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1
Jones, Campbell Llyr. "Neural control of a sea skimming missile." Thesis, University of Southampton, 1996. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.387900.
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Vural, Ozgur Ahmet. "Fuzzy Logic Guidance System Design For Guided Missiles." Master's thesis, METU, 2003. http://etd.lib.metu.edu.tr/upload/1026715/index.pdf.
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This thesis involves modeling, guidance, control, and flight simulations of a canard
controlled guided missile.
The autopilot is designed by a pole placement technique. Designed autopilot is
used with the guidance systems considered in the thesis.
Five different guidance methods are applied in the thesis, one of which is the
famous proportional navigation guidance. The other four guidance methods are
different fuzzy logic guidance systems designed considering different types of
guidance inputs.
Simulations are done against five different target types and the performances of the
five guidance methods are compared and discussed.
3
Lai, Haoyu. "On the design of nonlinear gain scheduled control systems." Ohio : Ohio University, 1998. http://www.ohiolink.edu/etd/view.cgi?ohiou1176486900.
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Doruk, Resat Ozgur. "Missile Autopilot Design By Projective Control Theory." Master's thesis, METU, 2003. http://etd.lib.metu.edu.tr/upload/4/1089929/index.pdf.
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In this thesis, autopilots are developed for missiles with moderate dynamics and
stationary targets. The aim is to use the designs in real applications. Since the real
missile model is nonlinear, a linearization process is required to get use of systematic
linear controller design techniques. In the scope of this thesis, the linear quadratic
full state feedback approach is applied for developing missile autopilots. However,
the limitations of measurement systems on the missiles restrict the availability of all
the states required for feedback. Because of this fact, the linear quadratic design will
be approximated by the use of projective control theory. This method enables the
designer to use preferably static feedback or if necessary a controller plus a low
order compensator combination to approximate the full state feedback reference.
Autopilots are checked for the validity of linearization, robust stability against
aerodynamic, mechanical and measurement uncertainties.
5
Bibel, John Eugene. "Missile autopilot design using Mu-Synthesis." Thesis, This resource online, 1998. http://scholar.lib.vt.edu/theses/available/etd-08252008-161841/.
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Counsell, John Mark. "Optimum and safe control algorithim (OSCA) for modern missile autopilot design." Thesis, Lancaster University, 1992. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.332382.
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White, David Paul. "Missile autopilot design using a gain scheduling technique." Ohio : Ohio University, 1994. http://www.ohiolink.edu/etd/view.cgi?ohiou1179860306.
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8
Ozkan, Bulent. "Dynamic Modeling, Guidance, And Control Of Homing Missiles." Phd thesis, METU, 2005. http://etd.lib.metu.edu.tr/upload/12606533/index.pdf.
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DYNAMIC MODELING, GUIDANCE, AND CONTROL OF HOMING MISSILES
Ö<br>ZKAN, Bü<br>lent
Ph. D., Department of Mechanical Engineering
Supervisor: Prof. Dr. M. Kemal Ö<br>ZGÖ<br>REN
Co-Supervisor: Dr. Gö<br>kmen MAHMUTYAZICIOgLU
September 2005, 236 pages
In this study, the dynamic modeling, guidance, and control of a missile with two relatively rotating parts are dealt with. The two parts of the missile are connected to each other by means of a roller bearing. In the first part of the study, the governing differential equations of motion of the mentioned missile are derived. Then, regarding the relative rotation between the bodies, the aerodynamic model of the missile is constructed by means of the Missile Datcom software available in TÜ<br>BiTAK-SAGE. After obtaining the required aerodynamic stability derivatives using the generated aerodynamic data, the necessary transfer functions are determined based on the equations of motion of the missile. Next, the guidance laws that are considered in this study are formulated. Here, the Linear Homing Guidance and the Parabolic Homing Guidance laws are introduced as alternatives to the Proportional Navigation Guidance law. On this occasion, the spatial derivation of the Proportional Navigation Guidance law is also done. Afterwards, the roll, pitch and yaw autopilots are designed using the determined transfer functions. As the roll autopilot is constructed to regulate the roll angle of the front body of the missile which is the controlled part, the pitch and yaw autopilots are designed to realize the command signals generated by the guidance laws. The guidance commands are in the form of either the lateral acceleration components or the flight path angles of the missile. Then, the target kinematics is modeled for a typical surface target. As a complementary part of the work, the design of a target state estimator is made as a first order fading memory filter. Finally, the entire guidance and control system is built by integrating all the models mentioned above. Using the entire system model, the computer simulations are carried out using the Matlab-Simulink software and the proposed guidance laws are compared with the Proportional Navigation Guidance law. The comparison is repeated for a selected single-body missile as well. Consequently, the simulation results are discussed and the study is evaluated.
9
Sefastsson, Ulf. "Evaluation of Missile Guidance and Autopilot through a 6 DOF Simulation Model." Thesis, KTH, Optimeringslära och systemteori, 2016. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-188897.
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Missile guidance and autopilot have been active fields of research since the second world war. There are lots of literature on the subjects, but the bulk of which are confined to overly simplified models, and therefore the publications of the methods applied to more realistic models are scarce. In this report a nonlinear 6 DOF simulation model of a tail-controlled air-to-air missile is considered. Through several assumptions and simplifications a linearized approximation of the plant is obtained, which then is used in the implementation of 5 guidance laws and 2 autopilots. The guidance laws are all based on a linearized collision geometry, and the autopilots are based on model predictive control (MPC). Both autopilots use linear quadratic MPC (LQMPC), and one is more robust to modelling errors than the conventional LQMPC. The guidance laws and autopilots are then evaluated with respect to performance in terms of miss distance in 4 interception scenarios with a moving target. The results show that the in this model the autopilots perform equally well, and that the guidance laws with more information about the target generally exhibit smaller miss distances, but at the cost of a considerably larger flight time for some scenarios. The conclusions are that the simplifying assumptions in the modelling are legitimate and that the challenges of missile control probably does not lie in the guidance or autopilot, but rather in the target tracking. Therefore it is suggested that future work include measurement noise and process disturbances in the model.<br>Det har forskats kring styrlagarna och styrautomaterna för robotar sedan an-dra världskrigets. Det finns mycket litteratur på områdena, men merparten av de publicerade resultaten behandlar enbart grovt förenklade modeller, och därför är tillgången på publikationer där metoderna applicerats i en mer realistisk modell begränsat. I denna rapport behandlas en olinjär simuleringsmodell av en jaktrobot som styrs med stjärtfenor och har sex frihetsgrader. Genom en rad antaganden och förenklingar erhålls en linjäriserad modell av missilen, vilket sedan används för implementering av fem styrlagar och två styrautomater. Styr-lagarna är alla baserade på en linjäriserad kollisionsgeometri och styrautomaterna är baserade på modellprediktiv styrning (MPC). Båda styrautomaterna använder linjärkvadratisk MPC, där den ena påstås vara mer robust gentemot modellfel. Styrlagarna och -automaterna utvärderas ur ett prestandaperspektiv med fokus på bomavstånd i fyra realistiska genskjutningsscenarier med ett rörligt mål. Resultaten visar att båda styrautomaterna presterar lika bra, och att de styrlagar med mer information om målets position/hastighet/acceleration generellt presterar bättre, men att de för vissa skjutfall får en väsentligt längre flygtid. Slutsatserna är att förenklingarna och antagandena i linjäriseringen är välgrundade, och att utmaningarna i missilstyrning inte ligger i utformning av styrlag/-automat, utan förmodligen i målsökningen. Därför föreslås det slutligen att framtida arbete bl. a. inkluderar mätbrus och störningar.
