DE1098371B - Aviation training device to simulate the flight angle - Google Patents

Description

Flight training device for simulating the flight angle The invention relates to a flight training device located on the ground with a simulated flight control system and associated simulated flight surveillance instruments as well as one the controls responding flight calculation system. The flight training device according to the invention is based on a flight calculation system that is assigned to the controls Contains tools for the division of electrical signals necessary for representation of flight control factors, e.g. B. the angular speeds of the rotary movements around the aircraft axles, as well as a number of interconnected and electrical systems responsive to the signals, for example for Representation of the angle of attack, as well as electrical integration systems for representation the aircraft's position in relation to the axes of movement, and is characterized by that facilities are provided, both on the angle of attack system as well respond to the transverse axis system and generate a signal that corresponds to the flight angle of the corresponds to the flight to be simulated.

The flight angle is defined as the difference between the pitch and the angle of attack. When solving aerodynamic equations of forces so far for this purpose only the vectors over the pitch angle are calculated while the vectors for the angle of attack were not taken into account. According to the invention the actual flight angle is taken into account in the calculations, which results in a better approximation of the real conditions in solving the aerodynamic Equations of forces results. An embodiment is shown in the drawing; 1 shows a block diagram of the pilot training device with the main instruments, Fig. 2 is a vector illustration of the rotary movements around the transverse axis and around the Longitudinal axis and the decomposition of gravity into lateral components, Fig. 3 a Diagram showing the relationship between different flight angles, Fig. 4 a circuit diagram of the computing device for determining the airspeed, the angle of attack, the longitudinal inclination and the gradient ratio, FIG. 5 is a circuit diagram of the arithmetic unit to determine the side slip angle, the rotational movements around the vertical axis and the Course direction, Fig. 6 a circuit diagram of the angle of attack servo system and the control loops for stormy air, wing freezing, landing flaps and undercarriage, Fig. 7 the servo systems for the sideslip angle and heading, including the slide signal amplifier and the regulation of the magnetic compass and the gyro compass, Fig. 8 the servo system for the rotary movement around the transverse axis together with the course direction signal amplifier and the controller for vertical airspeed, FIG. 9 the servo system for rotary movements around the longitudinal axis together with the airspeed signal amplifier, the aileron potentiometer and the gyrocompass control, Fig. 10 a modified form of the servo motors for the angle of attack for determining the rotational movements around the longitudinal and transverse axes the flight angle and feeding these values into the arithmetic operations.

The servo system There are nine servo units in the pilot training device connected to each other by an electromechanical arrangement (Fig. 1) to a To enable training in the relevant flight principles and flight maneuvers. Depending on the actuation of the simulated aircraft control units Creates tensions proportional to the various forces and movement or acceleration with respect to the three reference axes according to the basic principles of Generate aerodynamics. These three reference axes of the aircraft are: first, the longitudinal axis or "X" <axis, secondly the transverse axis or "Y" axis and thirdly the vertical axis or "Z" axis, with all axes passing through the aircraft's center of gravity.

Displacement and rotation movements with respect to these axes and with Relation to fixed axes that are perpendicular and parallel to the horizon are given by the servo systems generated. In one of these systems, forces are used to determine the Air speed is set; in another servo system, moments are added to the Calculation included to take into account the extent of the roll and in one third moments are evoked to the extent of the inclination to specify. Two auxiliary servomotors are provided to adjust the angle of attack and the to reproduce lateral slippage, with forces acting on the angle of attack servomotor act at right angles to the flight path in order to achieve aerodynamic values such as stroke (weight plus centrifugal force) to calculate the moments of resistance and inclination during the Slip servo motor the angle between the plane of symmetry of the aircraft and the Plane of trajectory 'calculated. The other four servomotors are used to make movements, both linear and angular motions derived from the three above Main servomotors are generated when depicting flight conditions such as altitude, Turning movements around the longitudinal axis, pitch and course direction must be taken into account.

As an introduction, the general basic relationship between the various servo motors and potentiometers is explained, which are used in achieving the combined output voltages, such as e.g. Be used in reading the airspeed meter 24 and gyrocompass 27 before describing the full flight calculation circuits. The representation is considerably simplified in order to give the entire overview, and therefore does not exactly correspond to the embodiment of the finished invention. Airspeed and Altitude Control First, a so-called »altitude control system« is described, which contains the altitude and throttle controls for determining the airspeed. According to the well-known principles of aerodynamics, the air speed (v) is a function of the engine thrust (T), which is always positive (with the exception of the propeller resistance when idling during flight below 1200 revolutions per minute) and the force of gravity (G), which either can be positive or negative depending on whether the aircraft is moving down or up and the drag, which is of course negative. The air resistance can be broken down into two components, namely a resistance with constant coefficient, which changes with the square of the air speed v2, and another resistance, which is expressed by the variable coefficient CD (a) , which changes with the angle of attack a , that is, the angle between the chord of the wing and the airflow.

From Fig. 4 it can be seen that a number of AC voltages, which different values of thrust, gravity and air resistance in accordance with the instantaneous values of the polarity and magnitude of these voltages, are fed individually in an amplifier 100, which is located in the servo system, which is labeled "airspeed". The outcome of the Amplifier100 is used to provide an automatically balancing servo circuit with a two-phase motor 101 to control. The control winding of this known per se Motor is excited by the output voltage of the amplifier, while the other Winding is fed with a constant reference alternating voltage plus e1. The motor rotates in one direction when the control and reference voltages are in the both windings have the same instantaneous value of polarity, and in opposite directions Direction when the instantaneous value of the control voltage compared to the reference voltage is reversed, the speed in both cases depending on the magnitude of the control voltage depends. The motor drives a two-phase feedback generator 101 a, whose one winding is excited by an AC reference voltage + e2 while the other Winding depending on the motor speed a feedback voltage Efb -for, creates a purpose described later. The engine also serves to get over one Reduction gear 101 b to control the contacts of a potentiometer system 102. The pointer of the air speed meter24 is driven directly by the motor adjusted, with suitable mechanical connections 101 c between the engine and the driven element are shown by dashed lines.

The individual potentiometer resistors can be known per se Wise to be wound and have circular or ribbon shape, but are schematically as flat shown. Each potentiometer is shaped or has such an outline that the value of the derived voltage at the potentiometer contact the desired relationship to the linear movement of the sliding contact, which depends on the special task of the Potentiometer depends.

According to FIG. 4, the potentiometer 104 is at its upper end point, which represents the maximum air speed, by a negative voltage -E powered and is grounded at its lower end, so that the diverted voltage represents the value -v2 at the sliding contact 107 and therefore also the air resistance with constant coefficient. This voltage can therefore be used as an input voltage of the airspeed amplifier 100 can be used, the positive shear stress T is opposite, the arrangement being such that when all input voltages be in equilibrium at the amplifier, d. H. if no change in airspeed takes place, the output voltage of the amplifier is zero and the motor 101 is not is excited. Any input voltage change which unbalances the system brings, be it in a positive or negative sense, such as B. a change in Throttle setting in level flight when the thrust and drag tensions are unequal, causes an actuation of the motor 101 in a corresponding direction, to bring the potentiometer contacts into a new equilibrium position in which new derived voltages restore equilibrium at the input of the motor.

