based Precision
Missile Guidance inspired by Honeybee Navigation,? Proceedings of the 42nd
IEEE Conference on Decision and Control Maui, Hawaii USA, 2006.
[37] Betts, J. T., ?Practical Methods for Optimal Control Using Nonlinear Program-
ming,? Industrial and Applied Mathematics.
[38] Holmstrom, H., Goran, A.O. and Edvall, M. M, ?User?s Guide For TOMLAB
/SNOPT,? Tomlab Optimization Inc., 2005.
[39] Gano, S. E., Perez, V. M. and Renaud, J. E., ?Development and Verification
of A MATLAB Driver For The SNOPT Optimization Software,? AIAA Paper
2001-1620, 2001.
[40] Phillips, J. P., ?Brachistochrone, Tautochrone, Cycloid?Apple of Discord,?
Math. Teacher 60, pp.506-508, 1967.
[41] Zernelo, E., ?Uber das Navigationsproblen bei ruhender oder veranderlicher
Windverteilung,? Z. Angrew. Math. und. Mech., 1931.
[42] Powers, W. F., ?Hamiltonian Perturbation Theory for Optimal Trajectory Anal-
ysis,? Master Thesis, Dept. of Aerospace Engineering, The University of Texas,
Austin, TX, 1966.
[43] Etkin, B. and Reid, L. D., ?Dynamics of Flight: Stability and Control (3rd
edition),? John Wiley and Sons Inc, 1995.
[44] Dai, R. and Cochran, J. E., ?Three-Dimensional Minimum-Time-To-Climb in
Constrained Aerospace,? 45th AIAA Aerospace Sciences Meeting and Exhibit,
Reno, Nevada, 2007.
[45] Dai, R. and Cochran, J. E., Three-Dimensional Minimum-Time Interception Tra-
jectory Planning Using Nonlinear Programming and Collocation, AIAA Guid-
ance, Navigation and Control Conference and Exhibit, Hilton Head, SC, 2007.
129
Appendix A
Aircraft Propulsion and Aerodynamic Data
This Appendix presents the aircraft propulsion data effected by Altitude and
March number and aerodynamic data effected by March number only in tables.
Table A.1: Thrust as a function of altitude and Mach number from Ref.[2] for aircraft
2.
Thrust T (thousands of lb)
Mach Altitude h (thousands of ft)
No.M 0 5 15 25 35 45 55 65 75 85 95 105
0 23.3 20.6 15.4 9.9 5.8 2.9 1.3 0.7 0.3 0.1 0.1 0.0
0.4 22.8 19.8 14.4 9.9 6.2 3.4 1.7 1.0 0.5 0.3 0.1 0.1
0.8 24.5 22.0 16.5 12.0 7.9 4.9 2.8 1.6 0.9 0.5 0.3 0.2
1.2 29.4 27.3 21.0 15.8 11.4 7.2 3.8 2.7 1.6 0.9 0.6 0.4
1.6 29.7 29.0 27.5 21.8 15.7 10.5 6.5 3.8 2.3 1.4 0.8 0.5
2.0 29.9 29.4 28.4 26.6 21.2 14.0 28.7 5.1 3.3 1.9 1.0 0.5
2.4 29.9 29.2 28.4 27.1 25.6 17.2 10.7 6.5 4.1 2.3 1.2 0.5
2.8 29.8 29.1 28.2 26.8 25.8 3.4 1.7 1.0 0.5 0.3 0.1 0.5
3.2 29.7 28.9 27.5 26.1 24.9 20.3 13.0 8.0 4.9 2.8 1.4 0.5
Table A.2: Lift and drag coefficients as a function of angle of attack and Mach Number
for aircraft 2.
M 0 0.4 0.8 1.2 1.6 2.0 2.4 2.8 3.2
CL? 2.240 2.325 2.350 2.290 2.160 1.950 1.700 1.435 1.250
CD0 0.0065 0.0055 0.0060 0.0118 0.0110 0.0086 0.0074 0.0069 0.0068
S = 500ft2, W = 34200lb, ? = 1
130
Appendix B
Gradient Method for 2-D MTTC problem
This Appendix listed the calculation procedure for the 2-D MTTC problem.