10
Sharma, Manu. "A neuro-adaptive autopilot design for guided munitions." Diss., Georgia Institute of Technology, 2001. http://hdl.handle.net/1853/12470.
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DeMott, Robert. "Development of a Flexible FPGA-Based Platform for Flight Control System Research." VCU Scholars Compass, 2010. http://scholarscompass.vcu.edu/etd/2321.
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This work is part of ongoing research conducted at Virginia Commonwealth University relating to unmanned aerial vehicles. The primary objective of this thesis was to develop a flexible, high-performance autopilot platform in order to facilitate research on advanced flight control algorithms. A dual FPGA-based system architecture utilizing a stacked, multi-board design was created to meet this goal. Processing tasks were split between the two FPGA devices, allowing for improved system timing and increased throughput. A combination of analog and digital filtering techniques were employed in the new system, resulting in enhanced sensor accuracy and precision compared to the previous generation autopilot system. Several important improvements to the safety and reliability of the overall system were also achieved.
12
Snyder, Mark. "NONLINEAR STABILIZATION AND CONTROL OF MEDIUM RANGE SURFACE TO AIR INTERCEPTOR MISSILES." Master's thesis, University of Central Florida, 2009. http://digital.library.ucf.edu/cdm/ref/collection/ETD/id/4081.
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Nonlinear stabilization and control autopilots are capable of sustaining nominal performance throughout the entire fight envelope an interceptor missile may encounter during hostile engagements and require no gain scheduling to maintain autopilot stability. Due to non minimum phase conditions characteristic of tail controlled missile airframes, a separation of time scales within the dynamic equations of motion between rotational and translational differential equations was enforced to overcome unstable effects of non minimum phase. Dynamic inversion techniques are then applied to derive linearizing equations which, when injected forward into the plant result in a fully controllable linear system. Objectives of the two time scale control architecture are to stabilize vehicle rotational rates while at the same time controlling acceleration within the lateral plane of the vehicle under rapidly increasing dynamic pressure. Full 6 degree of freedom dynamic terms including all coriolis accelerations due to translational and rotational dynamic coupling have been taken into account in the inversion process. The result is a very stable, nonlinear autopilot with fixed control gains fully capable of stable nonlinear missile control. Several actuator systems were also designed to explore the destabilizing effects second order nonlinear actuator characteristics can have on nonlinear autopilot designs.<br>M.S.E.E.<br>School of Electrical Engineering and Computer Science<br>Engineering and Computer Science<br>Electrical Engineering MSEE
13
Kahvecioglu, Alper. "A Tool For Designing Robust Autopilots For Ramjet Missiles." Master's thesis, METU, 2006. http://etd.lib.metu.edu.tr/upload/12607058/index.pdf.
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The study presented in this thesis comprises the development of the longitudinal autopilot algorithm for a ramjet powered air-to-surface missile. Ramjet Missiles have short time-of-flight, however they suffer from limited angle of attack margins due to poor operational-region characteristics of the ramjet engine. Because of such limitations and presence of uncertainties involved, Robust Control Techniques are used for the controller design. Robust Control Techniques not only provide an easy limitation/uncertainty/performance handling for MIMO systems, but also, robust controllers promise stability and performance even in the presence of uncertainties of a pre-defined class. All the design process is carried out in such a way that at the end of the study a tool has been developed, that can process raw aerodynamic data obtained by Missile DATCOM program, linearize the equations of motion, construct the system structure and design sub-optimal H&<br>#8734<br>controllers to meet the requirements provided by the user. An autopilot which is designed by classical control techniques is used for performance and robustness comparison, and a non-linear simulation is used for validation. It is concluded that the code, which is very easy to modify for the specifications of other missile systems, can be used as a reliable tool in the preliminary design phases where there exists uncertainties/limitations and still can provide satisfactory results while making the design process much faster.
14
Evcimen, Cagdas. "Development And Comparison Of Autopilot And Guidance Algorithms For Missiles." Master's thesis, METU, 2007. http://etd.lib.metu.edu.tr/upload/2/12608677/index.pdf.
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In order to have an interception with a target, a missile should be guided with a successful guidance algorithm accompanied with a suitable autopilot structure. In this study, different autopilot and guidance designs for a canard-controlled missile
are developed. As a first step, nonlinear missile mathematical model is derived by using the equations of motion with aerodynamic coefficients found by Missile DATCOM program.
Autopilot design starts by the linearization of the nonlinear missile model around equilibrium flight conditions. Controllers based on the concepts of optimal control theory results and sliding mode control are designed. In all of the designs, angle
of attack command and roll angle command type autopilot structures are used. During the design process, variations in angle of attack, Mach number and altitude can lead to significant performance degradation. This problem is typically solved by applying gain-scheduling methodology according to these parameters.
There are different types of guidance methods in the literature. Throughout this study, proportional navigation guidance and its modified forms are selected as a base algorithm in the guidance system design. Other robust forms of guidance methods, such as an optimal guidance approach and sliding mode guidance, are also formed for performance comparison with traditional proportional navigation guidance approach. Finally, a new guidance method, optimal proportional-integral guidance, whose performance is the best among all of the methods included in the thesis against highly maneuvering targets, is introduced.
15
Aydin, Gunes. "Aerodynamic Parameter Estimation Of A Missile In Closed Loop Control And Validation With Flight Data." Master's thesis, METU, 2012. http://etd.lib.metu.edu.tr/upload/12615044/index.pdf.