To generate a voltage proportional to the airspeed v, the linear potentiometer 103 is excited by a voltage -E, and the sliding contact 106 is adjusted according to the size of the air speed. This derived voltage is used in another part of the system to be described.

The shear stress is derived from the setting of the potentiometer 109, which reflects the engine throttling, the contact 110 being adjusted directly by the pilot in order to carry out the throttle control. This potentiometer is supplied with a voltage that is tapped at the contact 108 of the potentiometer 105, which is fed at its lower end point by a voltage + E, while the upper end point is grounded via a resistor R and is also connected directly to the contact 108 to derive a voltage that is proportional to the airspeed - is. This corresponds to the relationship whereby is. From this it can be seen that the thrust input voltage generally corresponds to the engine power, which depends on the throttle setting and the airspeed.

The coefficient CD of the air resistance changes, as mentioned above, with the angle of attack a. Another servo system called "Angle of Attack" is therefore provided to derive voltages that correspond to a number of factors that vary with angles of attack. A two-phase motor 111 (similar to motor 101) of the angle of attack system is fed from the output of an amplifier 112 in the manner described above in order to actuate a feedback generator 111 a and the contacts 113, 114 and 115 of the potentiometers 116, 117 and 118. These potentiometers are used to calculate the drag coefficient CD, the lift coefficient CL and the moment coefficient CM. The inputs of the a-amplifier 112 are supplied with voltages which correspond to the force of gravity, the force of lift (CL) and the centrifugal force (F i) due to the gradient.

The air resistance as a function of the angle of attack can be expressed by where D is the drag in kg, n is the air density, CD (a) is the drag coefficient and S is the projected wing area. The air resistance is therefore a function of v2, ie the square of the airspeed. To illustrate this relationship, the potentiometer 116 has a corresponding contour and is fed at its ends by a voltage (-v2) which is tapped from the potentiometer 104 of the airspeed system. The central part of the potentiometer 116 is grounded at the angle of attack where the drag coefficient CD (a) is zero, and the contact 113 is connected to the airspeed amplifier 100 through a conductor 113a. The derived voltage at contact 113, which changes with the angle of attack according to the above relationship, can therefore be used as input voltage CD for the airspeed amplifier. The gravity input (G), which depends on the inclination of the aircraft, requires an additional servo system, which will now be described.

The input voltages for the (a) pitch amplifier 112 are discussed below. The factor of gravity, which is affected by the upward or downward movement, can be divided into two components which are applied to the angle of attack and airspeed boosters 112 and 100. In practice, these gravity components are 90 ° components, ie the airspeed component is in the direction of the flight path and the angle of attack component is perpendicular to it. In the present example, the v and a components are derived from two contacts 122 and 123 of the potentiometer 119 of the "tilt" (O) servo system, the tilt amplifier 120 in turn being fed by the tilt change system now described to drive the motor 121 etc. to operate. The inclination potentiometer 119 has such an outline (cosine-shaped in the present case) and is grounded at two points shifted by 180 °, so that it represents normal flight and inverted flight. The potentiometer is fed by voltages -E and -fE at the points between the earthed points, which represent values of gravity for rising (negative) and falling (positive). The derived voltage at contact 122 gives the gravity component - W sin 0 again, which (at small angles of attack) represents the effect of the aircraft's weight as increasing or decreasing thrust and therefore airspeed, and is fed to the v-amplifier 100 via a conductor 122a. The voltage picked up at contact 123, which is offset by 90 ° from that of contact 122, represents the gravitational component W cos 0, which is supported by the lift and derived from the angle of attack and fed through the conductor 123a to the a-amplifier 112.

With reference to the angle of attack system, the lift L (in kg) can be given by the formula where CL (a) is the lift coefficient. The lift is therefore also a function of the square of the airspeed and depends on the properties of the aircraft being simulated. The potentiometer 117 of the a-system for determining the lift coefficient is therefore designed accordingly to simulate the coefficient CL (a) of the aircraft in question, and is grounded at its midpoint at the value of the angle of attack at which the lift coefficient is zero. It is fed with voltages -v2 and + v2, which are derived from the airspeed potentiometer 104, at its upper and lower end points. The positive instantaneous value of v2 can be obtained with the help of a 180 ° phase shifter. The contact 114 of the potentiometer 117 therefore generates a lift force voltage which is fed to the input of the a-amplifier 112. The angle of attack system is also supplied with an input voltage representing the centrifugal force, and this input voltage is derived from the servo system called "change in inclination" (coy), which takes into account the rotational movement about the transverse axis, the centrifugal force being the product of wy and v .

The input voltages of the system, which takes into account the rotary movement about the transverse axis, contain a so-called inclination torque input, which is derived from the potentiometer 118 of the angle of attack system. This moment of inclination is also a function of the square of the airspeed.

The potentiometer 118 is grounded in the middle part at the angle of incidence at which the tilting moment is equal to zero, and is fed by voltages (-v2) and (+ v2) similar to the potentiometer 117. It also has a similar outline so that the tension of the pitch moment on sliding contact 115 changes according to the characteristics of the aircraft concerned. This voltage is fed to the amplifier 125 via the conductor 115 a. Furthermore, the pitch moment Mp in mkg, which is generated by the altitude control when actuated by the pilot, goes into the amplifier 125. This change in pitch is derived from the altitude potentiometer 124 which is fed by voltages equal to the square of the airspeed + v2 and -v2. The middle part of the potentiometer is grounded to represent plane flight or zero inclination. The sliding contact 124a of the height potentiometer picks up a voltage which can be represented as inclination moment (Mp) in mkg and which is fed to the inclination change amplifier 125. A positive sign represents an increase in airspeed, angle of attack, change in pitch and pitch in the usual positive direction of FIG. 8.

As with the previous servomotors, the output of the amplifier feeds 125 a two-phase motor 126 to a feedback generator 126 a and the contacts 127 and 128 to be driven by potentiometers 129 and 130 via a reduction gear 126b. The linear potentiometer 129 is used to derive an input voltage, which represents the centrifugal force for the α-booster, and is therefore in the Middle grounded and is according to the airspeed of voltages + v and -v fed so that the derived voltage at the sliding contact 127 by the factors v and cuy is determined and is fed to the a amplifier via a conductor 127 a linear potentiometer 130, which supplies the input voltage to the inclination-O-system, is fed by voltages + E and -E, with the derived voltage at the sliding contact 128, which is fed to the i9 amplifier via line 128a, is proportional to the Is a change in inclination, the value of which, integrated over time, equates to the inclination angle 0 of the aircraft. This integration is performed by the tilt motor 121 and the feedback generator 121 a, wherein the potentiometer 119 voltages the sliding contacts 122 and 123 generated, which are not only the two gravity components but also the instantaneous value of the angle of inclination. The inclination element the gyrocompass can therefore be operated directly by the tilt motor.

The change in the various forces and moments, such as gravity, buoyancy, centrifugal force, thrust, air resistance, inclination moment, etc., are therefore generated by changes in the sliding contact of the various potentiometers together with changes in the potentiometer supply voltages, whereby the relative size of the forces and moments mentioned depends on the Depends on the value of the input resistance of the various amplifiers. As an example it should be mentioned that the relative size of the lift depends on the values of the air density (9) and the constant factor depends. In the present example, (n) is a constant and therefore these values determine the resistance of the input CL of amplifier 112. Decreasing the resistance increases the relative magnitude of the above constants.