1. Assume a nominal control load factor history n?(t), integrate the system differ-
ential equations forward from the starting point, store state variables history
V(t), ?(t) and h(t).
2. Derive the adjoint variables derivative equations as
??V = ??V (?T
?V ??VSCD0 +
1
2?V
2S?CD0
?V +
4n2W2
CL??V 3S +
2n2W2
C2L??V 2S
?CL?
?V )g/W
+?? gV 2(n?cos?)??h sin?
??? = ?Vgcos???? g
V sin???hV cos?
??h = ??V g
W
?T
?h
(B.1)
define the objective function
? = tf (B.2)
and the final conditions function as
? = h?hf = 0 (B.3)
131
There is no path constraints, so
? = 0 (B.4)
Then the lagrange multipliers with boundary conditions are defined as
?prime?(tf) = (???x)?t=t
f
= (???V ???? ???h)?t=t
f
= [0 0 0]
?prime?(tf) = (???x)?t=t
f
= (???V ???? ???h)?t=t
f
= [0 0 1]
(B.5)
Integrate ?? and ?? backward with stored state variables in step (1).
3. At the same time, calculate the quantities ??? and ??? defined as
?? = (???t + ???xf)?
t=tf = 1
?? = (???t + ???xf)?
t=tf =
?hf
?? = 0
??? = ?? ? ?????? = ????h
f
??? = ?? ? ?????? = [0 0 0]prime
(B.6)
4. Also, integrate the following numbers simultaneously with step(2)
I?? = integraltextTt0 ?prime??GW?1G???dt
I?? = integraltextTt0 ?prime??GW?1G???dt
I?? = integraltextTt0 ?prime??GW?1G???dt
(B.7)
132
where G is
G = ?f?n =
?
??
??
??
??
?
?f1
?n
?f2
?n
?f3
?n
?
??
??
??
??
?
=
?
??
??
??
??
?
? 4ngW?C
L?SV 2
?g
?V
0
?
??
??
??
??
?
(B.8)
and W is the pre-selected weighting matrix.
5. Select reasonable valuable dP and the correction of control variable value ?n is
calculated as
?n(t) = ?W?1Gprime???[(dP)
2
I?? ]
1
2 (B.9)
6. The new control variable history is obtained as
n(t) = n?(t) +?n(t) (B.10)
and repeate process (1)-(5) until convergent.
133
Appendix C
Proportional Navigation Guidance Law Algorithm
% This program is to use the Proportional Navigation Proportional
% law to simulate the three-dimensional intercepting trajectory
clear all close all
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
global K Vt
% set initial condition of target and interceptor
V_air_sea = 1116.9;
V_air_h = 1077.9;
Rt0 = [10000;-10000;10000];
Mt= 0.6;
Vty = V_air_h*Mt;
Vt = [0;Vty;0];
Rm0 = [0;-10000;0];
Mm = 0.4;
Vm0_abs = V_air_sea*Mm;
gama0 = 0.22;
kai0 = 0;
134
Vm0 = [Vm0_abs*sin(gama0);
Vm0_abs*cos(gama0)*cos(kai0);...