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Aerodynamic parameter estimation from closed loop data has been developed as another research area since control and stability augmentation systems have been mandatory for aircrafts. This thesis focuses on aerodynamic parameter estimation of an air to ground missile from closed loop data using separate surface excitations. A design procedure is proposed for designing separate surface excitations. The effect of excitations signals to the system is also analyzed by examining autopilot disturbance rejection performance. Aerodynamic parameters are estimated using two different estimation techniques which are ordinary least squares and complex linear regression. The results are compared with each other and with the aerodynamic database. An application of the studied techniques to a real system is also given to validate that they are directly applicable to real life.
16
Ng, Chuk Man 1974. "On the digital re-design of an analogue missile flight control system using PIM method." Thesis, McGill University, 1999. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=30263.
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This research examines some of the ways to design a digital control system with particular reference to the control of yaw plane dynamics of an air-to-air missile model. Specific attention is paid to compare two methods of global digital re-design of validated analogue closed loop control system. These two methods are the multi-input multi-output (MIMO) Plant Input Mapping (PIM) method and multi-loop PIM method. The thesis will first show how a specific class of closed loop MIMO feedback systems with a single-input multi-output (SIMO) plant can be re-structured as multi-loop control systems, which includes the closed loop missile control system. Some special characteristics pertaining to SIMO systems for control system design are explored. The aforementioned two methods of PIM global digital re-design are then applied to the MIMO analogue missile control system and its re-structured multi-loop counterpart. A comparison of the two methods is made on its design procedure, implementation structure and results from computer simulation.<br>This thesis also touches upon the topic of controller order reduction, particularly in consideration of the PIM digital re-design of analogue feedback systems. (Abstract shortened by UMI.)
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Ng, Chuk Man. "On the digital re-design of an analogue missile flight control system using PIM method." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 1999. http://www.collectionscanada.ca/obj/s4/f2/dsk1/tape4/PQDD_0030/MQ64239.pdf.
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Du, Yongliang. "Development of real-time flight control system for low-cost vehicle." Thesis, Cranfield University, 2011. http://dspace.lib.cranfield.ac.uk/handle/1826/8621.
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In recent years, more and more light aircraft enter our daily life, from Agricultural
applications, emergency rescue, flight experiment and training to Barriers to
entry, light aircraft always have their own advantages. Thus, they have become
more and more popular.
However, in the process of GDP research about Flight Control System design
for the Flying Crane, the author read a lot of literature about Flight Control
System design, then noticed that the research in Flight Control System have
apparently neglected to Low-cost vehicles. So it is necessary to do some study
about Flight Control System for this kind of airplane. The study will more
concern the control law design for ultra-light aircraft, the author hopes that with
an ‘intelligence’ Flight Control System design, this kind of aircraft could
sometimes perform flying tasks according to a prearranged flight path and
without a pilot.
As the Piper J-3 cub is very popular and the airframe data can be obtained
more easily, it was selected as an objective aircraft for the control law design.
Finally, a ¼ scale Piper J-3 cub model is selected and the aerodynamics
coefficients are calculated by DATCOM and AVL. Based on the forces and
moments acting on the aircraft, the trim equilibrium was calculated for getting
proper dynamics coefficients for the selected flight conditions. With the aircraft
aerodynamics coefficients, the aircraft dynamics characteristics and flying
qualities are also analyzed. The model studied in this thesis cannot answer
level one flying qualities in the longitudinal axis, which is required by MIL-F-
8785C. The stability augment system is designed to improve the flying qualities
of the longitudinal axis.
The work for autopilot design in this thesis includes five parts. First, the whole
flight profile is designed to automatically control aircraft from takeoff to landing.
Second, takeoff performance and guidance law is studied. Then, landing
performance and trajectory is also investigated. After that, the control law
design is decoupled into longitudinal axis and later-directional axis. Finally,
simulation is executed to check the performance for the auto-controller.
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Sleeman, William Clifford IV. "The Development of a Linux and FPGA Based Autopilot System for Unmanned Aerial Vehicles." VCU Scholars Compass, 2007. http://scholarscompass.vcu.edu/etd_retro/129.
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This project is part of research funded by NASA Langley in field of Unmanned Aerial Vehicles (UAVs) and is based on past work conducted at Virginia Commonwealth University. Dr. Mark A. Motter of NASA Langley intends to use the new autopilot system to test aircraft with many control surfaces. The goal of this project is to port an existing UAV autopilot system that has more computing power than the previous generation system to allow for more advanced flight control algorithms.The steps taken to complete this project include choosing a new hardware platform, porting C flight control software from a MicroBlaze platform to a PowerPC platform, and developing FPGA based hardware to interface with external sensors. The Suzaku-V based system was shown to have much better computing performance than the previous system, and several successful test flights have proved the viability of the new autopilot system.
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Kargin, Volkan. "Design Of An Autonomous Landing Control Algorithm For A Fixed Wing Uav." Master's thesis, METU, 2007. http://etd.lib.metu.edu.tr/upload/12608996/index.pdf.
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This thesis concerns with the design and development of automatic flight controller strategies
for the autonomous landing of fixed wing unmanned aircraft subject to severe environmental
conditions. The Tactical Unmanned Aerial Vehicle (TUAV) designed at the Middle East Technical University (METU) is used as the subject platform. In the first part of this thesis, a
dynamic model of the TUAV is developed in FORTRAN environment. The dynamic model is
used to establish the stability characteristics of the TUAV. The simulation model also incorporates ground reaction and atmospheric models. Based on this model, the landing trajectory
that provides shortest landing distance and smallest approach time is determined. Then, an
automatic flight control system is designed for the autonomous landing of the TUAV. The
controller uses a model inversion approach based on the dynamic model characteristics. Feed
forward and mixing terms are added to increase performance of the autopilot. Landing strategies are developed under adverse atmospheric conditions and performance of three different
classical controllers are compared. Finally, simulation results are presented to demonstrate the
effectiveness of the design. Simulation cases include landing under crosswind, head wind, tail
wind, wind shear and turbulence.
21
Peddle, Iain K. "Acceleration based manoeuvre flight control system for unmanned aerial vehicles." Thesis, Stellenbosch : Stellenbosch University, 2008. http://hdl.handle.net/10019.1/1172.