If the density o is to be made dependent on the height, then this can be done happen through a variable formation of the input resistances in all circuits, which contain n, and by changing these resistances in accordance with the Change from 0. However, it is usually easier to find the expression QQv11 by that the derived voltage v2 in a servo system, which depends on the altitude, will be changed. This can be achieved by making a drag outline of the elevation servomotor is fed with the voltage v2 at the end corresponding to the height zero, and that a voltage is tapped depending on the position of the potentiometer contact, if the contact with increasing height in the direction of the grounded end of the resistor is moved towards. In this way, the derived voltage v2 is changed and depending on a decrease with o of increasing height.

The use of the feedback generator for change control is particularly important, using the pitch servo integration system as an example can. If motor 121 were present alone to do the integration, then the inertia inherent in the drive device would be such a large error cause the system to be unusable from a practical point of view. if however, the feedback generator is built into the system as shown then the feedback voltage Efb forms a slope voltage for the slope amplifier and has such a phase relationship to the summed or resulting input signal, that it is directed against it, d. H. that there is kind of a negative feedback works. Therefore, if the gain in the control amplifier is large, the speed of the motor has a linear relationship with the magnitude of the input signal; H. to change the slope voltage without lag or lead, so that both large and small changes in the inclination can be integrated with the same accuracy. It is it can be seen that with a reversal of the main input signal and an actuation of the motor and generator in the reverse direction the phase of the feedback voltage is also reversed to counteract the input signal.

The operation of the circuit shown in Fig. 4 with reference to the The airspeed meter reading will now be described. If on a flight at a constant height z. B. the throttle is opened further, then the Airspeed increases, and the top of the aircraft is raised while at Closing the throttle the opposite occurs. If, according to Fig. 4, the contact 110 of the Throttle potentiometer z. B. is moved down to open the throttle, takes the derived thrust input voltage T for the amplifier 100, thereby increasing the airspeed servo system goes out of balance and the servo motor 101 runs in such a direction that that the potentiometer contacts 106, 107 and 108 are moved upwards in the sense an increase in airspeed.

The following process takes place in the airspeed potentiometer system 102: 1. The derived airspeed voltage v increases; 2. the derived voltage v2 increases with the square of the airspeed; 3. the derived voltage, which is the reciprocal of the air velocity corresponds to, decreases, and 4. the airspeed meter 24 shows a higher airspeed value.

However, the airspeed cannot increase infinitely because the air resistance increases with a constant coefficient corresponding to v2, as does the CD (a) air resistance. At the same time, the thrust, which changes reciprocally with the airspeed, decreases when the new equilibrium is reached. If the values of v and v2 increase, then the angle of attack system is again out of equilibrium, since the input voltages from the potentiometer 127 of the inclination change system and from the potentiometer 117 of the angle of attack system, which both depend on v and v2, have increased.

The gravity input of the inclination system also changes, with the a-servomotor 111 seeking a new equilibrium position and moving the potentiometer contacts 113, 114 and 115 downward so that the angle of attack is reduced. During this process, the derived voltages from the three a-potentiometers 116, 117 and 118 are used as follows: 1. The derived air resistance voltage (negative) from potentiometer 116 is used as the input voltage (CD) to the air velocity booster and increases in size so that it counteracts the increasing shear stress (positive) derived from the larger throttle opening.

2. Since the wing lift of an aircraft has to balance the centrifugal force and weight component which acts perpendicularly on the wing, the derived lift voltage (CL) of the potentiometer 117 has to balance the gravity factor Ga and the centrifugal force F. Assuming the airplane is in a level flight, the centrifugal force is zero and the increasing airspeed tends to reduce the angle of attack, which therefore becomes more negative.

3. The derived torque voltage of potentiometer 118, which is used as Input voltage (CM) of the slope change amplifier is used, becomes with decreasing Angle of attack more positive and therefore causes a balance disturbance of the pitch change servo system. This moves the contacts 127 and 128 upwards to contact the contact 127 a centrifugal force tension F, look for the booster 112, which is based on the a servo motor acts in the sense of a '' restoration of balance. Simultaneously the upward movement of contact 128 produces an increase in input voltage to the integrating tilt servo system O. All four servo systems participate now on a combined calculation and integration process, which is necessary to improve the behavior of the aircraft at the new airspeed and inclination determine.

When the inclination system is unbalanced in the sense a stronger positive inclination, d. H. for the ascending flight, then change the derived voltages at contacts 122 and 123, which are the gravitational (Weight) input components for the v and a - # 'represent stronger in their Size, with the v-component increasing in the present case and the a-component decreases. This is evident from the fact that when the top of the plane goes to the zenith would show the weight component in the direction of aircraft movement then equal - W would be, and the face component perpendicular to the wings, i.e. H. the a servo component, would be zero. When the aircraft is in intermediate positions, the components vectorially decomposed.

The negative weight component (-W sin O) of the airspeed servomotor tends to reduce the maximum speed of the aircraft that the aircraft wants to achieve by opening the throttle more. At the same time, the required wing lift is reduced, since the value H 'cos O of (Ga) at the (a) amplifier 12 decreases. This reduced wing lift allows for a further decrease in the angle of attack and an additional decrease in the negative pitch torque (Cm) on pitch change amplifier 125. This creates more upward movement of contacts 127 and 128, increasing the effect on the pitch and pitch servomotors until eventually overdrive these servomotors and have caused too great a change in weight components for equilibrium. The airspeed therefore decreases again, which in turn results in a reduction in the lift voltage (CL) in the a-amplifier 112, so that the angle of attack increases and generates a greater negative pitch torque voltage at the potentiometer 118 for the («) y) -amplifier 125 will. The (cvy) contacts 127 and 128 now move down to control the tilt integrating servo so that the tilt position is decreased until it eventually becomes negative. The W sin O component (G 1;) of the airspeed servomotor has now become positive, so that it supports the thrust and the airspeed increases again. The process is reversed, so that an equilibrium position is finally reached with damping which corresponds to the new throttle setting.

In this way the real vibration becomes equivalent to a damped vibration Orbit of the aircraft with vertical oscillation reproduced exactly, so that the replica largely corresponds to reality. The degree of damping of the undulating Orbit depends on the choice of the constants of the circuit including the percentage the speed feedback, the gear ratios, the relative input variables and the position of the potentiometer center taps.

In the above explanations, it has been assumed that only the throttle setting has been changed and that the altitude controls are in normal attitude or neutral Location remained. If the altitude control is adjusted, then a derived, The voltage corresponding to the turning torque is used to drive a servomotor for the change to control, d. H. the pitch change servo motor, one of which is the centrifugal force corresponding voltage is derived. This voltage controls as the input voltage the pitch servo motor to generate a change voltage of the opposite sign, but the same size as the first moment voltage to be derived. This same centrifugal tension controls the derivation of another a-input force voltage, which is the buoyancy represents and which has an opposite polarity and increases to the effect to compensate for the original tension. This essentially shows how a balance is achieved between the change in incline and the angle of attack will.

