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The precision approach path indicator (PAPI) system is installed alongside a runway or at a heliport to provide accurate visual approach path information to pilots of landing aircraft, either day or night. The indicator sends the information through different color combinations. The international standards have very rigid requirements for the intensity distribution and the transition angle from red light to white light of the indicator. These two parameters are the main parameters to judge the accuracy of the indicator. We propose a new method to measure these two optical parameters, which can improve the accuracy of the signals and guarantee safe landing of the aircraft. OCIS codes: 120.0280, 100.2000.

For a research aircraft, 'conventional' control laws (CLs) are implemented on a 'baseline' flight computer (FC) while research CLs are typically housed on a dedicated research computer. Therefore, for an experimental aircraft used to test specific fault tolerant flight control systems, a safety logic scheme is needed to ensure a safe transition from conventional to research CLs (while at nominal conditions) as well as from research CLs at nominal conditions to conditions with 'simulated' failures on specific control surfaces. This paper describes the design of such a safety scheme for the NASA Intelligent Flight Control System (IFCS) F-15 Program. The goals of the IFCS F-15 program are to investigate the performance of a set of fault tolerant CLs based on the use of dynamic inversion with neural augmentation. The different transitions are monitored using information relative to flight conditions and controller-related performance criteria. The testing of the scheme is performed with a Simulink-based flight simulation code and interface developed at West Virginia University for the NASA IFCS F-15 aircraft.

The performance of a phase-detection frequency-domain reflectometer (PD-FDR) for locating open and short circuits (hard faults) in a Navy F-18 flight control harness has been tested, and the analytical expressions for accuracy verified. Nine different types of aircraft wires appear in this harness: twisted pair, shielded wires with 1-4 inner conductors, 'filter wire,' and bundles of individual wires. PD-FDRs in a variety of frequency ranges (12-25, 100-220, 150-300, and 180-400 MHz) are compared. Signal processing techniques are utilized to remove the reflections where the PD-FDR is connected to the wire harness, which is critical to obtaining accurate measurements, particularly for short lengths of wire. For this specific application, open and short circuits are located to within 2.5 cm (1 in) for PD-FDR200 and 11 cm (5.5 in) for PD-FDR25 for wires ranging from 9 cm to 9.15 m (6-360 in). Index Terms--Aging wire, frequency domain reflectometer (FDR), wire fault location, reflectometry.

We investigated the in-orbit performance of a high-performance on-board computer developed with commercial off-the-shelf (COTS) technology in terms of its performance during the occurrence of single event effects. The processor worked and performed successfully both under normal and under solar flare conditions in 800 km altitude polar orbit. During a solar flare, the occurrence of single events increased by a factor of more than four compared with normal conditions. The area where single events occurred during the solar flare spread to the polar region, whereas normally they are limited to the region of South-Atlantic Anomalies (SAA). Our results suggest that the performance of our COTS processor is sufficient for future space applications.

A brief description on the making and features of an autopilot Optic flow based Control sysTem for Aerial Vehicles (OCTAVE) is presented. The results enable autonomous aerial robots to perform three further reputedly different tasks, autonomous take off, autonomous landing, and autonomous compensation for wind disturbances.

New concepts for microactivated flight control of projectiles have been investigated for operation in both subsonic and supersonic flight regimes. A successful concept can provide an inexpensive flight control device with many control redundancies, and it can reduce size weight and energy consumption of the control system. Two promising concepts have been explored in detail. The first one consists of a microflap in the form of a delta wing of 3 mm base by 4 mm height, which can be deployed and activated at a location on the ogive of the projectile or at the beginning of its cylindrical afterbody. The second concept is based on microballoons, which can be inflated and change the local pressure distribution on the projectile. A computational-fluid-dynamics (CFD) computer code has been used to model these two flight control concepts. The first concept was tested at two angles of extension, 9 and 21 deg and at seven different flight Mach numbers: 0.3, 0.5, 0.7, 0.8, 1.0, 1.5, 2.0, and 3.0. Some additional cases with 150-deg extension angle were also investigated. The second concept based on microballoons was tested at four different Mach numbers: 0.6, 1.0, 1.8 and 2.2. Particular emphasis has been given in the CFD analysis in resolving the flow characteristics around the deployed control surfaces. Longitudinal vortices were found above the upper surface of the delta wing in all investigated Mach numbers, which reduce the pressure there and shock waves acting on the lower surface in the case of supersonic flights. The results show that the ratio of lift to drag of the control surface is about nine in the subsonic flow regime and about five in the supersonic flow regime. These values suggest that this control surface can provide adequate control force for possible maneuvering of the projectile. However, microactuators deployment was also found to increase the total drag of the projectile by 1 to 4%. In cases with deployment of two microactuators located next to each other, mutual interference affects favorably the aerodynamic control forces.