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Thesis (PhD (Electrical and Electronic Engineering))--Stellenbosch University, 2008.<br>A strategy for the design of an effective, practically feasible, robust, computationally efficient
autopilot for three dimensional manoeuvre flight control of Unmanned Aerial Vehicles is
presented. The core feature of the strategy is the design of attitude independent inner loop
acceleration controllers. With these controllers implemented, the aircraft is reduced to a point
mass with a steerable acceleration vector when viewed from an outer loop guidance
perspective. Trajectory generation is also simplified with reference trajectories only required
to be kinematically feasible. Robustness is achieved through uncertainty encapsulation and
disturbance rejection at an acceleration level.
The detailed design and associated analysis of the inner loop acceleration controllers is carried
out for the case where the airflow incidence angles are small. For this case it is shown that
under mild practically feasible conditions the inner loop dynamics decouple and become
linear, thereby allowing the derivation of closed form pole placement solutions. Dimensional
and normalised non-dimensional time variants of the inner loop controllers are designed and
their respective advantages highlighted. Pole placement constraints that arise due to the
typically weak non-minimum phase nature of aircraft dynamics are developed.
A generic, aircraft independent guidance control algorithm, well suited for use with the inner
loop acceleration controllers, is also presented. The guidance algorithm regulates the aircraft
about a kinematically feasible reference trajectory. A number of fundamental basis trajectories
are presented which are easily linkable to form complex three dimensional manoeuvres.
Results from simulations with a number of different aircraft and reference trajectories illustrate
the versatility and functionality of the autopilot.
Key words: Aircraft control, Autonomous vehicles, UAV flight control, Acceleration control,
Aircraft guidance, Trajectory tracking, Manoeuvre flight control.
22
Náglová, Katarína. "Konvenční a adaptivní metody nastavení parametrů regulátoru pro řízení letové výšky letadla s pohyblivou pozicí těžiště v prostředí MATLAB - Simulink." Master's thesis, Vysoké učení technické v Brně. Fakulta elektrotechniky a komunikačních technologií, 2013. http://www.nusl.cz/ntk/nusl-219931.
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The thesis is focused on aircraft flight level control in case when the location of center of gravity changes during flight. The theoretical part of the thesis describes basics of aerodynamics, which are necessary to understand airplane's behavior and its control theory. Basics of state theory and stability of systems, as it is used for design of linear model. This part also mentions mathematical equations of nonlinear model of business jet and discusses conventional and adaptive methods used to design autopilot parameters. Several autopilots were designed in practical part of the thesis. The most important part is the logic used to determine location of center of gravity of the airplane. The models use nonlinear model of an airplane that better represents the real environment and conditions.
23
Cakir, Zeynep. "Development Of A Uav Testbed." Master's thesis, METU, 2011. http://etd.lib.metu.edu.tr/upload/12613209/index.pdf.
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The development and testing for a UAV testbed to be used in academic research and undergraduate education is proposed in this thesis. Analysis on commercial off-the-shelf UAV systems and autopilots lead to the development of a custom, open-architecture and modular UAV testbed. The main focus is to support research in UAV control field and education of the undergraduate students. The integration and use of commercial-off-the-shelf avionics and air vehicle are described in detail. System performance is examined both in flight and on the ground. Results of the system tests show that the developed system is a functional UAV testbed to be used in research of different flight control algorithms.
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DeBoy, Christopher C., Paul D. Schwartz, and Richard K. Huebschman. "Midcourse Space Experiment Spacecraft and Ground Segment Telemetry Design and Implementation." International Foundation for Telemetering, 1996. http://hdl.handle.net/10150/608390.
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International Telemetering Conference Proceedings / October 28-31, 1996 / Town and Country Hotel and Convention Center, San Diego, California<br>This paper reviews the performance requirements that provided the baseline for
development of the onboard data system, RF transmission system, and ground segment
receiving system of the Midcourse Space Experiment (MSX) spacecraft. The onboard
Command and Data Handling (C&DH) System was designed to support the high data
outputs of the three imaging sensor systems onboard the spacecraft and the requirement for
large volumes of data storage. Because of the high data rates, it was necessary to construct
a dedicated X-band ground receiver system at The Johns Hopkins University Applied
Physics Laboratory (APL) and implement a tape recorder system for recording and
downlinking sensor and spacecraft data. The system uses two onboard tape recorders to
provide redundancy and backup capabilities. The storage capability of each tape recorder
is 54 gigabits. The MSX C&DH System can record data at 25 Mbps or 5 Mbps. To meet
the redundancy requirements of the high-priority experiments, the data can also be
recorded in parallel on both tape recorders. To provide longer onboard recording, the data
can also be recorded serially on the two recorders. The reproduce (playback) mode is at
25 Mbps. A unique requirement of the C&DH System is to multiplex and commutate the
different output rates of the sensors and housekeeping signals into a common data stream
for recording. The system also supports 1-Mbps real-time sensor data and 16-kbps real-time
housekeeping data transmission to the dedicated ground site and through the U.S. Air
Force Satellite Control Network ground stations. The primary ground receiving site for the
telemetry is the MSX Tracking System (MTS) at APL. A dedicated 10-m X-band antenna
is used to track the satellite during overhead passes and acquire the 25-Mbps telemetry
downlinks, along with the 1-Mbps and 16-kbps real-time transmissions. This paper
discusses some of the key technology trade-offs that were made in the design of the system
to meet requirements for reliability, performance, and development schedule. It also presents some of the lessons learned during development and the impact these lessons will
have on development of future systems.
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Tanil, Cagatay. "Optimal External Configuration Design Of Missiles." Master's thesis, METU, 2009. http://etd.lib.metu.edu.tr/upload/12610873/index.pdf.
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The main area of emphasis in this study is to investigate the methods and technology for aerodynamic configuration sizing of missiles and to develop a software platform in MATLAB®<br>environment as a design tool which has an ability of optimizing the external configuration of missiles for a set of flight requirements specified by the user through a graphical user interface. A genetic algorithm based optimization tool is prepared by MATLAB is expected to help the designer to find out the best external geometry candidates in the conceptual design stage. Missile DATCOM software package is employed to predict the aerodynamic coefficients needed in finding the performance merits of a missile for each external geometry candidate by integrating its dynamic equations of motion.
Numerous external geometry candidates are rapidly eliminated according to objectives and constraints specified by designers, which provide necessary information in preliminary design. In this elimination, the external geometry candidates are graded according to their flight performances in order to discover an optimum solution.
In the conceptual design, the most important performance objectives related to the external geometry of a missile are range, speed, maneuverability, and control effectiveness. These objectives are directly related to the equations of motion of the missile, concluding that the speed and flight range are related to the total mass and the drag-to-lift ratio acting on missile. Also, maneuverability depends on the normal force acting on missile body and mass whereas the control effectiveness is affected by pitching moment and mass moment of inertia of missile. All of the flight performance data are obtained by running a two degree-of-freedom simulation.