Operation of the height control will now be described in detail. If the altitude control z. B. is set to flight downhill, then the contact 124a, and the derived altitude control potentiometer voltage, which is the pitch moment represents - assuming that contact 124a is originally in one position for ascending flight - only decreases in size, except for level flight and then reverses its polarity and increases in the opposite sense. Through this the incline change system is unbalanced and causes the Motor 126 operates in such a direction that contacts 127 and 128 follow hike below, the system seeking a new equilibrium position. If the Contacts get below the center or zero position of the incline change system, which correspond to the horizontal flight, then the derived voltages have instantaneous values opposite polarity as in the rise position, and the servo motor 121 of the tilt system, which is fed by the potentiometer 130, is now running in the Direction after negative inclination (downward flight), so that the derived Voltage at contact 122 increases, i.e. i.e., the weight component (-W sin 0) of the v-system becomes positive and increases the airspeed. The engine 111 of the a-system, which receives a control signal from the potentiometer 129, now runs in the opposite direction Direction towards negative a values. This process causes the Cm-voltage, that is fed to the incline change system becomes more positive and the system closes stabilize seeks. The movement of the servo motor has at the same time the derived voltage at contact 113 of the CD potentiometer changed so that the air resistance input is changed on the v-system and the airspeed reading is postponed. The interlinked processes described above between the four servo systems therefore repeat until the airspeed, the angle of attack and the aircraft attitude correspond to the setting of the drive and elevator controls.

During the operation of the altitude control described above, the search a system a position of equilibrium that depends on the input voltages that the Centrifugal force of the inclination change system or the gravity component of the Inclination system on the one hand and the lift coefficient of the changed angle of attack on the other hand derive. The resultant of these input voltages actuates the Motor 111 in positive and negative directions and causes compensation when the Systems of slope change and slope are stabilized.

It can therefore be seen that, apart from important operating voltages, the control voltages of the aforementioned interconnected servo systems of the throttle and height control potentiometers are derived, so that the movement in is controlled in a perpendicular direction and therefore a "perpendicular" system is created which is different from the azimuth or sliding flight system. The above description the operations of the airspeed servo system including the airspeed meter is intentionally simplified to show how the servo systems work together to facilitate. Each of these systems represents a certain flight condition or a system that revolves around a specific axis of the aircraft.

The airspeed meter reading and hence the reading vertical airspeed and altitude do not depend in the above system only from the thrust component of the engine, but also from retarding or changing Components, which in turn are affected by the angle of attack, the change in inclination and the angle of inclination as well as the height control are conditioned. A change either these factors or one of the components inevitably influences the relationship to them standing systems, so that they get out of balance and the complete System is constantly looking for a new equilibrium. This makes the aerodynamic Equilibrium of the aircraft simulated. Shift control system As mentioned above, can the flight calculation system of the invention roughly in the just described "vertical" System, a system for displaying the rotational movement around the longitudinal axis, which is described below is referred to as a "taxiing system" and a "sliding flight" system is split up, which is aimed at the turn or a change in the azimuthal position. As in the case of the vertical system, the sliding flight system is shown in Fig. 5 as an introduction shown in the complete flight calculation circuit.

Referring to Figure 5, a shift flight change servo system (ooz) is provided with a number of input voltages including an AC voltage representing the turning torque (Mt) of the rudder potentiometer 140. This potentiometer is grounded in the middle to represent straight flight and is fed by an alternating voltage which has instantaneous values + E and -E on the upper and lower terminals. The derived voltage at the rudder-dependent contact 141 corresponds to the sliding flight effect or the turning moment caused by the right or left rudder, and forms the input (Mt) for the ü9, amplifier. A change in the side control voltage brings the o), system out of balance, assuming that there is no compensation change in the other inputs, so that the output of amplifier 142 causes two-phase motor 143 to move potentiometer contacts 145, 146 and 147 to a new equilibrium position with respect to the relevant potentiometers 148, 149 and 150 , that is to say, for example, upwards with the right rudder. As in the servo systems described above, the motor drives a feedback generator 143a and adjusts the potentiometer contacts via a gear unit 143b and mechanical connections 143c.

The other inputs for the amplifier 142 includes the feedback voltage E fb of the generator 143 a, to ensure that the motor 143 is running at the correct speed, further comprising a voltage representative of an attenuation factor (1), which counteracts the increase of sliding, wherein this voltage is derived at a contact 146 of a potentiometer 149 which is grounded in the center and is fed by fixed voltages -E and -fE at the upper and lower terminals, and also a voltage which represents a second damping factor (2) , ie the reaction force of the trunk against the pushing, which is derived from a slip servo system (ß).

The ß-system contains a servo amplifier 151, the output of which is a Two phase servo motor 152 feeds, and the feedback generator connected therewith 152a drives the sliding contact via a gear 153b and mechanical connections 153 of the potentiometer 154 to operate. The input voltages for the ß-amplifier 151 consist of a feedback voltage Efb and a voltage equal to the represents the damping factor (2) mentioned, which is derived from the potentiometer contact 153 will. The potentiometer 153 is grounded in the middle and at its top and bottom Voltages -E and + E fed so that when ß increases or decreases, the derived Tension, which represents the aforementioned reaction force of the trunk, or the damping factor (2) changes accordingly. The voltage is fed through line 153a to the β amplifier and fed to the o), amplifier.

Another input voltage for the ß-amplifier consists of one Gravity factor referred to as W sin 0 cos O, where 0 is the roll angle and O is the angle of inclination. This input voltage depends on a roll servo system from, which is omitted in the present illustration for the sake of simplicity. It is sufficient, at the present time the finding that this factor is the component of gravity represents that in the usual X-Y plane of the aircraft (Fig. 2) in the direction lateral slipping acts.

Finally, an input voltage for the ß-amplifier is that of the Centrifugal force is derived from the potentiometer 148 of the w;, system. This potentiometer is grounded in the middle and at its top and bottom terminals fed by voltages -v and + v, which are determined by the airspeed system of Fig. 4 are derived, so that the voltage at contact 145, which the ß-amplifier over the line 145a is fed, the product of c) 2 and v, d. H. the centrifugal force, is equivalent to. It can therefore be seen that the ß-system, as in practice, is the most important Factors of gravity, centrifugal force and trunk reaction force or Represents damping, while the co "-system the factors of the torque as a result the Rudder control and the hull response or damping component.

It can be seen that the heading angle by integration the shifting flight change over time can be obtained. This is done through a Course direction servo system (7p), which is a simulated gyro compass which can be set according to the integrated heading angle. The, p-amplifier 155 is arranged as in the previous cases so that it the motor generator assembly 156-156a drives so that via the transmission 156b a Directional element 157 of the simulated gyrocompass is adjusted. The main input voltage to amplifier 155 is taken from the linear potentiometer 150 of the (o, system, which is grounded in the middle and to which the voltages -f-E and -E can be added to the upper and lower terminals. The one at the contact 147 is tapped and fed through the line 147a to the amplifier 155 corresponds to a change in the sliding process to the right or left, so that the Heading motor 156 rotates in a direction to align the compass until the rudder is in the middle and a straight flight again is recorded. At this point in time the (% system is in equilibrium and is switched off at the zero position of the shift change, whereby the voltage at contact 147 is equal to zero. The? P system is then switched off and results in a constant setting on the compass.