NASA is working to prevent 'loss-of-control' (LOC) on the part of pilots, an effort that is included in the U.S. government's $500 mil Aviation Safety Program. There were more aircraft accidents due to LOC than any other cause between 1990-96 for both the commerical transport and general aviation sectors. NASA aims to lower the accident rate by 9% by 2022.

Research is presented concerning the importance of simulated batch and pitch in the development of flight control systems of an advanced nature. The development of simulations which yield accurate results is discussed.

Issues concerning the use of wind-tunnels to test an aerodynamic flight control system are presented. The delta flap system is designed to aid recovery from spins and stalls by the generation of high lift of the wing.

Aircraft parameter identification techniques have become accepted as indispensable tools in the evaluation of prototype- and derivative aircraft in flight. Applications include estimation of stability- and control derivatives in the linearized equations of motion, synthesis of nonlinear aerodynamic and propulsion models in the context of performance certification and incorporation of information from dynamic flight test manoeuvres in a priori nonlinear flight simulation models. A variety of techniques in the time- as well as the frequency domain have been applied in the past. One of the successful techniques is the so-called two-step method (TSM) in which the original state-parameter estimation problem is decomposed into a nonlinear state/parameter estimation or reconstruction problem and a subsequent linear parameter identification problem. In the literature, the first step of the TSM is often referred to as 'flight path reconstruction'. The present paper focuses on the first step of the TSM. After a derivation of the system models describing the flight path relative to a flat earth as well as a spherical and rotating earth, and observation models for air data and GPS, the flight path reconstruction problem is introduced. Requirements with respect to type and quality of flight test transducers are discussed. Next follows an overview of different approaches to the solution of the flight path reconstruction problem with emphasis on Kalman filter/smoother and Maximum Likelihood methods. A new adaptive algorithm is presented, the Modified Recursive Maximum Likelihood Adaptive Filter (MRML) which is shown to be significantly more robust with respect to initialisation errors than earlier methods. A reconstructibility analysis is presented for different transducer combinations. Numerical examples are presented based on simulated as well as actual flight test data. Flight results are given of the flight path reconstruction part of an on-line pseudo real-time application of the TSM. The paper ends with concluding remarks.

This paper develops a method to tune fuzzy controllers using numerical optimization. The main attribute of this approach is that it allows fuzzy logic controllers to be tuned to achieve global performance requirements. Furthermore, this approach allows design constraints to be implemented during the tuning process. The method tunes the controller by parameterizing the membership functions for error, change-in-error, and control output. The resulting parameters form a design vector which is iteratively changed to minimize an objective function. The minimal objective function results in an optimal performance of the system. A spacecraft mounted science payload line-of-sight pointing control is used to demonstrate results. Index Terms - Attitude control, fuzzy logic controllers, numerical optimization, pointing control, spacecraft control.

As a part of trajectory modulation to increase terminal effectiveness, impact angle control is required in the terminal phase of tactical missile systems. Moreover, the missile systems are allowed to have only low altitude and minimum flight time to reduce the probability of detection by sensors of missile defense systems and thus to increase survivability. In this paper, a simple numerical solution for a time-optimal control law is suggested in the case of constrained missile maneuverability and impact angle for a first-order-lag autopilot. The numerical solution utilizes an analytic solution of a simplified problem with a zero-lag autopilot assumption to determine optimal switching instants. Performance of the proposed control law is tested by a simulation study.

The aircraft flight quantities and success of the mission depend to a great extent upon the actuator performance, and flight actuators must be designed to achieve the specified criteria. Electromechanical flight actuators driven by electric motors have begun to displace hydraulic technology in advanced flight vehicles. In aerospace application, permanent-magnet stepper motors are perfectly suited due to their efficiency and reliability, low volume-, weight-, and size-to-torque ratios, high power and torque densities, low cost and maintenance, simplicity and ruggedness, etc. Conventional open-loop stepper motor servos do not ensure the required accuracy and dynamic performance. An innovative method in motion control of advanced electromechanical flight actuators is developed, and nonlinear controllers are designed. The specified tracking accuracy, desired stability margins, microstepping capabilities, and disturbance attenuation are ensured by the robust nonlinear controllers synthesized. Analytical, numerical, and experimental results are documented to study the performance of flight actuators directly driven by stepper motors and to demonstrate the efficiency of control algorithms.