In order to solve the resulting multi-objective optimization problem with a set of constraint of linear and nonlinear nature and in equality and inequality forms, genetic-algorithm-based methods are applied. Hybrid encoding methods in which the integer configuration variables (i.e., nose shape and control type) and real-valued geometrical dimension (i.e., diameter, length) parameters are encoded in the same individual chromosome.
An external configuration design tool (EXCON) is developed as a synthesis and external sizing tool for the subsonic cruise missiles. A graphical user interface (GUI), a flight simulator and optimization modules are embedded into the tool. A numerical example, the re-configuration problem of an anti-ship cruise missile Harpoon, is presented to demonstrate the accuracy and feasibility of the conceptual design tool. The optimum external geometries found for different penalty weights of penalty terms in the cost function are compared according to their constraint violations and launch mass values. By means of using EXCON, the launch mass original baseline Harpoon is reduced by approximately 30% without deteriorating the other flight performance characteristics of the original Harpoon.
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Mattei, Giovanni. "Robust nonlinear control : from continuous time to sampled-data with aerospace applications." Thesis, Paris 11, 2015. http://www.theses.fr/2015PA112025/document.
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La thèse porte sur le développement des techniques non linéaires robustes de stabilisation et commande des systèmes avec perturbations de model. D’abord, on introduit les concepts de base de stabilité et stabilisabilité robuste dans le contexte des systèmes non linéaires. Ensuite, on présente une méthodologie de stabilisation par retour d’état en présence d’incertitudes qui ne sont pas dans l’image de la commande («unmatched»). L’approche récursive du «backstepping» permet de compenser les perturbations «unmatched» et de construire une fonction de Lyapunov contrôlée robuste, utilisable pour le calcul ultérieur d’un compensateur des incertitudes dans l’image de la commande («matched»). Le contrôleur obtenu est appelé «recursive Lyapunov redesign». Ensuite, on introduit la technique de stabilisation par «Immersion & Invariance» comme outil pour rendre un donné contrôleur non linéaire, robuste par rapport à dynamiques non modelées. La première technique de contrôle non linéaire robuste proposée est appliquée au projet d’un autopilote pour un missile air-air et au développement d’une loi de commande d’attitude pour un satellite avec appendices flexibles. L’efficacité du «recursive Lyapunov redesign» est mis en évidence dans le deux cas d’étude considérés. En parallèle, on propose une méthode systématique de calcul des termes incertains basée sur un modèle déterministe d’incertitude. La partie finale du travail de thèse est relative à la stabilisation des systèmes sous échantillonnage. En particulier, on reformule, dans le contexte digital, la technique d’Immersion et Invariance. En premier lieu, on propose des solutions constructives en temps continu dans le cas d’une classe spéciale des systèmes en forme triangulaire «feedback form», au moyen de «backstepping» et d’arguments de domination non linéaire. L’implantation numérique est basée sur une loi multi-échelles, dont l’existence est garantie pour la classe des systèmes considérée. Le contrôleur digital assure la propriété d’attractivité et des trajectoires bornées. La loi de commande, calculée par approximation finie d’un développement asymptotique, est validée en simulation de deux exemples académiques et deux systèmes physiques, le pendule inversé sur un chariot et le satellite rigide<br>The dissertation deals with the problems of stabilization and control of nonlinear systems with deterministic model uncertainties. First, in the context of uncertain systems analysis, we introduce and explain the basic concepts of robust stability and stabilizability. Then, we propose a method of stabilization via state-feedback in presence of unmatched uncertainties in the dynamics. The recursive backstepping approach allows to compensate the uncertain terms acting outside the control span and to construct a robust control Lyapunov function, which is exploited in the subsequent design of a compensator for the matched uncertainties. The obtained controller is called recursive Lyapunov redesign. Next, we introduce the stabilization technique through Immersion \& Invariance (I\&I) as a tool to improve the robustness of a given nonlinear controller with respect to unmodeled dynamics. The recursive Lyapunov redesign is then applied to the attitude stabilization of a spacecraft with flexible appendages and to the autopilot design of an asymmetric air-to-air missile. Contextually, we develop a systematic method to rapidly evaluate the aerodynamic perturbation terms exploiting the deterministic model of the uncertainty. The effectiveness of the proposed controller is highlighted through several simulations in the second case-study considered. In the final part of the work, the technique of I\& I is reformulated in the digital setting in the case of a special class of systems in feedback form, for which constructive continuous-time solutions exist, by means of backstepping and nonlinear domination arguments. The sampled-data implementation is based on a multi-rate control solution, whose existence is guaranteed for the class of systems considered. The digital controller guarantees, under sampling, the properties of manifold attractivity and trajectory boundedness. The control law, computed by finite approximation of a series expansion, is finally validated through numerical simulations in two academic examples and in two case-studies, namely the cart-pendulum system and the rigid spacecraft
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Novák, Jiří. "Návrh autopilota a letových řídících módů v prostředí Simulink." Master's thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2020. http://www.nusl.cz/ntk/nusl-416616.
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Tato diplomová práce je zaměřena na vývoj simulačního prostředí v Matlab/Simulink zvoleného letadla ve známém letovém režimu. Pozice a orientace letadla pohybujícího se ve vzduchu je popsána pohybovými rovnicemi se šesti stup\v{n}i volnosti. Soustava translačních, rotačních a kinematických rovnic tvoří soustavu devíti nelineárních diferenciálních rovnic prvního řádu. Tyto rovnice lze linearizovat okolo nějakého rovnovážného stavu, který budeme nazývat letovým režimem. Součástí simulačního prostředí je řídící systém letadla založený na PID regulaci. Základem je návrh autopilota, který řídí úhel podélného sklonu a úhel příčného náklonu. Součástí návrhu jsou takzvané „flight director\textquotedblright \phantom{s}m\'dy jako udržení výšky, volba kursu, regulace vertikální rychlosti, změna výšky, zachycení požadované výšky a navigační m\'{o}d založený na nelineárním navigačním zákonu. Optimalizace regulátorů za použití PSO algoritmu a Pareto optimalitě je využita pro nastavení parametrů PID regulátoru. Simulační prostředí je vizualizováno v softwaru FlightGear.
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Lavergne, Fabien. "Méthodologie de synthèse de lois de commandes non-linéaires et robustes : application au suivi de trajectoire des avions de transport." Toulouse 3, 2005. http://www.theses.fr/2005TOU30248.