If it is now assumed that the pilot controls the rudder e.g. B. hard sets to the right, then the input voltage of the rudder torque for the eiz amplifier to the maximum value. This brings the coz system out of balance, and motor 143 moves potentiometer contacts 145, 146 and 147 up when the system seeks a new equilibrium position that corresponds to the change in the shift flight. Since the derived potentiometer voltages change in size, ß- and, p-systems, where the ß-system due to the increasing centrifugal force and the i /) system is actuated by the increasing change in shear. In the assumption, that the roll angle 0 has not changed, the ß-system shows a lateral Slipping, with the motor 152 moving the contact 153 downward in the direction which corresponds to a position of the known ball bank pointer "ball to the left", d. H. "pushing" as a result of insufficient bank slope. This state can go through Enlargement of the roll angle 0 can be corrected by increasing the size of the input voltage the gravity component W sin 0 cos 0 is increased so that the "pushing" Centrifugal force o), xv is counteracted. When these two input voltages are in the right relationship, the ß-system is balanced for a state in which there is no lateral slipping, the contact 153 being in the Center or zero position and the ball of the ball inclinometer is in stands in the middle. The voltage tapped at contact 153 is, as previously described, used as input voltage for the ß-amplifier to represent the attenuation and serves to dampen the lateral slipping and to center the contact 153.

The shear force components which influence the lateral slippage according to the angle of inclination and roll are shown schematically in FIG. Here an aircraft .P is shown in a combined ascent and right-hand roll position. The pitch angle 0 and the roll angle 0 are measured between a horizontal plane given by the fixed axes X and Y and the X and Y axes of the aircraft. The vertical gravity or weight vector W from the center of gravity can therefore be projected onto the YZ plane of the aircraft as W cos O, and this component in turn can be projected onto the XY plane as W cos O sin 0. This last-mentioned component therefore represents the component of gravity, which acts along the wing and causes the lateral slipping. If this component exceeds the force of gravity and the damping component, then the side slip or the fall occurs. If it is smaller than this component, the lateral sliding takes the form of "pushing". At a given rate of turn, the centrifugal force o) zv becomes a direct function of the airspeed. In order to perform a turn without sliding sideways, the bank angle must be increased with increasing airspeed so that the expression W cos O sin 0 compensates for the force (,) "v. This results in an important simulation of the actual flight conditions Feedback voltages of the ß and (,) z servomotors and the transmission ratios can be made to vibrate the calculation system in order to reproduce the lateral vibrations of an aircraft with their damping.

Pitch Change Servo Motor and Elevation Control The pitch change (coy) servo system is shown in Fig. 4 together with the pilot controlled altitude potentiometer. Since the derived potentiometer signals are usually relatively weak, is it is desirable to have some important and frequently used signals such as v and v2 e.g. B., to strengthen.

The summed wy signal activates the MG set of the servomotor in order to adjust the sliding contacts of the oy potentiometers with linear characteristics. These potentiometers supply voltages in phase opposition, which represent a change in inclination corresponding to o y , which are resolved in the roll servo system in FIG. The third resistor also supplies a voltage for the acceleration error system of FIG. All coy resistors are grounded midway between the two opposing fed terminals to represent zero slope change.

Angle of attack servo motor, strong air movements, icing and flap control The angle of attack servo system of FIG. 6 is mainly influenced, as in the case of FIG by buoyancy, centrifugal force and gravity. The system contains three potentiometer windings energized by the v2 and (v2 + C) signal amplifiers and a pair of cam switches 360 and 361 which operate in response to a be to the device of moving air and also the incline changing system to control when the critical angle of attack is reached. For more precise control can be the ratio of the angular movement of the servo motor shaft to the actual Change in the angle of attack exceed the standard value, e.g. B. a relationship from 4: 1.

The winding (1) is fed with v2 at two terminals offset by about 165 °, this voltage being modified by the angle of attack in order to represent the lift CL at the input of the amplifier. The sliding contact 238 is in the neutral position about 10 ° (actually 21/2) above the center tap (in the direction of a) in order to simulate the normal state of positive lift, ie + a during plane flight. The resistor (2) is fed with a voltage (v2 + C), the derived voltage representing the tilting moment Cu and being fed to the input of the amplifier. The outline of this resistance simulates the characteristic properties of the aircraft (INT inclination moment coefficient), especially with critical values of a. Mathematically, the contour of the resistance can be expressed as a derivative of the tilting moment, which is plotted against a. Sudden changes in incline, such as B. at the critical point, can be simulated in that resistors such. B. the resistor 398, can be inserted into the potentiometer circuit. The sliding contact 239 of the resistor (2) is in the neutral position about 4 ° (actually 1 °) below the center tap (in direction a) in order to simulate a small negative moment of inclination, i.e. when the center of gravity of the aircraft is close to the pressure point ( CP) lies. The resistor (3) is fed at its opposite ends by a voltage -v2, and the derived voltage, which represents the air resistance CD , is used as a negative input voltage for the v-amplifier. The pitch change and airspeed servomotors are controlled according to the change in the angle of attack, as shown in FIG.

The a-input voltages also contain the modified v2 voltage of resistor (1) a negative velocity feedback voltage Efb, a positive voltage to increase the angle of attack if icing is reproduced becomes, a critical acceleration voltage at the terminal 206, which in case of need is derived from the flight speed servo motor to provide a "critical" To bring about state faster, a voltage vaoy at terminal 229 of the coy servo motor, which represents the vertical centrifugal force, a gravitational tension E cos 0 at terminal 235, a gravity load factor voltage E cos 0 cos 0 at the Clamp 236, which represents loss of lift due to rolling, and a Buoyancy component tension from the flaps.

The input voltage, which represents the »wing freezing«, is derived from a potentiometer 362 which is fed with a positive voltage + E who strives to make a more positive. The potentiometer sliding contact 362 'is adjusted by a reversing motor 363 via a reduction gear, see above that the icing voltage fluctuates between zero and the positive maximum value can. The icing motor is used in a suitable manner for the simulated icing and de-icing controlled, e.g. B. by a switch 364 on the pilot's seat for the De-icing or wing heating, which is in series with motor winding 365. at actuation of the switch rotates the motor in such a direction that the derived voltage is lowered. A switch 366 of the flight instructor is located in series with reverse winding 367 and also in series with switch 368. This one Switch is mechanical, as indicated at 369, with switch 365 of the pilot coupled so that one switch is on when the other is off is. The pilot's control switch is therefore superordinate, and icing can occur only enter when the pilot's switch is off. The motor windings can therefore individually from a voltage source Ea @ by the pilot or under certain requirements are switched on by the flight instructor for icing and de-icing will.

The icing motor is also used to apply a negative voltage the potentiometer 370 to decrease the airspeed by the air resistance increases with increasing icing. The sliding contact 371 taps the negative voltage for this purpose depending on the amount of icing.

