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A digital avionics system was installed in the CV-340 transport aircraft. Flight tests were made to obtain preliminary performance data in the manual flight director mode using time controlled guidance. These tests provide a basis for selection of terminal area guidance, navigation, and control system concepts for short haul aircraft and for investigating operational procedures.
A laboratory and field test program was conducted to quantify the optical navigation parameters of the Mariner vidicons. A scene simulator and a camera were designed and built for vidicon tests under a wide variety of conditions. Laboratory tests characterized error sources important to the optical navigation process and field tests verified star sensitivity and characterized comet optical guidance parameters. The equipment, tests and data reduction techniques used are described. Key test results are listed. A substantial increase in the understanding of the use of selenium vidicons as detectors for spacecraft optical guidance was achieved, indicating a reduction in residual offset errors by a factor of two to four to the single pixel level.
Difficulties arise in guiding a solar electric propulsion spacecraft due to nongravitational accelerations caused by random fluctuations in the magnitude and direction of the thrust vector. These difficulties may be handled by using a low thrust guidance law based on the linear-quadratic-Gaussian problem of stochastic control theory with a minimum terminal miss performance criterion. Explicit constraints are imposed on the variances of the control parameters, and an algorithm based on the Hilbert space extension of a parameter optimization method is presented for calculation of gains in the guidance law. The terminal navigation of a 1980 flyby mission to the comet Encke is used as an example.
The sensitivity of the trajectory of an Aeromaneuvering Orbit-to-Orbit Shuttle (AMOOS) to unpredictable variations in atmospheric density is shown and a compensating guidance law is developed. The sensitivity was determined using a varying atmosphere model whose statistical properties matched those of actual atmosphere data. This study showed that a compensating guidance law is crucial to the feasibility of AMOOS. Using an optimal regulator approach, a linear feedback law for bank angle was developed and checked out with the atmosphere model. This law uses the altitude profile of the trajectory to control the amount of energy dissipated within the atmosphere.
An approach has been developed to determine the guidance and control system sensitivity to off-nominal aerodynamics for the Space Shuttle Orbiter during entry. This approach, which uses a nonlinear six-degree-of-freedom interactive, digital simulation, has been applied to both the longitudinal and lateral-directional axes for a portion of the orbiter entry. Boundary values for each of the aerodynamic parameters have been identified, the key parameters have been determined, and system modifications that will increase system tolerance to off-nominal aerodynamics have been recommended. The simulations were judged by specified criteria and the performance was evaluated by use of key dependent variables. The analysis is now being expanded to include the latest shuttle guidance and control systems throughout the entry speed range.
This paper, assuming steady wind and zero sideslip, presents a discrete-time mathematical model to obtain a control law and three-dimensional flight path to guide an aircraft in a given time from a given initial state (position, velocity and heading) to a prescribed final state subject to the constraints on airspeed acceleration, and pitch and bank angles of the aircraft. For ease in implementing the control law, the control inputs are assumed to be delayed and are applied in a sequential fashion. The guidance problem is formulated as a discrete nonlinear optimal control problem with time delays in dynamics and a cost functional of Bolza form. With a quadratic penalty function to handle terminal constraints on velocity and heading, a solution technique to the control problem based on conjugate gradient algorithm is investigated. Numerical examples are presented to illustrate the applicability of this approach to solution of a terminal area guidance problem in an automated air traffic control environment.
A nonlinear six degree of freedom entry simulation study was conducted to identify space shuttle guidance and control system software modifications which reduce the control system sensitivity to the guidance system sampling frequency. Several modifications which eliminated the control system sensitivity and associated control limit cycling were examined. The result of the modifications was a reduction in required reaction control system fuel.
