Abstract: A process and apparatus for operating a steam power plant having a coal-fired steam generator, in which at least two partial air streams are heated by means of the exchange of heat in at least one exhaust gas/air heat exchanger with hot exhaust gases that are obtained in the steam generator by the combustion of coal. The coal is dried in a milling-drying system with a first partial air stream of the heated air supplied as a primary air stream and a second partial air stream of the heated air is supplied to the steam generator as a secondary air stream for the combustion of the coal. The secondary air stream is preheated in an air/air heat exchanger upstream from the exhaust gas/air heat exchanger, with respect to the direction in which the secondary air stream flows, by the exchange of heat with the heated primary air stream.

Abstract: The invention relates to a control system to which a state vector representing the states of a controlled system is applied. The control system provides a correcting variables vector of optimized correcting variables. The relation between state vector and correcting variables vector is defined by a matrix of weights. These weights depend on the solution of the state dependent Riccati equation. An equation solving system for solving the state dependent Riccati equation in real time is provided. The state vector is applied to this equation solving system. The solution of the state dependent Riccati equation is transferred to the control system to determine the weights.

Abstract: A method of training a neural network for the guidance of a missile to a target includes the steps of: computing a solution of the non-linear guidance problem in analytical form, generating numerical solutions for a number of flights of a virtual missile to a virtual target, determining a human pilot's behavior and of the missile by simulating a number of flights of the virtual missile to a virtual target, and “cloning” a neural or fuzzy-neural network with the knowledge about the guidance of the missile to the target as obtained by the preceding steps. For determining the behavior of the human pilot, a scenario of missile and target is represented. This scenario is transformed into slow-motion. The flight of the missile to the target is simulated in slow-motion, the human pilot guiding the missile to the target. The pilot's behavior and the behavior of the missile resulting therefrom is stored for a number of such simulated flights.

Abstract: The invention relates to a control system to which a state vector representing the states of a controlled system is applied. The control system provides a correcting variables vector of optimized correcting variables. The relation between state vector and correcting variables vector is defined by a matrix of weights. These weights depend on the solution of the state dependent Riccati equation. An equation solving system for solving the state dependent Riccati equation in real time is provided. The state vector is applied to this equation solving system. The solution of the state dependent Riccati equation is transferred to the control system to determine the weights.

Abstract: A power plant which combusts fossil fuel and biomaterial or waste includes a large boiler plant having at least one device for the combustion of fossil fuel. At least one additional combustion device is connected to the large boiler plant via at least one connection channel. The additional combustion device includes a grate hearth for the combustion of biomaterial or waste. The additional combustion device includes at least one heat exchanger device which is integrated into the water/steam circuit of the large boiler plant.

Abstract: For monitoring flight safety of aircraft, a monitoring device accommodated in a pod is provided. The monitoring device has a monitoring unit to be attached to an aircraft. Sensors for measuring aircraft data independently of the sensors of the aircraft itself are provided in the monitoring unit. There is a device for automatically monitoring the operation of the technical equipment of the monitoring unit. In addition, the monitoring unit contains a system for automatically monitoring the pilot's reactions on the basis of aircraft data provided by the independent sensors. Furthermore, the monitoring unit contains a device which responds to deviations of the aircraft from the range of safe flight states. This device triggers an alarm. There is a device responding to the aircraft inadmissibly approaching ground and a collision warning device, which responds to the risk of collision with other aircraft. Also these devices trigger an alarm in the case of danger.

Abstract: Measuring signals are derived from a plurality of redundantly provided sensors, the sensor signals of which are sampled and digitized at a predetermined rate, and are processed by computers also provided redundantly. Each of the computers processes the sensor signals with a main program and a monitoring program. A first computer is arranged to process in each computing cycle (n) sensor signals sampled in a predetermined sampling cycle (n), which is preferably identical with the computing cycle. Further computers are provided. Each of the further computers processes in each computing cycle n the sensor signals which have been sampled in sampling cycles (n-1), (n-2), (n-3), respectively, which lead the computing cycle by predetermined time intervals. A fault logic serves to make decisions on program faults from the output data obtained with the main program and the monitoring program in the various computing cycles.

Abstract: A filter for obtaining a useful signal, constant with respect to time, from a noisy measuring signal, having a measuring signal integrator and an optimal estimator (54) which represents a model of the integrated measuring signal and to which the difference of the integrated measuring signal, and an estimate of the integrated mesuring signal resulting from the model, is applied, for estimating that coefficient of the model, which corresponds to a linear signal rise.

