DE2250219A1 - FLIGHT CONTROLLER FOR CONTROLLING THE AIRWAY AND AERODYNAMIC FLOW CONDITION - Google Patents

Description

Flight controller for regulating flight path and aerodynamic flow condition The invention relates to a flight controller for controlling trajectory and aerodynamics Flow condition (e.g. airspeed or angle of attack) with measuring devices for aircraft position and aerodynamic flow condition and comparison facilities to compare these variables with reference variables and to generate control deviation signals, Measuring devices for aircraft dynamics (acceleration and angular velocity) and actuators for thrust and elevator.

Common flight controllers have separate control systems for the aerodynamic State of flow, e.g. the speed of sighting in relation to the surrounding air, and intended for the flight path. The airspeed is e.g.

Measured via the dynamic pressure and compared with a reference variable. The control deviation signal acts on an actuator in a propulsion controller for the gas throttle and thus causes a change in the engine thrust in one of the Airspeed control deviation counteracting meaning. Completely independent of this, the control system for the slug track control works. This contains a measuring device for the position of the aircraft, e.g. an altimeter. The signal thus obtained is compared with eim: reference variable and the system deviation signal formed in this way acts on an autopilot which, via a servomotor, acts on the elevator, for example acts. If there is an altitude deviation, the elevator is then operated, see above that the aircraft goes into climb or descent and thus the control deviation corrected.

It is also known on the propulsion controller and the autopilot to connect additional status variables to improve the control behavior, For example, in a known propulsion controller, the control deviation signal the signal from a longitudinal accelerometer counteracted to the thrust smoothness to improve. Namely, when the aircraft experiences a gust of wind from the front, so that it is temporary the relative speed between the aircraft and the surrounding air is increased with that at the same time Plane delayed against ground, so that a negative acceleration signal arises. You can put these signals on top of each other agree that they compensate and the controller to such ßöen - öhnlicgh as the human pilot - does not react by adjusting the throttle (PS 1 190 199).

There are also longitudinal acceleration signals with other propulsion regulators switched on in order to obtain a damping of the regulation, woberi the longitudinal acceleration signal replaces the time derivative of the speed signal, which is otherwise difficult to form. On an autopilot new. the servomotor only the elevator is for example for damping, a pitch angle speed that is tapped from a rate gyro proportional signal applied.

These known flight controllers or propulsion controllers and autopilots have the disadvantage that the apparatus-wise separate control systems for airspeed or the like and trajectory coupled to one another via the behavior of the aircraft are. A singular in one control system causes a disturbance in the other and vice versa It is known from modern control theory that an optimal Control receives when all state variables relevant to the behavior of the controlled system connected to all existing actuators in a suitable linerakomizinatiorl are. In practice, however, this theoretical requirement is usually not feasible, because in the case of complex rule systems, the number of those to be taken into account State variables become too large and some of these state variables are not without further is measurable.

The invention is based on the object of a flight controller for precise Control of flight path and aerodynamic flow condition (e.g. flight path speed, Angle of attack or lift coefficient).

According to the invention, a flight controller of the type mentioned is designed so that that from the control deviation signals control technology state variables through direct Measurement, filtering (observation filter) or model simulation can be generated and that each of these state variables is switched to each of the servomotors.

The invention is based on the knowledge that for the regulation of flight path and aerodynamic flow condition in an aircraft are relevant control-related state variables in a level that is sufficient for precise control Either measure the selection directly or obtain it from the measured values and thus an integrated flight controller can be created, trajectory and aerodynamic Allowed to regulate the flow state at the same time in a desired manner.

Advantageously, each of the state variables is also on one Actuating motor for a spoiler for direct lift control. There are then three servomotors for Disposal made by all educated State variables can be controlled in a suitable linear combination.