The arrangement for strong air movements (R. A.) can be both automatic in a critical condition as also arbitrarily activated by the flight instructor will. It should be shown that the moving air mainly changes the inclination and affects the sliding flight system, for it is clear that the effects of disturbances can also be introduced in other places by moving air and also through physical influence on the pilot. The disturbances are caused intermittently, preferably in an arbitrary manner by appropriately shaped motor-driven Cams 372 and 373. The drive motor 374 is from the voltage source E. operated by either the a-switch 360 or the switch 375 of the teacher who parallel to it. The instructor's switch is operated by a button 376 of his Control panel controlled so that after closing switch 375 the intensity the disturbance can be increased by moving air. For that purpose is the button 376 mechanically via the gears 377 and 378 with two rotatable sliding contacts 379 and 380 connected by potentiometers 381 and 382. Each potentiometer is incremental and alternately by opposing voltages via corresponding cam switches fed. The cam switch 372 z. B. applies a positive and a negative voltage + E and -E to the relevant potentiometer 382 during each cam disk rotation, so that the supply voltage at each potentiometer over a wide range changed in paragraphs. The voltages for moving air that are applied to the potentiometer sliding contacts 379 and 380 are removed and the intensity of which is changed by the flight instructor appear as input voltages for the slide change and Inclination change system to influence these systems including instruments, which are thereby operated in a realistic manner. The reason for the operation of the R.A. motor by the a servo motor is the imitation of fluctuations or Impact of the aircraft in critical conditions.

The simulation of the actuation of the flaps is made by three potentiometers 383, 384 and 385 caused the drag, the pitch moment and the Represent buoyancy. The sliding contacts are with a common drive shaft 389 connected to a control in accordance with the operation of a flap control handle 390 to be brought about by the pilot. The simulated position of the flaps influences hence the airspeed through the air resistance potentiometer 380 the moment potentiometer 384 the slope change and through the lift potentiometer 385 the approach angle. The air resistance potentiometer 383 is with its low voltage terminal Grounded through a resistor 391 which is short-circuited by the pilot switch 392 can be used to indicate the retracted or extended position of the chassis. When switch 392 is in the up position, the low voltage terminal is of the potentiometer 383 is directly earthed to zero the air resistance on the chassis and in the lower position of the switch the resistor 391 in switched on the circuit in order to increase the voltage on the sliding contact 386 and the enlarged Air resistance with the chassis extended to reproduce.

In order to visibly indicate the position of the flaps, be it for the pilot or the flight instructor, a switching contact 393 on the shaft 389 can be used for the flap actuation be attached, which turns the signal lamp 394 on and off. A signal lamp can also be operated in a similar manner by the chassis switch. By the shaft 389 is also actuated a switch 395 at the top or bottom Position of the flaps to cause a sliding correction. In the upper position a tension representing the angle of attack changed by rolling, of the terminal 257 is fed to the airspeed servomotor in the normal way, while in the lower position the terminal 257 is grounded via a resistor 396 so that the glide correction voltage for the airspeed servomotor is interrupted. Heading servo system, gyrocompass and magnetic compass with turning error The heading servo system (zp), Fig. 7, controls both the gyrocompass 27 via a direct mechanical connection with the servo motor as well as the remote controlled ones magnetic compass (RMC) 59 via an electrical synchronous transmission and reception system. The RMC is also operated through acceleration and flight maneuvers influenced by north rotation to eliminate the known errors of rotation and acceleration mimic the compass reading.

The input voltage for the heading servo amplifier contains a negative speed feedback voltage E fd and a signal voltage 7; , sec 0, ie sec 0, which represents the angular velocity of the aircraft about a fixed vertical Z axis according to FIG. Therefore, when the heading signal is zero, the aircraft heading is constant. The signal voltage is obtained in the following way: The heading signal amplifier ip (Fig. 8) is fed with input voltages from the aileron servo system which represent the rate of turn components - # - (,) y sin 0 and ± coz cos (P, and the summed output des Signal amplifier, the or? p, excites the primary winding 454 of the transformer 455. The secondary winding 456 of the transformer is grounded in the middle and connected at its voltage points to the individual quadrants of the secant resistor (2). The V signal is therefore changed by the secant resistor (2) of the 0 servo system to correspond to the aforementioned heading signal voltage sec O ie yj sec 0 or (o), cos 0 + coy sin 0) sec 0. The heading servo system therefore functions as an integrator to detect changes in the direction of the aircraft, the heading servo motor being directly mechanically connected at 29 to the gyro 27 as mentioned. The heading generator feedback circuit is connected to a low voltage circuit 418, as in the described β servo system.

Three course direction potentiometers are provided. The cosine resistor (1) is fed by two voltages -v cos 0 and -fv cos 0 of the resistor (1) of the tilt servo system (Fig. 18). Two derived voltages v cos 0 sin y and v cos 0 cos y are taken from this resistor at the contacts 431 and 432, which are offset by 90 ° on the resistor. The linear resistance (2) for determining the north rotation error is fed by voltages of opposite polarity, which correspond to the centrifugal force vcoz. The centrifugal force voltage from the to, servo system of FIG. 23, which is used as the input voltage for the β signal amplifier, is fed via a conductor 433 and amplified by the north rotary error amplifier-419 to close the resistor (2) at points 180 ° apart dine, which represent the north and south azimuth positions. The centers of the resistance, which are also 180 ° apart, are grounded to reflect the east and west azimuth positions at which no north rotation error occurs. The voltage derived from the resistor (2) can therefore reflect the north rotation error. The linear resistance (3) is used to introduce the acceleration error. This resistor is also fed at two points offset by 180 °, namely from the output of the acceleration error amplifier 420. The maximum voltage and the earthed sections of the resistor (3) are aligned with those of the resistor (2), but the sliding contact of the resistor (3) is offset by 90 ° compared to that of the resistor (2), since in the case of an acceleration error the maximum error occurs in the east and west direction, while there is no acceleration error when flying in the north-south direction. The acceleration error amplifier has two input voltages, one of which is derived from the generator feedback circuit of the airspeed servomotor and that is, the acceleration represents, while the other represents the centrifugal force ± vo) y due to the inclination. In addition to the gyrocompass 27, the azimuth servomotor also operates the rotating coil 421 of the synchronous transmitter 414. The coil 421 is fed by a suitable voltage E via a slip ring connection. The stationary winding of the transmitter contains two coils 422 and 423 which cross each other and are electrically connected via slip rings to the crossed coils 424 and 425 of the rotatable element of the synchronous receiver 426. A fixed coil 427 is fed by the voltage E, like the transmitter coil 421, and is inductively coupled to the cross winding so that the latter follows the angular movements of the transmitter coil. The display element of the RMC 59 is mechanically connected in the manner shown to the cross-coil assembly 424 and 425, and a suitable damping device, such as. A magnetic damper 428 is attached to the common shaft to realistically dampen the RMC pointer, especially when the transmitter requests a sudden change of direction.

Pitch servo system and vertical air speed display The pitch servo system 0 together with the heading signal amplifier and the amplifiers for the vertical airspeed meter are shown in Fig. 8 shown. The system assumes that the change in inclination wy is integrated and control the pitch element of the attitude gyro.