Vm0_abs*cos(gama0)*sin(kai0);];
% set the constant effective navigation ratio
K = 5;
% set integration step size
dt = 1;
% set integration initial condition
Rt = Rt0;
Rm = Rm0;
Vm = Vm0;
Rtm = Rt-Rm;
Vtm = Vt-Vm;
Rtm_abs =norm(Rtm);
Vc =-(Rtm(1)*Vtm(1)+Rtm(2)*Vtm(2)+Rtm(3)*Vtm(3))/Rtm_abs;
% integrate the velocity and coordinates forward using trapezoidal method
for i = 1:1000000
% set loop stop condition, that is when the close velocity changes its
% sign
if Vc<=0
135
break
end
if Rtm_abs <1000
dt=0.001;
else
dt=0.1;
end
Rt_old = Rt;
Rm_old = Rm;
Vm_old = Vm;
Rtm = Rt-Rm;
Vtm = Vt-Vm;
Rtm_abs = norm(Rtm);
Vc = -(Rtm(1)*Vtm(1)+Rtm(2)*Vtm(2)+Rtm(3)*Vtm(3))/Rtm_abs;
A = acc(Rm,Vm,Rt);
Rt = Rt+dt*Vt;
Rm = Rm+dt*Vm;
Vm = Vm+dt*A;
Rtm = Rt-Rm;
Vtm = Vt-Vm;
136
Rtm_abs = norm(Rtm);
Vc = -(Rtm(1)*Vtm(1)+Rtm(2)*Vtm(2)+Rtm(3)*Vtm(3))/Rtm_abs;
A = acc(Rm,Vm,Rt);
Rt = 0.5*(Rt_old+Rt+dt*Vt);
Rm = 0.5*(Rm_old+Rm+dt*Vm);
Vm = 0.5*(Vm_old+Vm+dt*A);
Rmx(i) = Rm(1)/10000;
Rmy(i) = Rm(2)/10000;
Rmz(i) = Rm(3)/10000;
Rtx(i) = Rt(1)/10000;
Rty(i) = Rt(2)/10000;
Rtz(i) = Rt(3)/10000;
if i==1
t(i) = dt;
else
t(i) = dt+t(i-1);
end
vi(i) = norm(Vm);
acel(i) = norm(A);
end
137
time = t(i-1);
fprintf(?The interception time usd is:\n[%3.6f] seconds\n ?, time);
fprintf(?The final miss distance is:\n[%3.6f] feet\n ?, Rtm_abs);
figure (1)
plot3(Rmx,Rmy,Rmz)
hold on
plot3(Rtx,Rty,Rtz)
xlabel(?X(10000ft)?)
ylabel(?Y (10000ft)?)
zlabel(?Altitude (10000ft)?)
figure(2)
plot(t,vi)
ylabel(?flight speed (ft/sec)?)
% This function calculate the acceleration command of the interceptor at
% each step time
function A=acc(Rm,Vm,Rt)
global K Vt
% calculate the relative position and relative velocity
Rtm=Rt-Rm;
Vtm=Vt-Vm;
138
% calculate the closing velocity and line-of-sight angular velocity
Rtm_abs=norm(Rtm);
dRtm=-(Rtm(1)*Vtm(1)+Rtm(2)*Vtm(2)+Rtm(3)*Vtm(3))/Rtm_abs;
omega=cross(Rtm,Vtm)/Rtm_abs^2;
% calculate the acceleration
A=-K*dRtm/Rtm_abs*cross(Rtm,omega);
139
Appendix D
Chebyshev Pseudospectral Collocation and Nonlinear Programming
Codes in Solving Brachistochrone Problem
% This is the major function which will use SNOPT to run the nonlinear
% programming solver
clear all Name = ?Brachistochrone Problem?;
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
global nc_defect ndiffeq nnodes nlp_state ncv nlpv x_0 xf2 xf1 x01
x02
% number of differential equations
ndiffeq = 2;
% number of control variables
ncv = 1;
% number of discretization nodes
nnodes = 25;
% number of state nlp variables
nlp_state = ndiffeq * nnodes;
% number of control nlp variables
nlp_control = ncv * nnodes;
140
% total number of nlp variables
nlpv = nlp_state + nlp_control
% number of state vector defect equality constraints
nc_defect = nlp_state;
% number of auxiliary equality constraints (boundary conditions)
nc_aux = 4;
% total number of equality constraints
nc_total = nc_defect + nc_aux;
% set the initial input of NLP variables
x01 = 0; % starting point
x02 = 0;
xf1 = 5; % ending point
xf2 = -1;
% guess the initial input of the control variables
u_guess=atan((xf2-x02)/(xf1-x01));
% guess the initial input of the state variables
for i=1:1:nnodes
x_0(2*i-1)=x01+(xf1-x01)/(nnodes-1)*(i-1);
x_0(2*i)=x02+(xf2-x02)/(nnodes-1)*(i-1);
x_0(nlp_state+i)=u_guess;
141
end
x_0(nlpv+1)=1.3; % guess the final time
% set the NLP variables bounds
for i=1:1:nnodes
x_L(2*i-1)=-20;
x_U(2*i-1)=20;
x_L(2*i)=-20;
x_U(2*i)=20;
x_L(nlp_state+i)=-2*pi;
x_U(nlp_state+i)=2*pi;
end
x_L(nlpv+1)=0;
x_U(nlpv+1)=10;
% set the nonlinear function constraints
for i=1:1:nc_total
c_L(i)=0;
c_U(i)=0;
end fLowBnd = 0;
Prob = conAssign(?trapm3_f?, [], [], [], x_L, x_U,Name, x_0,[],
fLowBnd, [], [], [], ?trapm3_cbr?, [], [], [], c_L, c_U);
142
Prob.Warning = 0; % Turning off warnings.