A suitable selection of the control state variables is essential for the invention so that, on the one hand, good control of the flight path and aerodynamic flow state is ensured and, on the other hand, the state variables used for the control can actually be measured or obtained. It has proven to be particularly advantageous that the following control-related state variables are formed and connected to all servomotors with different factors: (b) Integral of the control deviation of the aerodynamic flight condition (Jdt) by pure integration of this control deviation, (c) Airspeed (u) by measuring the longitudinal acceleration (b) of the aircraft using a longitudinal accelerometer and by pseudo-integration of the same with elimination of the stationary final value over a High-pass-low-pass filter d) Time derivative of the airspeed (b) from the longitudinal acceleration with elimination of the stationary end value via a high-pass filter (e) Angular acceleration around the pitch axis (gy) by measuring the angular velocity (#y) by means of a turning circle and pseudodifferentiation via a high-pass filter (f) Angular velocity around the pitch axis (#y) through direct measurement using a rate gyro (g) Pitch angle (+ p) through pseudo integration that of the angular velocity (#y) (h) vertical velocity (h) via complementary filter on the one hand by measuring the vertical acceleration (h) by means of a vertical accelerometer and pseudo integration of the same and on the other hand from the height control deviation by pseudo differentiation via a high-pass filter, (i) height control deviation (A h) by direct hook measurement by means of an altimeter and difference formation with a guide variable , (å) Time integral of the height control deviation (iA h dt) through pure integration of the latter.

The elevator deflection cannot be measured directly.

It can be measured which adjustment movement of the elevator servomotor but the actual elevator deflection can be influenced by slack, elasticity and vibrations of the elevator can be very different from the servomotor output. It is therefore advantageous that the elevator deflection is also used as a state variable and its time derivative from the elevator servomotor output by means of a das Behavior of the elevator simulating model route formed and in linear combination are connected as feedback to the input of the elevator servomotor.

As is well known, the rudder effectiveness increases with the standing pressure. Around To take this into account, the elevator servomotor is advantageously one of a back pressure signal applied computer to take into account the increased Dynamic pressure upstream of increased rudder effectiveness. It's the feedback signals of the model route as well as the angle of attack control deviation and its integral directly to the servomotor but the other state variables to the input of the computer.

In general, it is advantageous that from the system deviation signals and the aircraft dynamics signals, control-related state variables are formed, which at least partially redundant the technical control status of the aircraft describe. Control intervention then takes place even in the event of a fault in a channel in the right direction via another state variable.

The articulation angle # 2 from measured Angle of attack ob, vertical speed h and airspeed u according to the relationship 2 u are formed.

The invention is referred to below using an exemplary embodiment explained in more detail on the associated drawing: The drawing shows a signal flow diagram an integrated flight controller designed according to the invention.

The flight controller according to the invention contains an angle of attack sensor 10, accelerometer for longitudinal acceleration t and vertical acceleration h , which are represented by the block 12, a rate gyro 14 for measurement the pitch angle speed and an altimeter 16. The measured angle of attack At 18, α is compared with a setpoint value> setpoint by a setpoint generator 20 and a control deviation signal 8 α is formed. The altimeter signal is accordingly H with a setpoint hsoll from a setpoint generator 22 in the point 24 compared and an altitude control deviation signal h is formed.

There are also three servomotors provided, namely a thrust servomotor 26, through which an adjustment of the gas throttle takes place, an elevator servomotor 28 for adjusting the elevator and a spoiler actuator 30, through which a spoiler of the wing is adjusted, which affects the direct lift. The elevator servomotor 28 is preceded by a computer 32, which receives a dynamic pressure signal is acted upon and performs a division by the back pressure. This is supposed to the increased rudder effectiveness with increased dynamic pressure (increased airspeed) of the elevator must be taken into account.

A high-pass filter 34 is connected upstream of the spoiler servomotor.

Since spoilers or buoyancy clearing only in a given work area should work and only two reference variables should be present, the stationary one must Status of all signals for direct supply control via a high-pass filter 34 be eliminated. The output of the elevator servomotor 28 is #. From the measurands 4 u, h, g y and th a number of state variables are formed. This is done by means of a "pseudo observer", generally designated 34. That pseudo-observer 34 is also the output of the elevator servomotor 28 at an input 36 in Form of a signal 1 switched on. The ones supplied by the pseudo observer 34 State variables are as follows: (a) The control deviation # α of the angle of attack, that in point 18 by forming the difference between the target value αsoll and the measured one Angle of attack α is.