The 0 servomotor control amplifier input voltages include a voltage at terminal 217 which represents the pitch moment when the aircraft is tilted downward, this voltage being introduced when v is approximately 65 km (40 miles) per hour or less and the instantaneous value of Polarity is determined by the 0-cam switch 445. Another input voltage is the roll-modified slip flight change voltage co, sin 0 at the terminal 250, which represents a vertical tilt component about the x-axis due to the shear. A further voltage is the inclination change voltage coy cos 0 at the terminal 251, which is changed by the rolling, which represents the vertical inclination component as a function of the inclination about the y-axis. The resultant of the added input voltage represents the time derivative of O and causes the inclination servo motor including the four potentiometer resistors of the inclination system to be actuated. If the servomotor amplifier is supplied with a negative feedback voltage E fb proportional to the speed of the servomotor, then the latter moves the four sliding contacts of the potentiometer resistors according to the integrated value - the sum of the input voltages.

The four potentiometers have the following tasks: The resistor (1) has a sin-cos form and is fed by the air speed signal and influenced accordingly (9 to a) pick up a voltage at contact 446, the v sin 0 for the buzzer of the display circuit corresponds to the vertical airspeed, and b) to tap voltages -v cos 0 and + v cos O at contacts 447 and 448, respectively.

The 0 resistance (1) is made of voltages + v and -v on two by 180 ° offset points, with each voltage terminal at 90 ° from the center position the slope is offset by zero. The resistor is at the 90 ° from the voltage terminals staggered points grounded to ensure level flight in the normal and in the supine position to represent. Characteristic resistances -449 and 450 can be put into the potentiometer circuit be switched on. The resistor (2) has an outline for the secant derivative and is fed by the ip signal. The derived value y) sec 0 at the contact 219 is the input voltage for the integrating heading servo system (Fig. 7). It should be noted that a tip and an -il) signal separate the neighboring Quadrant of the two isolated halves of the potentiometer is fed to the Correct heading reading when 0 changes from 90 to 270 degrees. This Condition occurs when an aircraft is in a straight flight to the north and then loops so that the heading indicator follows South changes, and again to North when this maneuver is carried out further returns. The cosine resistance (3) is fed by voltages -E, + E. One tapped voltage -E sin 0 at contact 451 forms the input voltage for the Airspeed booster and corresponds to gravity, which is called thrust Plane accelerates; two more tensions are shifted by 180 ° at the Contacts 452 and 453 are tapped and are used to 1. the resistance (6) of the servo system to feed and 2. to supply the input voltages for the a-amplifier, which the gravity components + E sin 0 and -E sin 0, which influence the lift, represent, depending on the 19 switch (Fig. 9).

When the 0 servo system is in the zero position shown, is the airspeed input voltage at contact 451 which is the gravity component perpendicular to the plane of the Y-Z axes of Fig. 2, equals zero and the voltage on the sliding contact 452 or 453, which is supplied by the roll servo switch 489 of FIG is and. is influenced by the roll servo system to keep the component of gravity perpendicular to represent the X-Y plane, a maximum. The linear resistor (4) is used to derive a tilt voltage to stabilize the coy servo system, the represents the longitudinal damping of the aircraft and partly the continuous vibration properties the exercise device, d. H. the damped wave of vertical oscillations.

The 0 servo motor also controls via a mechanical link 457 the angular position of the rotor coil 458 of a synchronous transmitter 459, which is electrically with a transmitter unit 460 (Fig. 9) is connected to the inclination element 493 of the position gyro 30 to control.

The tilt controlled cam 445 operates a number of Switching contacts depending on whether normal or inverted flight are reproduced, whereby the position shown corresponds to normal flight. This switch holds in combination with the 0 switch 485 the correct relationship of the trigonometric characters separate or combined inversions of the incline and roll position upright, as they occur in looping, split S, Immelmann and similar maneuvers. The cam pin 461 simultaneously actuates the switches 462 and 463, and the opposite one Cam pin 464 actuates switches 465 and 466 simultaneously. Switch 466 provides an a-tension which is influenced by the rolling and which in turn the output voltage of the buzzer amplifier 467 of the vertical airspeed circuit influenced to obtain a sliding correction depending on the angle of attack. To this The purpose is the switch 466 through a conductor 444 with a 180 ° phase shifter connected, the output voltage of which is fed to the input of the buzzer amplifier 467 will.

In addition to the a-sliding correction voltage, there are two other input voltages available for the vertical airspeed buzzer amplifier 467, which is ± v represent sin 0. This is the vertical component of airspeed. the Another input voltage is side slip voltage influenced by rolling β sin 0 (Fig. 9) which is a sliding correction voltage for β. The measuring device 54 for vertical airspeed can be formed by a voltmeter, which fed by the output voltage of the series connected amplifiers 467 and 469 will. The output of amplifier 467 is also normally connected to the input for the altitude servo system connected to the vertical airspeed for the altitude display to integrate.

Finally, there is a switch 468 that operates through the cam 445 is operated, which is normally the output circuit of the buzzer amplifier 467 of Fig. 18 for vertical air velocity grounds provided that there is a negative angle of inclination, and an intended position of the height servo cam 533 represents sea level or zero altitude. As the input voltage for the amplifier 469 for vertical airspeed is zero in this case, none can further downward movement of the altimeter can be made, and there can be no negative Display on the meter 54 for vertical air velocity occur to further Display altitude losses.

Servo system for displaying the rotation around the longitudinal axis (roll servo system), aileron control and attitude gyro In Fig. 9, the aileron servo system (0) is shown together with the pilot controlled aileron potentiometer and air velocity v signal amplifier. The roll system is primarily influenced by the roll torque of the aileron control and the torque of the motor, and also by lateral slipping and changes in sliding flight. The input voltages for the roll servo amplifier contain a negative speed feedback voltage E fb to make the roll change proportional to the roll moment, and two voltages that represent the aerodynamic coupling between roll and roll, one of the two voltages proportional to the roll change co and the other an the terminal 241 is proportional to ß, also a voltage of resistor (2), which represents the roll stabilization torque of the aircraft, also a voltage which is derived from the sliding contact 470 of the aileron potentiometer 471 and reflects the main roll torque when the aileron is operated, and finally a motor torque voltage on the Terminal 215.

The oar potentiometer 471 is fed from the v signal amplifier so that the oar torque voltage at contact 470 is a function of airspeed. This relationship is explained by the fact that the helix angle of the aircraft when rolling is equal to v, although the lift on the aileron is proportional. The helix angle, which can also be expressed as the ratio of the tangential speed of the wing tip during a roll to the forward airspeed v, is proportional to the roll speed of the aircraft as a result of a given roll moment is determined by a damping factor. It can therefore be shown that the taxiing speed for a given aileron deflection corresponds to the value v and not v2. By feeding the aileron potentiometer according to v, the damping factor can therefore be automatically included in the determined roll speed, so that the determination process is simplified. By changing the location of the ground contact 472 of the shunted trim potentiometer 473, the aileron input voltage can still be changed to mimic the trim.

The v signal voltage, which the aileron and trim potentiometers is obtained from the v-signal amplifier of FIG. This amplifier will in turn fed with a v signal that is amplified to the primary winding 474 of a transformer 475 which has two secondary windings. A secondary winding 476, which has a grounded center tap, is with potentiometers in the other Servo systems connected, and the other untapped secondary winding 477 feeds the aileron and trim potentiometers.