This paper addresses the motivations for using a distributed control system architecture, technical challenges, typical functions which are off-loaded to remote terminals, sensor/effector interface issues, data bus selection, technology insertion issues, lessons learned, and objectives for future distributed control implementations. Typical design requirements, constraints, environmental conditions, and operational challenges will be described. Examples of various distributed control system implementations will be discussed, including both propulsion control and flight control examples.

A method of designing control laws for uncertain nonlinear systems is presented. Dynamic inversion is used to partially linearize the dynamics and then a nonlinear version of quantitative feedback theory (QFT) is applied to the resulting system which assures robustness to plant uncertainty. The design yields good performance with low bandwidth. An application to the design of flight control laws for a high performance aircraft is presented. The control laws demonstrate good performance by accurately following large angle of attack commands at flight speeds ranging from 53 to 150 m/s. Robustness is verified by including [+ or -]20 percent variations in pitching moment derivatives. The reduced bandwidth compared to a fixed-gain, linear design, leads to greatly reduced actuator transients, which should give improved reliability and longer life for the actuators and associated structure.

A regulator design technique is presented for linear parameter varying (LPV) systems. This technique may be applied to many different types of systems, including nonlinear, due to the broad class of systems that may be represented by LPVs. The regulator, consisting of an inner loop and an outer loop, renders the closed-loop system's steady-state input-output to be linear time invariant (LTI) and causes the output to track a commanded trajectory. With real-time, accurate parameter data, the inner loop effectively cancels the parameter dependent terms. The outer loop is designed using LTI [H.sub.[infinity]] synthesis to enable the closed loop system to meet stability and performance goals. Due to the inner loop controller and imperfect parameter cancellation, the complete closed-loop system is likely to be a nonlinear function of the parameters and their derivatives. To assess the stability using the quadratic Lyapunov test, we model the closed-loop system as a polytopic system. The key ideas are illustrated with a nonlinear aircraft flight control example.

A research was conducted to examine tangential forebody blowing on a 6% generic fighter model in the 2.1 x 1.5-m low-speed wind tunnel. Three distinct forebodies were also used to study the effects of differential azimuthal slot positions. Results show that slot locations demonstrate significant yawing moments even at angles of attack above 60 degrees, thereby, implying that control moment did not weakened at the maximum angle of attack tested.

A relative navigation system based on both the Inertial Navigation System (INS) and the Global Positioning System (GPS) is developed to support situational awareness during formation flight. The architecture of the system requires an INS/GPS integration across two aircraft via a data !ink. A fault-tolerant federated filter is used to estimate the relative INS errors based on relative GPS measurements and a range measurement obtained from the data link. The filter is constructed based on a reduced-order model of the relative INS error process. A method for analyzing the filter performance is presented. A case involving two helicopters in formation flight is studied under three different flight trajectories to account for the effect of vehicle motion on the INS state transition matrix. The results of the covariance analysis are compared with actual flight results over an instrumented test range.

A Kalman filter based multiple sensor fusion method for determining the position of a general aviation aircraft with respect to the runway during night landing and takeoff is discussed. Known structure of the airport lights, video images acquired by an onboard video camera, position estimates from an onboard Global Positioning System (GPS), and data from an attitude indicator are integrated in a Kalman filtering algorithm. Simulation results are presented to demonstrate the feasibility of the proposed concept.

A system identification methodology can be efficiently applied to flight vehicle modeling problems. An accurate estimation of aircraft parameters is possible due to the developments in measurement techniques and data processing facilitated by digital computers. Examples of updating and development of databases from flight test data, and interdisciplinary modeling characteristics of rotor dynamics and rigid body are given. The modern methods of parameter estimation and the application areas of system identification are discussed.

The feasibility of using pneumatic blowing to attain optimum directional and yaw control of aircrafts at high angles of attack was evaluated using a generic subscale model. Experiments were carried out in the University of Notre Dame Low Speed Wind Tunnels wherein the model was tested at various angle-of-attack values. Results show that the effectiveness of the proposed pneumatic control is strongly influenced by the tail configuration of the aircraft.