(b) The integral of the angle of attack control deviation ## αdt, which Value of # α formed from point a 18 by means of a pure integrator 38 is formed.

(c), The aircraft speed and this is derived from the longitudinal acceleration t by means of a pseudo-integration via a high-pass filter 40 and a low-pass filter 42 obtained with elimination of the stationary end value.

(d) The time derivative of the airspeed. This is made from the longitudinal acceleration provided by accelerometer 12 upon elimination of the stationary end value obtained via the high-pass filter 40.

(e) The angular acceleration about the pitch axis #y. This is from the Angular velocity signal #y of the tropics 14 by pseudodifferentiation obtained by means of a high-pass filter 44.

(f) The kink angle #. This is also derived from the angular velocity signal #y derived by pseudo integration via a filter 46.

(g) The vertical speed h. This is done in two ways obtained, namely, on the one hand from the vertical acceleration h and on the other hand from the Height control deviation signal 4 h formed in point 24. These two signals will be linked to one another via mutually complementary filters 48 and 50 at point 52, according to the following relationship: s 1 ..

h h = 1 + Ts #h + Ks 21 + Ts s + Ks² 1 + Ks = #h = s h 1 + Ts 1 + Ts With K = T becomes h = s # h It becomes in this way the influence of the noise on the h signal is reduced.

(h) The height control deviation is calculated directly at point 24 from the from Altimeter 16 measured height h and the target height supplied by the setpoint generator 22 hsoll formed by forming the difference.

(i) The time integral of the height control deviation J A h dt is derived from the Signal A h obtained at point 24 by pure integration by means of integrator 54.

(j) From the output signal # of the elevator servomotor 28, which is fed to the pseudo observer 34 via the input 36, are by means of a the behavior of the elevator model track 56 the rudder deflection # and its time derivative # formed.

The elevator model can, for example, be designed as a second-order transducer, which is formed by connecting two low-pass filters 58 and 60 in series with appropriate feedback, with the signal for the elevator deflection 4 at the output of the filter 60 and the signal for the elevator running speed # between the filters 58 and 60 is tapped. The following relationship then results: with = elevator deflection = = output of the elevator servomotor D = damping of the elevator system = = natural frequency of the elevator system 0 The 12 state variables thus obtained are linearly traced back to the servomotors with amplification factors kiγ.

The status variables # and # are thus fed back the elevator deflection and the elevator run speed, as well as the angle of attack control deviation # »And its integral directly to the input of the elevator servomotor 28. Elevator deflection and elevator speed are only available on the elevator servomotor 28 returned. The other state variables are all on all servomotors 26, 28, 30 returned, namely in the case of the elevator servomotor to the input of the upstream computer 32 and in the case of the spoiler servomotor to the input of the upstream filter 34.

The gain factors for the connection of the individual control-related state variables to the various servomotors result from the following feedback matrix. Feedback matrix State Servomotor # α ## α uu #y #y # h #h ## h # # Thrust K1.1 K1.2 K1.3 K1.4 K1.5 K1.6 K1.7 K1.8 K1.9 K1.10 0 0 Spoiler K2.1 K2.2 K2.3 K2.4 K2.5 K2.6 K2.7 K2.8 K2.9 K2.10 0 0 Elevator K3.1 K3.2 0 0 0 0 0 0 0 0 K3.11 K3.12 (direct) Dynamic pressure 0 0 K3.3 K3.4 K3.5 K3.6 K3.7 K3.8 K3.9 K3.10 0 0 customized The various state variables in the described embodiment of the invention are obtained in different ways: some are obtained by direct measurement, that is # α, #y, #h. Other quantities are derived from the measured quantities by filtering (observation filter). These are, for example, the integrals ## αdt and ## h dt or the pitching acceleration or the pitching angle. Other state variables are obtained through a model simulation which, for example, simulates the behavior of the elevator as a model.