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Le travail présenté dans ce mémoire de thèse s'inscrit dans le cadre de la commande non-linéaire et robuste des avions de transport. Le but de cette thèse est de coupler les propriétés de la commande non-linéaire (adaptation aux non-linéarités de l'avion, synthèse de correcteurs explicites, facilité de réglage une fois la synthèse réalisée, généricité des lois de commande obtenues) à des propriétés de robustesse indispensables à l'activité aéronautique. En effet, pour garantir la sécurité des vols tant en pilotage manuel qu'en pilotage automatique, les lois de commande doivent présenter des propriétés fortes de stabilité et de performances robustes. Après une introduction au contexte industriel et de recherche du sujet de thèse, une partie "techniques, méthodes et outils" nous permet de mettre en avant les contributions du travail de thèse dans les domaines de la commande non-linéaire robuste et de la modélisation automatique. La technique de commande non-linéaire robuste présentée, appelée commande RMI (Robust Multi-Inversion) s'appuie sur la technique désormais classique d'inversion de la dynamique, notamment étudiée à Airbus depuis quelques années (Fabrice VILLAUME, Jean DUPREZ) et qu'elle robustifie par l'adjonction d'une boucle supplémentaire d'observation. Nous présentons aussi un outil de génération automatique de modèles non-linéaires, multivariables et embarquables, ainsi que les méthodes afférentes basées sur les réseaux de neurones. Cet outil est nécessaire à l'industrialisation des lois de commandes non-linéaires basées modèles. La partie applicative de la thèse souligne ensuite les particularités du système "avion" et propose des architectures de lois de commande, des trajectoires de référence associées, et la validation avancée de l'ensemble par simulations sur simulateur certifié. Enfin, après une conclusion sur le bilan de la thèse et les perspectives envisageables, nous proposons des annexes permettant d'approfondir certains aspects de notre étude<br>The work presented in this PhD thesis report is situated within the framework of the nonlinear and robust control of transport aircrafts. The purpose of this thesis is to couple the properties of nonlinear controllers (adaptation to the aircraft nonlinearities, explicit controllers synthesis, easy and decoupled setting once the synthesis is achieved, genericity of the obtained control laws) with essential robustness properties. Indeed, to guarantee the flight safety, both in manual handling and in automatic control, the control laws have to present strong robust stability and performances properties. After an introduction to the industrial and research context, a "techniques, methods and tools" part allows us to point out the thesis contributions in the nonlinear robust control and automatic modelling domains. The nonlinear robust control technique presented, called RMI control (for "Robust Multi-Inversion") is based on the now classical Nonlinear Dynamic Inversion (NDI) technique, notably studied at Airbus for some years (Fabrice VILLAUME, Jean DUPREZ), and is robustified by adding a complementary observation loop. We also present an automatic tool creating nonlinear, multivariable and embeddable models, as well as neural networks correlated methods. This tool is mandatory for the industrialization of our model-based flight control laws. Then the applicative part of the thesis underlines the specificities of the "aircraft" system and proposes flight control laws architectures, associated reference trajectories, and the advanced validation of the whole system by simulations performed on Airbus' certified simulator. Finally, after a conclusion on the main results and perspectives linked to the thesis, we propose annexes allowing to go further into the details of certain parts of our study
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Roussel, Emmanuel. "Contribution à la modélisation, l'identification et la commande d'un hélicoptère miniature." Thesis, Strasbourg, 2017. http://www.theses.fr/2017STRAD030/document.
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La stabilisation et l’automatisation du vol de tout véhicule aérien nécessite la mise en oeuvre d’algorithmes de commande. La synthèse et la simulation des lois de commande reposent sur un modèle mathématique du véhicule, qui doit être de complexité et de précision appropriées. Cette thèse présente une méthodologie complète d’identification appliquée à un hélicoptère coaxialminiature. L’étude théorique de son comportement en vol permet d’établir plusieurs modèles basés sur la mécanique du vol, qui diffèrent par les phénomènes aérodynamiques pris en compte. Ils sont identifiés, comparés et validés grâce à des données de vol, mettant en évidence l’importance de certains phénomènes dans la précision du modèle. Différentes lois de commande sont alors étudiées et évaluées en simulation puis par des expérimentations sur un prototype. Les résultats obtenus sont conformes aux simulations numériques, validant ainsi l’ensemble de la démarche<br>Control algorithms are at the heart of the stability and automatic flight capabilities of any aerial vehicle. Synthesis and simulation of control laws are based on a mathematicalmodel of the vehicle, which must be a trade-off between simplicity and accuracy. This work presents a complete system identification methodology applied on a miniature coaxial helicopter. Based on flight mechanics and aerodynamics, several models are built. They differ in the aerodynamic phenomena taken into account. They are identified, compared and validated thanks to flight data, highlighting important phenomena in the accuracy of the model. Several flight control strategies are then studied and evaluated through simulations and experiments with a prototype. The results are in accordance with numerical simulations, thus validating the whole approach
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Powly, A. A. "Variable Structure Control Based Flight Control Systems For Aircraft And Missiles." Thesis, 2004. http://etd.iisc.ernet.in/handle/2005/1136.
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Lei, Keng-Lon, and 李勁輪. "Application of Robust Control Theory to Missile Autopilot Design with High Off-Boresight Angle." Thesis, 1999. http://ndltd.ncl.edu.tw/handle/81681921771563892277.
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碩士<br>國立成功大學<br>航空太空工程學系<br>87<br>An autopilot system for the heading reversal maneuvering (HRM) phase in missile engagement with high off-boresight angle up to 180 degrees has been successfully synthesized with the combination of aerodynamic control and thrust vector control (TVC). The application of H∞ control theory has provided the missile robustness (including robust stability and robust performance) properties in the presence of model uncertainty caused by aerodynamic coefficient variation. Simultaneously, autopilot has enabled the missile to resist the maximal possible yawing moment with any switching frequency induced by the vortex shedding during high angle-of-attack flight. Performance of this autopilot system was validated with a high-fidelity nonlinear six degree-of-freedom simulation. Excellent results have been brought out and the results manifest the feasibility of robust control for various missile control problems.
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Chen, Wei-Yu, and 陳韋佑. "Integration and Simulations for Mission and Flight Control Systems of a Missile." Thesis, 2002. http://ndltd.ncl.edu.tw/handle/hv2j3f.
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碩士<br>國立成功大學<br>航空太空工程學系碩博士班<br>90<br>The objective of this thesis is to investigate the method of integrating the mission and control systems of a missile.
To validate the method, simulations with a 6-d.o.f. rigid body model are conducted.