This paper presents a study of the effects of structural flexibility on one of the control loops that constitute the Attitude and Articulation Control Subsystem of the Galileo Spacecraft. Both stability analysis and time domain simulation studies are discussed. Results obtained indicate that the control loop of interest - the Clock Controller - will interact stably with the spacecraft structure as long as the spacecraft scan platform boresight points at least 30 deg away from the spacecraft spin axis (poles), Flexibility effects become more and more pronounced with proximity of the bore-sight to the poles as several structural modes become excitable to resonance.

This paper examines the performance enhancement of a statically unstable aircraft subject to the input and state constraints. Under control saturation, i/o linearizability is destroyed and the state trajectories may not be attracted to the sliding surface. If the reference signals are sufficiently large and the zero-dynamics is lightly damped, the i/o linearizing control may become unreasonably large in magnitude, making the closed-loop system susceptible to the damaging effects of control saturation. In addition to performance degradations such as increased tracking errors, control saturation can drive the closed-loop system to instability. In this paper, a new design method called approximate i/o linearization is presented to enhance the performance of the SISO longitudinal flight control problem under saturation. The new approximate i/o linearization law is obtained by solving a pointwise minimization problem. The function to be minimized consists of a surface whose relative degree is one, its derivative, and weighted square of the input u. The advantages of the approximate i/o linearization is that the adverse effects of control saturation can be minimized by properly selecting the weight on the usage of the control. The only requirement for the new technique is that the original plant be locally i/o linearizable. Thus approximate i/o linearization does not impose additional strict requirements on the plant. In the remaining sections of the paper, stability and bounded tracking properties of the approximate i/o linearization are proven. Finally, a longitudinal flight control problem is used to demonstrate the application of approximate i/o linearization.

In this paper a computational scheme is presented for accurately predicting the farfield amplitude and phase characteristics of Global Positioning System (GPS) antennas flush-mounted to a corrugated groundplane. The algorithm developed is particularly well-suited in beamshaping of (GPS) antennas in order to provide a high level of multipath rejection. The usefulness of the analytical model has been verified by the excellent agreement achieved between experimental data and predicted amplitude and phase patterns.

We analytically obtain the capture regions for the true proportional navigation (TPN) missile guidance law against an intelligently maneuvering target. Two versions of TPN are considered. The first is the original TPN which assumes the commanded lateral acceleration to be directly proportional to the line-of-sight (LOS) rate only, the proportionality factor being an arbitrary constant or dependent only on the initial closing velocity. The other, known as RTPN (realistic TPN), assumes the commanded lateral acceleration to be directly proportional to the LOS rate and also the current closing velocity. The target is assumed to maneuver in such a way as to increase the LOS rate and thus directly oppose the proportional navigation (PN) philosophy of annulling the LOS rate. A necessary and sufficient condition for capture is derived for the original TPN, and using it, the exact capture region is obtained. A sufficient condition for capture is derived for RTPN and is used to obtain its capture region partially. Some necessary conditions for capture are also derived for RTPN and are used to obtain an upper bound on its complete capture region. Using these conditions some important results on the existence of capture regions and a comparative study of capturability of TPN laws are also presented.

A mathematical model describing the dynamics of a self-contained flight control actuator is established. Typical failures of the actuator are defined in the space of the model parameters in the form of a failure pattern library. A diagnostic model of the actuator, intentionally sensitive to particular failures, is defined. A diagnostic procedure, featuring on-line parameter estimation of the diagnostic model, failure detection, and isolation, is proposed. A failure prediction procedure, utilizing trends exhibited by parameter estimates, is formulated. The results are verified by computer simulation.

An active noise control system using control loudspeakers in the area between two walls separating source and space where control is required, attenuates the noise greatly. The system also significantly attenuates sound across the curved double wall of an aircraft fuselage mock-up. The system includes error microphones and attenuation increases proportionally with the distance between microphone and speakers.

Fundamentals and novel similarity transformations that simulate poststall (PST) and thrust-vectoring flight controlled (TVFC) aircraft yield designs for scale model flight tests that maximize agility during PST maneuvers. The tests reveal that combining TFVC with conventional flight controls enables agility maximization.

Flight control systems can benefit by being designed to emulate functions of natural intelligence. Intelligent control functions fall in three categories. Declarative actions involve decision making, providing models for system monitoring, goal planning, and system/scenario identification. Procedural actions concern skilled behavior and have parallels in guidance, navigation, and adaptation. Reflexive actions are more or less spontaneous and are similar to inner-loop control and estimation. Intelligent flight control systems will contain a hierarchy of expert systems, procedural algorithms, and computational neural networks, each expanding on prior functions to improve mission capability, to increase the reliability and safety of flight, and to ease pilot workload.