Result = tomRun(?snopt?, Prob, 1);
X_OUT= Result.x_k;
for i=1:1:nnodes
yff(i)=X_OUT(2*i);
xff(i)=X_OUT(2*i-1);
uff(i)=X_OUT(nlp_state+i);
end
figure (1) % plot the trajectory
plot(xff,yff)
function c = trapm3_cbr (x,prob)
% equality constraints
% Chebyshev pseudospectral collocation method
% inputs: x = current nlp variable values
% outputs: c = vector of nonlinear equality constraints evaluated at x
%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%
global nc_defect ndiffeq nnodes nlp_state ncv nlpv xf2 xf1 x01 x02
% compute state vector defect equality constraints
t_total = x(nlpv+1);
if t_total==0
143
t_total=0.001;
end
for k = 1:1:nnodes
% state vector elements
% reset to previous node
nks = (k - 1) * ndiffeq;
for i = 1:1:ndiffeq
xk(i) = x(nks + i);
end
% control variable elements
% reset to previous node
nkc = nlp_state + (k - 1) * ncv;
for i = 1:1:ncv
uk(i) = x(nkc + i);
end
% save one node states in one column of a matrix
sx(:,(nnodes-k+1)) = xk?;
% compute state vector defects for current node
f = deriv2 (xk, uk);
% save one node f value in one column of a matrix
144
sf(:,(nnodes-k+1)) = f?;
end
% calculte the Chebyshev-Gauss-Lobatto differentiation matrix D
N = nnodes-1;
t = cos(pi*(0:N)/N)?;
p = [2; ones(N-1,1); 2].*(-1).^(0:N)?;
T = repmat(t,1,N+1);
dT = T-T?;
D = (p*(1./p)?)./(dT+(eye(N+1))); % off-diagonal entries
D = D - diag(sum(D?)); % diagonal entries
% compute the defects array for current node
for k=1:1:nnodes
nks = (k - 1) * ndiffeq;
for i=1:1:ndiffeq
d(i,k) = D(k,:)*sx(i,:)?; % calculate the derivative of f at node k
resid(nks + i) = 2/t_total*d(i,k)-sf(i,k);
end
end
% set active defect constraints
% (offset by 1)
145
for i = 1:1:nc_defect
c(i) = resid(i);
end
% current final state vector
xfinal(1) = x(nlp_state - 1);
xfinal(2) = x(nlp_state);
% current initial state vector
xinitial(1) = x(1);
xinitial(2) = x(2);
% initial boundary conditions
c(nc_defect + 1) = xinitial(1)-x01;
c(nc_defect + 2)=xinitial(2)-x02;
% final boundary conditions
c(nc_defect + 3) = xfinal(1)-xf1;
c(nc_defect + 4) = xfinal(2)-xf2;
c = c?; % transpose
function f = trapm3_f (x,prob)
% objective function
% inputs: x = current nlp variable values
% outputs: f = objective function evaluated at x
146
global nlpv
tfinal = x(nlpv+1);
f = tfinal; % objective function (maximize final time)
function xdot=deriv2(x,u)
% equations of motion
% input: x = current state vector, u = current control vector
% output: xdot = derivative of x and y
global x02
% constant parameter
g = 9.82;
% evaluate equations of motion at current conditions
xdot(1) = sqrt(2*g*(x02-x(2)))*cos(u);
xdot(2) = sin(u)*sqrt(2*g*(x02-x(2)));
147