The Saturn launch vehicle's guidance and control system is so complex that the reliability of a simplex system is not adequate to fulfill mission requirements. Thus, to achieve the desired reliability, redundancy encompassing a wide range of types and levels was employed. At one extreme, the lowest level, basic components (resistors, capacitors, relays, etc.) are employed in series, parallel, or quadruplex arrangements to insure continued system operation in the presence of possible failure conditions. At the other extreme, the highest level, complete subsystem duplication is provided so that a backup subsystem can be employed in case the primary system malfunctions. In between these two extremes, many other redundancy schemes and techniques are employed at various levels. Basic redundancy concepts are covered to gain insight into the advantages obtained with various techniques. Points and methods of application of these techniques are included. The theoretical gain in reliability resulting from redundancy is assessed and compared to a simplex system. Problems and limitations encountered in the practical application of redundancy are discussed as well as techniques verifying proper operation of the redundant channels. As background for the redundancy application discussion, a basic description of the guidance and control system is included.
One candidate for a reusable upper stage to be carried by the space shuttle is an aeromaneuvering orbit-to-orbit shuttle (AMOOS). This concept uses the drag of the vehicle during a pass through the atmosphere rather than the propulsion system to slow the vehicle on a return from a high energy orbit. The nature and magnitude of the sensitivity of AMOOS to uncertainties in the properties of the atmosphere are shown. Various guidance schemes for correcting for the effects that the unpredictable variations in the atmosphere have on the trajectory are discussed. For the mission studied here, a payload retrieval from geosynchronous orbit with aerodynamic plane change, a linear feedback guidance scheme was developed. A relatively simple heuristic law was used to demonstrate the concept. Using optimal control theory a feedback law was developed analytically. Testing with a large number of different atmospheres showed this law to be a feasible means of controlling the AMOOS trajectory. Refinements to the technique offer promise of significant improvement, and these are discussed.
A derivation of an explicit solution to the two point boundary-value problem of exoatmospheric guidance and trajectory optimization is presented. Fixed initial conditions and continuous burn, multistage thrusting are assumed. Any number of end conditions from one to six (throttling is required in the case of six) can be satisfied in an explicit and practically optimal manner. The explicit equations converge for off nominal conditions such as engine failure, abort, target switch, etc. The self starting, predictor/corrector solution involves no Newton-Rhapson iterations, numerical integration, or first guess values, and converges rapidly if physically possible. A form of this algorithm has been chosen for onboard guidance, as well as real time and preflight ground targeting and trajectory shaping for the NASA Space Shuttle Program.
The verification process and requirements for the ascent guidance interfaces and the ascent integrated guidance, navigation and control system for the space shuttle orbiter are defined as well as portions of supporting systems which directly interface with the system. The ascent phase of verification covers the normal and ATO ascent through the final OMS-2 circularization burn (all of OPS-1), the AOA ascent through the OMS-1 burn, and the RTLS ascent through ET separation (all of MM 601). In addition, OPS translation verification is defined. Verification trees and roadmaps are given.
The Terminal Configured Vehicle (TCV) program operates a Boeing 737 modified to include a second cockpit and a large amount of experimental navigation, guidance and control equipment for research on advanced avionics systems. Demonstration flights to include curved approaches and automatic landings were tracked by a phototheodolite system. For 50 approaches during the demonstration flights, the following results were obtained: the navigation system, using TRSB guidance, delivered the aircraft onto the 3 nautical mile final approach leg with an average overshoot of 25 feet past centerline, subjet to a 2-sigma dispersion of 90 feet. Lateral tracking data showed a mean error of 4.6 feet left of centerline at the category 1 decision height (200 feet) and 2.7 feet left of centerline at the category 2 decision height (100 feet). These values were subject to a sigma dispersion of about 10 feet. Finally, the glidepath tracking errors were 2.5 feet and 3.0 feet high at the category 1 and 2 decision heights, respectively, with a 2 sigma value of 6 feet.