The potentiometers of the rolling system work as follows: The resistor (1) has an outline that corresponds to a sin-cos function and has two sliding contacts offset by 90 @. It is fed at terminals offset by 180 ° according to the shift change, and the voltages taken from the sliding contacts offset by 90 ° are fed as input voltages to the? P-signal and 0 servo amplifiers. One derived voltage + co, cos (P on contact 478 serves as the y) signal amplifier input voltage (Fig. 8), and the other on contact 479 is an input voltage - co, sin 0 for the 0 servo motor. The resistor is grounded at points 90 ° from the voltage clamps to represent the zero roll position for plane and inverted flight. The resistor (2) has a linear characteristic and is fed by a voltage E. It serves as a stabilization potentiometer to generate a voltage that corresponds to the characteristic damping and roll stabilization of the aircraft. Resistor (3) has a sin-cos contour and, like resistor (1), is fed by a ß-signal voltage. The derived voltage ß sin (P forms the input voltage for the summv amplifier 467 of the circle for the vertical airspeed in order to introduce the vertical component of the lateral slip speed into this display system. The resistor (4) with a sincos outline has two sliding contacts and is shown in fed in a similar manner according to a, the derived voltage at the one contact 480 in the circle being used for the vertical airspeed, possibly after polarity reversal, depending on the position of the contact 466 of the O-switch 445 (FIG. 8) Another voltage tapped at contact 481, which is offset by 180 °, is used for gliding correction, if necessary also after reversing the polarity at contact 466 of the 0 switch both in the circle for the vertical airspeed and in the flap circle in FIG. 6. The resistance ( 5) with sin-cos outline has three sliding contacts and is similarly depending on the change in inclination g eats. The derived voltage cov cos 0 at contact 482 forms an input voltage for the 0 servo amplifier, and the voltage + co, y sin 0 at contact 483 offset by 900 is dependent on the input of the course direction signal amplifier and the input of the (o, servo amplifier supplied by the 0 switch 484, which is actuated by a cam disk 485 which is mechanically connected to the 0 servo shaft. A third derived voltage - co, sin 0 at contact 486, which is 180 ° with respect to contact 483 of the resistor ( 5) is offset, is used as input voltage for the c,.), Servo system, depending on the position of switch 484, ie when the roll angle in inverted flight exceeds 90 ° and is in the range between 90 and 270 ° . The resistor (6) with a sin-cos outline has two sliding contacts that are 270 ° apart and is fed with a 0 voltage (Fig. 8). The derived voltage -1- E cos O sin 0 at contact 487 forms a gravity component of the input voltage for the ß-signal amplifier (Fig. 7), and the derived voltage E cos O cos 0 at contact 488 is a gravity load factor of the input voltage at a- Amplifier. Angle of Flight Control Fig. 10 shows the apparatus used to make the aviator training device respond to angles of flight of the mimicked flight. The flight angle according to FIG. 3 is measured from the horizontal and is determined by the path of the aircraft and not by its attitude or angle of inclination. The flight angle therefore represents the difference between the angle of inclination and the angle of attack or Y = O - a. The gravity or weight vector is denoted by sin (O - a) .

The a and 0 servo systems can be operated automatically to determine the flight angle as shown in FIG. The a-servomotor adjusts the cosine resistance 701 via a direct mechanical connection 700 in accordance with the angle of attack. This resistor is fed by voltages -y- E and -E in phase opposition at two points offset by 180 °. The 0 servo motor independently adjusts the potentiometer contacts 702, 703 and 703 'via a mechanical connection 704, the contacts 703 and 703' being offset from one another by 180 ° and shifted by 90 ° in relation to the contact 702. The voltage tapped at contact 702, which represents sin (O - a) , is connected to conductor 706 via a slip ring connection 705 and is connected to the input of the airspeed amplifier. In the same way, the antiphase gravity voltage components -f- cos (O - a) and - cos (0 - a) are connected to the contacts 703 'and 703 via slip rings 707' and 707 to the conductors 708 'and 708 in order to increase the resistance ( 6) of the 0 servo potentiometer, which controls the heavy. resolves force-tension components in dependence on rolling. The derived voltage + E sin 0 cos (0 - a) at contact 487 of the 0 resistor (6) is an input voltage to the amplifier (Fig. 7) for side slip to determine the gravity vector of slip and represents the component of gravity along the y-axis for flight angle (O - a). The other component - E cos 0 cos (0 - a) of the contact 487 'shifted by 90 ° is at the input 236 of the angle of attack servo motor of FIG. 6 in order to add the gravity load factor determine.

Claims (7)

PATENT CLAIMS: 1. Flight training device for simulating the flight angle with a simulated flight control system and associated simulated flight monitoring instruments as well as a flight calculation system which responds to the controls and which the controls contains associated tools for deriving electrical signals that are used for Representation of flight control factors, e.g. B. the angular speeds of the rotary movement around the aircraft axles, as well as a number of interconnected and electrical systems responsive to the signals, for example for Representation of the angle of attack, as well as electrical integration systems for representation provides for the aircraft position in relation to the axes of movement, characterized in that that facilities are provided, both on the angle of attack system as well address the lateral axis system (Fig. 10) to generate a signal indicative of the flight angle of the flight to be simulated. 2. Device according to claim 1, characterized in that da.ß with the transverse axis system and with the longitudinal axis system full circular rotary movements carried out and by flight control instruments (gyrocompass 30 in Fig. 9, more magnetic Compass 59 and gyrocompass 27 in Fig. 7) are controlled and that switching means (445 in Fig. 8 and 485 in Fig. 9) each of which is provided by one of the two Systems are controlled and with the signals of the systems (Fig. 6 and 9) for the Attitude are connected so that all maneuvers to be carried out around the transverse and longitudinal axis be imitated instrumentally and realistically. 3. Device according to claim 1, characterized by signal generating devices (Fig. 7) in connection with the course direction system to imitate acceleration and northing errors, which are controlled by a simulated compass device (59). 4. Apparatus according to claim 1, characterized by signal generating devices (383, 384, 385 in Fig. 6) to imitate the wing flaps and the air resistance, the torque and the lift by controlling the angle system (Fig. 6). 5. Device according to claim 1, characterized by signal-generating aids (370, 362. in Fig. 6) for imitation the formation of ice on the wings, which serve the angle of attack system as well to influence other electrical systems representing airspeed. 6. Apparatus according to claim 1, characterized in that means (381, 381 in Fig. 6) to imitate stormy air movements are available, the optional can be used to generate arbitrary signals to generate additional electrical Systems of the flight calculation system to influence which the rotational movement around the Transverse axis (Fig. 4) and the angular speed of the rotary movement around the vertical axis (Fig. 5) to imitate. 7. Apparatus according to claim 6, characterized in that the angle of attack system (Fig. 6) Provides switching devices (360, 361) with which the stormy air movements Signals representing also include critical values of the angle of attack for stalling to be able to imitate the airplane. References Considered: U.S. Patents No. 2,494,508, 2,366,603, 2,510,174, 2,463,602, 2,522,434, 2,475 31.4, 2,529,468, 2431749.