In order to obtain a set of aerodynamic coefficient data, a typical configuration of twin-inlet missile is provided
and the data compendium method is used accordingly.
With the estimated aerodynamic data, a control system is designed to stabilize the missile and track the output of
the guidance system.
In this study, the mission is designed to be the maximum range of flight with mass of fuel fixed.
To obtain the optimal trajectory, the singular perturbation is used.
Since the control obtained from the optimal trajectory analysis is not an adequate input for the 6-d.o.f. rigid
body system, tracking the optimal trajectory output of the mission system is chosen to build up a robust integration
of the mission and control systems.
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Yu, Tsungte, and 游宗德. "The Study of Appling Support Vector Machines to the Missile Flight Control." Thesis, 2012. http://ndltd.ncl.edu.tw/handle/02253581127040028986.
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碩士<br>國防大學理工學院<br>兵器系統工程碩士班<br>100<br>This thesis applied the support vector regression (SVR) technique to establish an intelligent controller with online training and adaptive controls. This controller can determine the extremely short-term changes of control commands to predict the output of the next time step precisely.
This research employed LABVIEW (laboratory virtual instrument engineering workbench), an image programming language, and MATLAB, a support vector machine toolbox. The application of a DC motor speed controller provided good control results. Then we integrated the SVR technique into longitudinal autopilot that adopted a dual-feedback flight control system structure of acceleration and angular velocity. The simulation results show that the SVR controller (SVRC) can control missile longitudinal acceleration commands precisely and achieve adaptive control.
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Wang, Chun-Bi, and 王俊筆. "Base on Fuzzy PID to Design and Simulate the UAV Longitudinal Autopilot Flight Control System." Thesis, 2009. http://ndltd.ncl.edu.tw/handle/76730328713319491116.
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Lin, Hung-Wei, and 林煥巍. "A Study of Parameters Selection for an Integrated Guidance and Flight Control Autopilot Used in UAV." Thesis, 2004. http://ndltd.ncl.edu.tw/handle/65122643883255035998.
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碩士<br>國立成功大學<br>航空太空工程學系碩博士班<br>92<br>A F-16 as the UAV dynamic model and its integrated guidance and flight control autopilot is chosen in this research as the main focus of this study. The objective is to investigate the effect of autopilot parameters to the flight performance of an unmanned aerial vehicle (UAV) and proposes an analysis method for selecting a set of parameters best suited for many different flight conditions. The guidance law of the autopilot transforms the relative position between the UAV and a desired location and the velocity of the UAV into flight control command. The flight control law, a robust nonlinear controller, based on feedback linearization and sliding mode control receives the flight control command, directs the UAV, and reaches the desired location. The autopilot includes a large set of parameters worthy of in-dept study to achieve stable and optimal flight performance. In order to better understand the UAV flight behavior, a hardware-in-the-loop real-time simulation platform is formed with two PCs and avionic instrument. The UAV flight simulation software written with Visual C++ is placed in one PC, which connects with a position GPS and a high precision attitude compass through a data modulating single chip processor. The simulation data are transported to the other flight animation PC, in which an animation software written with graphical language, LabVIEW is developed. The platform helps engineers in designing or verifying the UAV autopilot. Finally, four flight control examples are included to demonstrate the importance of parameters selection to the integrated guidance and control autopilot.
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Chen, Feng-Wei, and 陳峰緯. "Based on Input Estimator and LQG Control Theory to Design Flight Satbility Controller for Supersonic Cruise Missile." Thesis, 2012. http://ndltd.ncl.edu.tw/handle/37259789151044730553.
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碩士<br>國防大學理工學院<br>兵器系統工程碩士班<br>101<br>In this thesis, the recursive input estimation method and LQG (Linear-Quadratic-Gaussian) control theory are utilized to design high-precision and robustness flight stability controller for supersonic cruise missile. The method for obtaining the disturbance input estimation ability and enhancing flight stability control performance are also proposed. The recursive input estimation with characteristics about faster computation, less measurements and low pass filter. LQG controller is based on optimal control theory has the characteristics of easy to implement. Combination of these two methods, the flight stability robust controller architecture that enables high-speed flight vehicle to maintain a stable flight performance in conditions of serious environmental disturbance are proposed. The thesis explore the flight vehicle two control issues
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Kumar, Shashi Ranjan. "Sliding Mode Control Based Guidance Strategies with Terminal Constraints." Thesis, 2015. http://etd.iisc.ernet.in/2005/3876.
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In the guidance literature, minimizing miss distance along with optimizing the energy usage had been an objective for several decades. In current day applications, additional terminal performance such as impact angle and impact time are of paramount importance. These terminal constraints increase warhead effectiveness and survivability of the interceptor. This thesis contributes to the design of guidance laws addressing terminal constraints such as impact angle, impact time, and both impact time as well as impact angle, in addition to interception of targets.
In the first part of the thesis, the guidance laws which ensure the alignment of the interceptor at a desired impact angle within a finite time is proposed using different variants of sliding mode control(SMC).The impact angle is first redefined in terms of line-of-sight angle and then the impact angle problem is converted to a simpler problem of controlling line-of-sight angle and their rates. The sliding mode capturability and interpretation of the guidance laws are presented. In order to cater to very large heading angle errors, which give rise to negative closing speed initially, modifications to the guidance laws are also suggested. The modifications to the guidance laws for avoiding singularities, which may be encountered during implementation, due to the inherent nature of terminal SMC, are suggested. However, the guidance laws, which alleviates the possibility of such singularities completely, are also designed by using non singular terminal SMC. The two loop guidance and control, for a skid-to-turn cruciform interceptor in the pitch plane, is also proposed with an autopilot designed using the concept of dynamic SMC. The guidance laws addressing impact angle constraint for three dimensional scenarios are also presented. Unlike the usual approach of decoupling the three dimensional engagement in to two mutually orthogonal planar engagements, the guidance laws are derived using coupled engagement dynamics. These guidance laws are designed using conventional and non singular terminal SMC and provide asymptotic and finite time alignment of the intercept or to the desired impact angles, respectively.
Next, the SMC based guidance laws which ensure the interception of targets at pre-specified impact times is proposed in this thesis. The guidance law is first designed for stationary targets and then extended to constant velocity targets using the notion of predicted interception point. A switching surface is designed using the concepts of collision course and time-to-go with non-linear engagement dynamics and its role in achieving the objectives is also discussed. In order to account for large heading angle errors and even for negative initial closing speeds, different methods of estimation of time-to-go, resulting in two different guidance laws, are used. Unlike the existing guidance laws, the proposed guidance laws achieve an impact time even less than its initially estimated value. The flexibility in selecting a desired impact time is also exploited using the maximum available acceleration information. A cooperative salvo attack strategy, based on the proposed impact time guidance law, with a desired impact time chosen in real time using a centralized coordination algorithm, is proposed for stationary targets. The coordination manager determines a common impact time based on time-to-goof the interceptors, by minimizing the total switching surface deviations which in turn reduces the control effort.