An airdata sensing system is developed for the accurate analysis of aircraft aerodynamics at high angles-of-attack. The high angle-of-attack airdata sensing system system (HI-FADS) consists of a simple hardware and a set of 25 pressure orifices arranged in concentric circles on the aircraft's nosecap. It uses modified system matrix equations. To evaluate its reliability, the HI-FADS is installed on the high alpha research flight vehicle for flight testing. Results indicate that the HI-FADS perform with greater accuracy and wider angle-of-attack range than conventional intrusive airdata systems.

Forward flight effects on the counterflow thrust vector control (CFTVC) of a supersonic jet has been studied by generating a coflowing stream around the periphery of the CFTVC system made up of a rectangular Mach 1.4 jet. Results reveal that the jet can be easily and continuously vectored to as much as 22 degrees for all the three coflow Mach numbers. The system response with coflow is comparable to that without coflow with just a slight decrease in performance.

This Technical Brief addresses the problem of robust pole assignment for uncertain dynamic systems. A control design methodology is proposed to determine a fixed controller assigning all poles of uncertain systems in a pre-specified disk. The robust pole assignment problem is reduced to a convex optimization problem. The proposed method is applied in design of a flight control system for a small aircraft successfully.

The laminated plate theory and kinematics can be successfully used to compute the deflections produced by piezoceramic Flexspar flight control surface actuators. Experimental model testing indicated that pitch deflections up to +/-11 degrees may be obtained with Flexspar control surfaces with 4-in. span and 3.33-in. mean geometric chord. Dynamic testing showed an 11-Hz natural frequency and a 23-Hz break frequency with a 40 mW of power used at the most extreme actuation condition.

An aeroservoelastic interaction between the structural mode response, the buffet on the aircraft and the flight-control systems in the lateral-directional axis was evident during flight envelope expansion operation through 45-degree angle-of-attack in a forward-swept wing aircraft. The buffet-excited interaction, consisting of three modes, influences the control surface actuator commands.

High intensity electromagnetic radiation has been demonstrated to be a source of computer upsets in commercially available digital flight control systems. Such upsets can degrade the quality of the control signal ranging from a perturbation error over a few sample periods to a permanent error mode or computer failure. Under these conditions, the primary concern of the control engineer is to insure that the closed-loop system remains stable. A stochastic disturbance model and a set of associated stability assessment tools are introduced for determining stability robustness of a nominal closed-loop system subject to electromagnetic disturbances. The focus is primarily on flight control applications, but the methodology is suitable for any application where highly reliable digital control is needed. The technique is demonstrated on a simple test example and on a stabilizing controller for the longitudinal dynamics of the AFTI/F-16 aircraft.

The mean and covariance of a Kalman filter residual are computed for specific cases in which the Kalman filter model differs from a linear model that accurately represents the true system (the truth model). Multiple model adaptive estimation (MMAE) uses a bank of Kalman filters, each with a different internal model, and a hypothesis testing algorithm that uses the residuals from this bank of Kalman filters to estimate the true system model. At most, only one Kalman filter's model will exactly match the truth model and will produce a residual whose mean and standard deviation have already been analyzed [9]. All of the other filters use internal models that mismodel the true system. We compute the effects of a mismodeled input matrix, output matrix, and state transition matrix on these residuals. The computed mean and covariance are compared with simulation results of flight control failures that correspond to mismodeled input matrices and output matrices.

Athena Technologies uses the 'straight-up mathematics' control theory to reduce the size of the control systems of unmanned aerial vehicles. The company uses non-linear control design theory and modeling to improve the accuracy of the aircraft's control system and navigation.

The thrust vectoring flight control (TVFC) technology would make the air journey safe and prevent the damage to aircraft operation. TVFC is spin free and independent of the external flow regime upon which the conventional aerodynamic flight control and air safety regulations and training of pilots depend. Application of military TVFC to civil air transport would enhance combat capability and flight safety standards. Incorporation of new thrust reversal methods to TVFC systems would maximize the potential and reliability of aircraft.

Several programs on integrating artificial intelligence (IA) in aircraft cockpits have been established. Among them are the 'a reliable computer-human interaction environment' project and the cockpit assistance system program. However, several problems such as lack of standards and high installation costs have to be resolved before AI systems can be fitted in commercial aircraft.