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NASA supercritical wing technology, and the digital fly-by-wire program are discussed as used on the F-8 research aircraft. Furthermore, the Highly Maneuverable Aircraft Technology (HiMAT) program, using a 0.44-scale, 3500 lb jet-powered remotely piloted research aircraft, is analyzed. Highly accurate data on aerodynamic and structural loads and deflection data were obtained from model flight tests. In addition, an oblique-wing technology is discussed through the example of the piloted AD-1 research aircraft, noting the low cost of the experiment. Other topics include: integrated system technology; generic forces of the research aircraft program; and aeronautical technology application trends.

The formulation and flight test results of an algorithm to detect and isolate the first failure of any one of 12 duplex control sensor signals (24 in all) being monitored are reviewed. The technique uses like signal differences for fault detection while relying upon analytic redundancy relationships among unlike quantities to isolate the faulty sensor. The fault isolation logic utilizes the modified sequential probability ratio test, which explicitly accommodates the inevitable, irreducible low frequency errors present in the analytic redundancy residuals. In addition, the algorithm used sensor output selftest, which takes advantage of the duplex sensor structure by immediately removing a hard failed sensor from control calculations and analytic redundancy relationships while awaiting a definitive fault isolation decision via analytic redundancy. This study represents a proof of concept demonstration of a methodology that is applied to duplex or higher flight control sensor configurations and, in addition, can monitor the health of one simplex signal per analytic redundancy relationship.

A triplex digital fly by wire flight control system was developed and installed in an F-8C aircraft to provide fail operative, full authority control. Hardware and software redundancy management techniques were designed to detect and identify failures in the system. Control functions typical of those projected for future actively controlled vehicles were implemented.

A triplex digital fly-by-wire flight control system was developed and then installed in a NASA F-8C aircraft to provide fail-operative, full authority control. Hardware and software redundancy management techniques were designed to detect and identify failures in the system. Control functions typical of those projected for future actively controlled vehicles were implemented. This paper describes the principal design features of the system, the implementation of computer, sensor, and actuator redundancy management, and the ground test results. An automated test program to verify sensor redundancy management software is also described.

Flight test results of the F-8 digital fly-by-wire (DFBW) control system are presented and the implications for application to active control technology (ACT) are discussed. The F-8 DFBW system has several of the attributes of proposed ACT systems, so the flight test experience is helpful in assessing the capabilities of those systems. Topics of discussion include the predicted and actual flight performance of the control system, assessments of aircraft flying qualities and other piloting factors, software management and control, and operational experience.

The emergence of highly reliable fly-by-wire flight control systems makes it possible to consider a strong reliance on automatic control systems in the design optimization of future aircraft. This design philosophy has been referred to as the control configured vehicle approach or the application of active control technology. Several studies and flight tests sponsored by the Air Force and NASA have demonstrated the potential benefits of control configured vehicles and active control technology. The present status and trends of active control technology are reviewed and the impact it will have on aircraft designs, design techniques, and the designer is predicted.

A digital fly-by-wire flight control system was designed, built, and for the first time flown in an airplane. The system, which uses components from the Apollo guidance system, is installed in an F-8 airplane as the primary control system. A lunar module guidance computer is the central element in the three-axis, single-channel, multimode, digital control system. A triplex electrical analog system which provides unaugmented control of the airplane is the only backup to the digital system. Flight results showed highly successful system operation, although the trim update rate was inadequate for precise trim changes, causing minor concern. The use of a digital system to implement conventional control laws proved to be practical for flight. Logic functions coded as an integral part of the control laws were found to be advantageous. Although software verification required extensive effort, confidence in the software was achieved.

An in-flight investigation of the effect of pure time delays on low L/D space shuttle type landing tasks was undertaken. The results indicate that the sensitivity of the pilot ratings to changes in pure time delay in pitch is strongly affected by the task and only slightly affected by changes in control system augmentation mode. Low L/D spot landings from a lateral offset were twice as sensitive to pure time delay as normal low L/D landings. For comparison purposes, formation flying was also investigated, and was found to be less sensitive to time delay than the landing tasks.

The experience gained in digital fly-by-wire technology through a flight test program being conducted by the NASA Dryden Flight Research Center in an F-8C aircraft is described. The system requirements are outlined, along with the requirements for flight qualification. The system is described, including the hardware components, the aircraft installation, and the system operation. The flight qualification experience is emphasized. The qualification process included the theoretical validation of the basic design, laboratory testing of the hardware and software elements, systems level testing, and flight testing. The most productive testing was performed on an iron bird aircraft, which used the actual electronic and hydraulic hardware and a simulation of the F-8 characteristics to provide the flight environment. The iron bird was used for sensor and system redundancy management testing, failure modes and effects testing, and stress testing in many cases with the pilot in the loop. The flight test program confirmed the quality of the validation process by achieving 50 flights without a known undetected failure and with no false alarms.

The NASA Dryden Flight Research Center is flight testing a triply redundant digital fly-by-wire (DFBW) control system installed in an F-8 aircraft. The full-time, full-authority system performs three-axis flight control computations, including stability and command augmentation, autopilot functions, failure detection and isolation, and self-test functions. Advanced control law experiments include an active flap mode for ride smoothing and maneuver drag reduction. This paper discusses research being conducted on computer synchronization, fault detection, fault isolation, and recovery from transient faults. The F-8 DFBW system has demonstrated immunity from nuisance fault declarations while quickly identifying truly faulty components.

Flight test results of the F-8 digital fly by wire control system are presented and the implications for application to active control technology are discussed. The F-8 DFBW system has several of the attributes of proposed ACT systems, so the flight test experience is helpful in assessing the capabilities of those systems. Topics of discussion include the predicted and actual flight performance of the control system, assessments of aircraft flying qualities and other piloting factors, software management and control, and operational experience.

Motion cue and simulation fidelity aspects of validation of general purpose airborne simulator

Analysis and validation of characteristics of general purpose airborne simulator for simulation of aerodynamic characteristics of large transport aircraft

Airborne simulator program for evaluation of motion and visual cue effects on pilot performance in roll, using compensatory tracking tasks

Response feedback control loops effects on lateral-directional dynamics of variable stability transport aircraft

This paper discusses the design and initial flight tests of the first digital fly-by-wire system to be flown in an aircraft. The system, which used components from the Apollo guidance system, was installed in an F-8 aircraft. A lunar module guidance computer is the central element in the three-axis, single-channel, multimode, digital, primary control system. An electrohydraulic triplex system providing unaugmented control of the F-8 aircraft is the only backup to the digital system. Emphasis is placed on the digital system in its role as a control augmentor, a logic processor, and a failure detector. A sampled-data design synthesis example is included to demonstrate the role of various analytical and simulation methods. The use of a digital system to implement conventional control laws was shown to be practical for flight. Logic functions coded as an integral part of the control laws were found to be advantageous. Verification of software required an extensive effort, but confidence in the software was achieved. Initial flight results showed highly successful system operation, although quantization of pilot's stick and trim were areas of minor concern from the piloting standpoint.