A simulator study was undertaken to compare and evaluate the design features of the electronic displays for possible use in V/STOL aircraft: a combined transition display (Display A), a perspective display (Display B), and a hover display (Display C). Display B presents height information via integrated elements; Displays A and C present information similarly except that Display C presents vertical and lateral guidance via conventional cross pointers. High pilot work load was attained by evaluating the displays only as situational guidance displays in a simulated vehicle without stability augmentation. Glide slopes of 6 and 15 deg were used, and steady-state and no-wind conditions were randomly presented. Six pilots participated, and fifty-five objective performance measures were taken along with pilot opinions
The method of linear covariance analysis is applied to the study of an aerospace vehicle integrated guidance-and-navigation system. This requires that the rms trajectory deviations of the vehicle be computed along with the rms errors of the navigation system. The technique is used to examine the performance of the shuttle orbiter vehicle's guidance-and-navigation system during the terminal phase of the landing. Essentially the same rms trajectory deviations were obtained from a single covariance run as from a 30-sample Monte-Carlo run set. Of particular interest in the study was the result that rms trajectory deviations for the integrated-system were generally significantly larger than the rms navigation errors.
This paper describes the navigation and guidance system developed for the TCV B-737, a Langley Field NASA research aircraft, and presents the results of an evaluation during final approach, landing, rollout and turnoff obtained through a nonlinear digital simulation. A Kalman filter (implemented in square root form) and a third order complementary filter were developed and compared for navigation. The Microwave Landing Systems (MLS) is used for all phases of the flight for navigation and guidance. In addition, for rollout and turnoff, a three coil sensor which detects the magnetic field induced by a buried wire in the runway (magnetic leader cable) is used. The outputs of the sensor are processed into measurements of position and heading deviation from the wire. The results show the concept to be both feasible and practical for commercial type aircraft terminal area control.
The six themes identified by the Workshop have many common navigation guidance and control needs. All the earth orbit themes have a strong requirement for attitude, figure and stabilization control of large space structures, a requirement not currently being supported. All but the space transportation theme have need for precision pointing of spacecraft and instruments. In addition all the themes have requirements for increasing autonomous operations for such activities as spacecraft and experiment operations, onboard mission modification, rendezvous and docking, spacecraft assembly and maintenance, navigation and guidance, and self-checkout, test and repair. Major new efforts are required to conceptualize new approaches to large space antennas and arrays that are lightweight, readily deployable, and capable of precise attitude and figure control. Conventional approaches offer little hope of meeting these requirements. Functions that can benefit from increasing automation or autonomous operations are listed.
An approach developed to improve the ground instrumentation time tagging accuracy and adapted to support the Minuteman ICBM program is desired. The Timing Insertion Unit (TIU) technique produces a telemetry data time tagging resolution of one tenth of a microsecond, with a relative intersite accuracy after corrections and velocity data (range, azimuth, elevation and range rate) also used in missile guidance system analysis can be correlated to within ten microseconds of the telemetry guidance data. This requires precise timing synchronization between the metric and telemetry instrumentation sites. The timing synchronization can be achieved by using the radar automatic phasing system time correlation methods. Other time correlation techniques such as Television (TV) Line-10 and the Geostationary Operational Environmental Satellites (GEOS) terrestial timing receivers are also considered.
A twin-jet commercial transport equipped with digital navigation, guidance, and control systems and advanced electronic display system was used for airborne operational research. The results of flight tests which evaluated a second-order horizontal-path guidance control law and the autopilot and airplane response are presented. This evaluation was accomplished through analysis of recorded flight data and through pilot opinion of the airplane maneuvers during automatic path-capture scenarios and path tracking. Four different path captures were flown at a ground speed of approximately 160 knots and repeated at approximately 300 knots. Path tracking tracking error was measured in terms of cross track error along two paths: one pat at cruise speeds and one at airport-terminal-area speeds. The path tracking accuracy and the smoothness of the airplane maneuvers were judged satisfactory for high speeds; however, at lower speeds the control law design should be improved so that tracking will be more accurate for operations in the airport terminal area.