Abstract: In order to eliminate gyro errors of electrically caged, two-axis gyros, which have a pick-off (52,54) and a torquer (46,48) on each input axis, and each pick-off (52,54) arranged on one input axis energizes the torquer (48,46) on the other input axis, a first measurement is made with a first direction of rotation of the gyro rotor (42). Subsequently, for a second measurement, the direction of rotation of the gyro rotor (42) is reversed and, at the same time, the polarity of the signals from the pick-offs (52,54) is inverted. A measuring value is formed from the gyro signals obtained with both of these measurements.

Abstract: A plurality of two-axis rate gyros are arranged in relation to an aircraft-fixed coordinate system such that they redundantly supply angular rate information. A plurality of accelerometers supply correspondingly redundant acceleration information. First signal processing means comprise means for failure detection and elimination, such that angular rate and acceleration information subjected to error are eliminated. The angular rate and acceleration information thus cleared from errors supply stabilization signals for the autopilot. Second signal processing means obtain the angular rate and acceleration information cleared from errors and integrated if required, and therefrom supply heading and attitude reference signals.

Abstract: The north direction shall be determined by a carrier fixed two-axis rate gyro having a substantially vertical spin axis from the two components of the horizontal component of the rotary speed of the earth. Deviations of the position of the rate gyro from an exact vertical alignment of the spin axis are detected by two accelerometers which provide acceleration signals. The ratio of acceleration signals and acceleration due to gravity provide values of the angles of inclination. Correction signals are formed out of the accelerometer signals and are subtracted from the signals obtained by the rate gyro. One of the correction signals compensates for the component of the vertical component of the rotary speed of the earth falling into the direction of the input axes of the rate gyro. Another signal compensates the rotary speed of the carrier relative to the earth and thus rotatory interferences. The tangent of the true azimuth angle is obtained by forming quotients of the rate gyro signals corrected in this way.

Abstract: By means of a two-axis, electrically restrained gyro with vertical spin axis, North direction is determined from the ratio of the horizontal components of the angular rate of the earth. The gyro is rotatable about a vertical axis into 0.degree.-, 90.degree.-, 180.degree.- and 270.degree.-positions. The signals thus obtained are stored. A signal processing circuit operating with various differences of these signals and ratios of such differences provides the North deviation unaffected by gyro drift and scale factor error.

Abstract: The position of a vehicle is derived from vehicle speed and heading. A two-axis electrically restrained gyro, the spin axis of which is parallel to the vehicle vertical axis, serves at first to determine the north direction with stationary vehicle. The attitude of the vehicle about the longitudinal and transverse axes is measured by means of a pair of vehicle-fixed accelerometers. True north direction is derived from the signals of the accelerometers and of the gyro. Subsequently the gyro is rotated through 90.degree. about one input axis and serves as heading-attitude reference during the mission. The rotary speed about a third axis perpendicular to the input axes of the gyro is measured by means of a rotary acceleration meter the output of which is applied to an integrator. During the mission, the attitude parameters of the vehicle are computed continuously from the initial attitude parameters and the rotary speeds.

Abstract: The vehicle is operated over land which is described by a grid coordinate system having a north direction offset from geographic north. The vehicle has a pick-up to determine its heading speed along its longitudinal center line. The spin axis of a free gyro establishes a reference axis. A meridian gyro determines the angle, if any, between said reference axis and geographic north at the start of a mission. At the start of the mission a computer is supplied with information as to the starting point in said grid coordinate system, the angle between geographic north and grid north, and an estimated corrective factor (based on the type of terrain to be traversed) to be applied to the measured heading speed. The computer continuously determines, for display, the position of the vehicle in said grid system.

Abstract: A highly-amplified pickoff signal responding to positional deviations of a wire-suspended gyro are fed to a torquer connected to the gyro. In response to these the torquer exerts on the gyro a torque counteracting the gyro directing torque. The pickoff signals are sampled at time internals T and an analogue-digital converter produces therefrom a timed sequence of digital signals with each digital signal being proportioned to the pickoff signal at a time of sampling. This sequence of digital signals are fed to a digital filter which produces an output signal, indicative of the deviation, in accordance with the equation:U.sub.M (nT) = U.sub.M ((n-1).multidot.T) +1/n [u.sub.m (nT) - U.sub.M ((n-N).multidot.T]wheren is a consecutive whole number,U.sub.M (jT) the average value output at the time jT,u.sub.m (jT) the digitalized measuring value at the time jT, andN a fixed whole number.