The thesis also proposes a SMC based guidance strategy which addresses impact angle and impact time constraints simultaneously. This guidance scheme is based on switching between impact time and impact angle guidance laws based on certain conditions. Unlike existing impact time guidance laws, the proposed guidance strategy takes into account the curvature of the trajectory due to the impact angle requirement. The interceptor first corrects its course to nullify the impact time error and then aims to achieve interception with desired impact angle. In order to reduce the transitions between the two guidance laws, a novel hysteresis loop is introduced in the switching conditions. Initially stationary targets are considered, and later the same guidance scheme is extended to constant velocity targets using the notion of predicted interception point.
Theclaimsofalltheguidancelawsarevalidatedwithextensivesimulationsandtheir performances are compared with existing guidance laws. Although all the guidance laws derived in the thesis are based on the assumption of constant speed interceptors, their performances are evaluated with a time-varying speed interceptor model, subjected to aerodynamic conditions, to validate their efficacy. The implementation of impact time guidance on time-varying speed interceptors is a formidable challenge in the guidance literature. Such implementations have also been presented in the thesis after introducing the notion of average speed and shown to yield satisfactory performance.
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Saroj, Kumar G. "An Integrated Estimation-Guidance Approach for Seeker-less Interceptors." Thesis, 2015. http://etd.iisc.ernet.in/2005/3828.
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In this thesis, the problem of intercepting highly manoeuvrable threats using seeker-less interceptors that operate in the command guidance mode, is addressed. These systems are more prone to estimation errors than standard seeker-based systems. Several non-linear and optimal estimation and guidance concepts are presented in this thesis for interception of randomly maneuvering targets by seeker-less interceptors. The key contributions of the thesis can be broadly categorized into six groups, namely (i) an optimal selection of bank of lters in interactive multiple model (IMM) scheme to cater to various maneuvers that are expected during the end-game, (ii) an innovative algorithm to reduce chattering phenomenon and formulate effective guidance algorithm based on 'differential game guidance law' (modi ed DGL), (iii) IMM/DGL and IMM/modified DGL based integrated estimation/guidance (IEG) strategy, (iv) sensitivity and robustness analysis of Kalman lters and ne tuning of lters in filter bank using innovation covariance,
(v) Performance of tuned IMM/PN, tuned IMM/DGL and tuned IMM/modi ed DGL against various target maneuvers, (vi) Performance comparison with realistic missile model.
An innovative generalized state estimation formulation has been proposed in this the-sis for accurately estimating the states of incoming high speed randomly maneuvering targets. The IMM scheme and an optimal selection of lters, to cater to various maneu-vers that are expected during the end-game, is described in detail. The key advantage of this formulation is that it is generic and can capture evasive target maneuver as well as straight moving targets in a uni ed framework without any change of target model and tuning parameters.
In this thesis, a game optimal guidance law is described in detail for 2D and 3D engagements. The performance of the differential game based guidance law (DGL) is compared with conventional Proportional Navigation (PN) guidance law, especially for 3D interception scenarios. An innovative chatter removal algorithm is introduced by modifying the differential game based guidance law (modified DGL). In this algorithm, chattering is reduced to the maximum extent possible by introducing a boundary layer around the switching surface and using a continuous control within the boundary layer. The thesis presents performance of the modified DGL algorithm against PN and DGL, through a comparison of miss distances and achieved accelerations. Simulation results are also presented for varying fiight path angle errors.
Apart from the guidance logic, two novel ideas have been presented following the evolving "integrated estimation and guidance" philosophy. In the rst approach, an in-tegrated estimation/guidance (IEG) algorithm that integrates IMM estimator with DGL law (IMM/DGL), is proposed for seeker-less interception. In this interception scenario, the target performs an evasive bang-bang maneuver, while the sensor has noisy measure-ments and the interceptor is subject to an acceleration bound. The guidance parameters (i.e., the lateral acceleration commands) are computed with the help of zero e ort miss distance. The thesis presents the performance of the IEG algorithm against combined IMM with PN (IMM/PN), through a comparison of miss distances. In the second ap-proach, a novel modi ed IEG algorithm composed of IMM estimator and modi ed DGL guidance law is introduced to eliminate the chattering phenomenon. Results from both of these integrated approaches are quite promising. Monte Carlo simulation results re-veal that modi ed IEG algorithm achieves better homing performance, even if the target maneuver model is unknown to the estimator. These results and their analysis o er an insight to the interception process and the proposed algorithms.
The selection of lter tuning parameters puts forward a major challenge for scien-tists and engineers. Two recently developed metrics, based on innovation covariance, are incorporated for determining the filter tuning parameters. For predicting the proper combination of the lter tuning parameters, the metrics are evaluated for a 3D interception problem. A detailed sensitivity and robustness analysis is carried out for each type of Kalman lters. Optimal and tuned Kalman lters are selected in the IMM con guration to cater to various maneuvers that are expected during the end-game.
In the interception scenario examined in this thesis, the target performs various types of maneuvers, while the sensor has noisy measurements and the interceptor is subject to acceleration bound. The tuned IMM serves as a basis for synthesis of e cient lters for tracking maneuvering targets and reducing estimation errors. A numerical study is provided which demonstrates the performance and viability of tuned IMM/modi ed DGL based modi ed IEG strategy. In this thesis, comparison is also performed between tuned IMM/PN, tuned IMM/DGL and tuned IMM/modi ed DGL in integrated estimation/guidance scheme. The results are illustrated by an extensive Monte Carlo simulation study in the presence of estimation errors. Simulation results are also presented for end game maneuvers and varying light path angle errors .
Numerical simulations to study the aerodynamic e ects on integrated estimation/ guidance structure and its e ect on performance of guidance laws are presented. A detailed comparison is also performed between tuned IMM/PN, tuned IMM/DGL and tuned IMM/modi ed DGL in integrated estimation/guidance scheme with realistically modelled missile against various target maneuvers. Though the time taken to intercept is higher when a realistic model is considered, the integrated estimation/guidance law still performs better. The miss distance is observed to be similar to the one obtained by considering simpli ed kinematic models.