In the present case is the selection of the status variables used taken in such a way that 1. precise guidance of the trajectory, 2. precise regulation aerodynamic flow condition, 3. good stability, 4. good gust suppression (in flight path, aerodynamics and thrust), 5. an elimination of the wind shear influence, 6. a simple implementation of the status variables, 7. extensive redundancy, 8. parameter insensitivity over a wide range and 9. a frequency-based division of labor of the various manipulated variables is guaranteed. These aspects can be addressed with the present combination of state variables take into account adequately. Redundancy of the status variables means that certain influences are taken into account several times, so that even with disturbances in one Channel, this influence is still taken into account in the regulation.

In the case of high-frequency processes, #y # # Da both applies, for example # at the output of the elevator model 56 as well as N - are observed in this way a redundant observation of #, which leads to an increase in parameter insensitivity leads.

The pitch angle 4 can also be formed from the measured angle of attack d and the obtained state variables ß and u according to the relation = U

Claims (7)

Claims 1.) Flight controller to control flight path and aerodynamic State of flow (e.g. airspeed or angle of attack) -init measuring devices for aircraft position and aerodynamic flow condition and comparison devices to compare these variables with reference variables and to generate system deviation signals, Measuring devices for aircraft dynamics (acceleration and angular velocity) and actuators for thrust and elevator, characterized in that from the Control deviation signals (# α, 4 h) and the flight dynamics signals (t,?, w) Control technology y state variables through direct measurement, filtering (observation filter) or model simulation and that each of these state variables each of the servomotors (26, 28) is switched on. 2.) flight controller according to claim 1, characterized in that each the state variables also to a servomotor (30) for a spoiler for direct Buoyancy control is switched on. 3.) Flight controller according to one of claims 1 or 2, characterized in that that from the control deviation signals (# α, #h) and the aircraft dynamics signals (u, h, #y) control-related state variables are formed, which correspond to the control-related Describe the state of the aircraft at least partially redundantly. 4. Flight controller according to claim 3, characterized in that the following technical control state variables are formed and connected to all servo motors with different factors: (a) control deviation (a α) of the aerodynamic flight state (angle of attack α) by direct measurement and difference formation with a; Control variable, (b) integral of the control deviation of the aerodynamic flight condition (## αdt) by pure integration of this control deviation, (c) flight speed (u) by measuring the longitudinal acceleration (b) of the aircraft by means of a longitudinal accelerometer and by pseudo-integration of the same with elimination of the stationary end value via a high-pass-low-pass filter (d) Time derivative of the airspeed (u) from the longitudinal acceleration with elimination of the stationary end value via a high-pass filter e) Angular acceleration around the pitch axis (#y) by measuring the angular non-speed (#y) by means of a rate gyro and pseudodifierentiation via a high-pass filter (f) Angular velocity around the pitch axis (X y) by direct measurement using a rate gyro (g) Pitch angle (IQ) by pseudointegration the angular velocity (X y) (h) vertical velocity (h) via complementary filter on the one hand by measuring the vertical acceleration (h) by means of a vertical accelerometer and pseudo integration of the same and on the other hand from the height control deviation by pseudo differentiation via a high-pass filter, (i) height control deviation (h) by direct height measurement by means of an altimeter and difference formation with a reference variable, ( j) Time integral of the height control deviation (# h dt through pure integration of the latter. 5. flight controller according to claim 4, characterized in that in addition the elevator deflection (and its time derivative from the Elevator servomotor output (#) using an output that simulates the behavior of the elevator Model line (56) formed and in linear combination as a return to the entrance of the elevator servomotor (28) are switched on. 6. flight controller according to claim 5, characterized in that the elevator servomotor (28) a computer (32) acted upon by a dynamic pressure signal for consideration the rudder effectiveness, which is increased with increased dynamic pressure, is connected upstream, and that the feedback signals (#, #) from the model line (56) and the angle of attack control deviation (# α) and its integral directly on the actuator (2t)) but the others State variables are switched to the input of the computer (32). 7. flight controller according to one of claims 4 to 6, characterized in that that as a redundant state variable the pitch angle # 2 from the measured angle of attack α, vertical speed h and airspeed u according to the relationship # 2 = α + h / u is formed.