DED0018685MA - - Google Patents

Description

FEDERAL REPUBLIC OF GERMANY

Registration date: September 17, 1954. Advertised on February 2, 1956

GERMAN PATENT OFFICE

The invention relates to flight training devices for trainee pilots and relates in particular to on the ground located devices of the type mentioned for the calculation of flight conditions and for the simulation of forces on the trainee pilot, who occur in the real flight, when a turn around the three Axes of the aircraft during flight maneuvers, such as pitch, bank and turning, takes place.

In actual flight, the combined effects of gravity, acceleration and the centrifugal force acting on the student's body is represented as a resultant force acting either in a direction perpendicular to the pilot's seat, e.g. B. at a correctly executed turn in which the trainee pilot feels that he is sitting normally upright, or which works at an angle to the seat, as in "sliding sideways" or "sliding flight" maneuvers, the trainee pilot feeling that he is tipping one way or the other. The forces that The trainee pilots are essentially influenced by gravity and centrifugal force as a result of yawing and the centrifugal force due to the pitch. It is therefore not sufficient to use the mentioned maneuvers simply by tilting the chassis in the general direction of stick movement to imitate, as such tipping can be very misleading. So is z. B. at a flat

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Curve to the left at a speed that causes sliding greater than the centrifugal force the force of gravity, and the resulting force acting on the trainee pilot, creates the impression that

• 5 his seat tilts to the right instead of to the left. Even in the case of a precisely flown curve, a Inclination of the seat give the student pilot a completely wrong impression, as with such Curve the sitting reaction would be normal. Similar

ίο incorrect impressions can also occur with a longitudinal incline of the aircraft are generated, z. B. the acceleration during a dive actually for the student pilot can give the feeling that his seat is tilting backwards instead of forwards, while the opposite is true for slowing down during a shallow climb.

A main object of the invention is therefore to provide a device for calculating flight conditions and to mimic the effects commonly referred to as the "sitting sensation" to create and those in real flight maneuvers due to the combined force of gravity, acceleration and centrifugal force act on the student pilot.

Another object of the invention is the creation of improved calculation and control means, which respond to the operation of simulated flight controls to determine the position of the seat of the trainee pilot in accordance with a resulting force vector determined by the mimicked flight conditions is derived.

Another and more specific object of the invention is the creation of improved electrical computing and control facilities based on the Address operation of mock flight controls in order to obtain control voltages which Represent force vectors that act on the student pilot, using these tensions for this purpose will be able to control the position of the student pilot's seat in the training device.

The invention is explained in more detail in the following description in conjunction with the drawings.

Fig. Ι is a schematic representation of an electrical Calculation and control system in a flight training device to simulate the effects of forces on the student pilot;

Fig. 2 shows schematically the flight computing system of Fig. Ι in greater detail, and

Figs. 3 and 4 are vector diagrams showing the resulting Show force vectors.

In fig. ι is a cabin S of the exercise device with the position of the pilot schematically in connection with the associated calculation and control devices of the invention- shown. As such, the exercise device may be of and include any suitable type a seat 1 for the trainee pilot and replicas of the aircraft controls including a throttle 2, a stick or a control column 3 and the rudder pedals 4. The controls concerned the throttle and the aileron, elevator and rudder are according to a preferred embodiment with Means for dissipating voltages operationally connected, e.g. B. with potentiometers 15, i6, 17 and 18, which have movable contacts 15 ', 16', 17 'and 18'. The potentiometer contacts can be used to derive voltages depending on the aircraft control actuation concerned can be adjusted in a manner known per se. It should be noted that the invention also applies to others Kinds of exercise equipment is applicable, such as B. for those that are built for rotation in azimuth, without the particular type of operating medium that matters.

The cabin S of the trainee pilot is suitably gimbaled at 9 so that a longitudinal inclination around the transverse axis and a transverse or rolling inclination can be carried out around the longitudinal axis, the axes being designated as Y and X axes (Fig. 3 ). The gimbal 9 can, for. B. a yoke 10 which is connected to a rotatable shaft 11 which can be pivoted by means described below to produce a transverse inclination movement, while a transverse axis 12, to which the car S is directly attached, rotatable in the yoke is mounted and can be operated by additional means described below for generating a longitudinal inclination.

The electrical calculation and control systems for the lateral and longitudinal inclination of the student's cabin to imitate the above-mentioned “sitting feeling” during flight maneuvers is shown schematically in FIG indicated, while the calculation and integration system in Fig. 2 is shown as a block diagram.

In Fig. Ι a reference AC voltage source E is provided for feeding the entire system, and the various derived and control voltages are taken from this source, the positive and negative signs represent the instantaneous value of the polarity with respect to the reference current source. So z. B. the various aircraft control potentiometers fed by voltages which represent certain functions of the airspeed ν , which are obtained in the calculation system. The throttle potentiometer 15 is fed at its lower terminal with a voltage that is the reciprocal of the Eigengeschwindig-

speed - represents, and is at its upper terminal

grounded so that the removed from the sliding contact 15 ' Voltage when adjusting the throttle

represents the thrust T according to the relation T = ~. The elevator and rudder potentiometers 17 and 18 are each fed with a positive voltage + u ² at their upper terminals and with a negative voltage - v ² at their lower terminals, which correspond to the square of the vehicle's own speed. The aileron potentiometer 16 is fed with a voltage ν that is linear to the airspeed. Each potentiometer is provided with a grounded center tap in order to represent positive and negative turning moments about the usual aircraft axes according to FIG. 3 with reference to normal plane flight. The various derived torque voltages of the control potentiometers 15, 16, 17, and 18 are fed through lines 19, 20, 21 and 22 to the calculation system 23, from which

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Control voltages for operating the longitudinal and lateral inclination device of the student cabin S derived will.

The flight computing and integrating device of FIG. 2 will now be described. The device contains after In this illustration there are essentially seven servo units, each of which has a flight condition, e.g. B. Airspeed, angle of attack, change in pitch, etc., represented by an electromechanical Network are interconnected so that they work according to certain flight principles in order to simultaneously and continuously to calculate the relevant flight conditions. The different with each other in Connected circles of the servo units shown are so far for a better understanding simplified so that the invention emerges clearly therefrom.

When operating the system shown, in

primarily derived from the operation of the simulated aircraft controls described above voltages, which are proportional to the various speeds and forces that cause a movement or acceleration with respect to the three reference axes in accordance with the basic aerodynamic conditions. The three reference axes of FIGS. 3 and 4 are firstly the longitudinal or% axis of the aircraft, secondly a y-axis in the direction of the aircraft wing perpendicular to the longitudinal axis and thirdly an axis z which is perpendicular to the other two axes, all of which Axes go through the center of gravity of the aircraft.

The translational and rotational movement with respect to these axes and with respect to fixed ones Axes that are perpendicular and parallel to the horizon are determined by the servo systems. In one of these systems is used to determine the forces that determine the airspeed (air speed) result, in another servo system, moments are calculated to increase the yaw rate and in a third, moments are calculated to produce changes in inclination. Two auxiliary servomotors are provided to show the angle of attack and the side slip, where the pitch servo motor speeds around the y-axis integrated to the aerodynamic Sizes of lift (weight plus centrifugal force), drag and To calculate the tilting moment, while the servo motor for lateral slipping the angle between the plane of symmetry of the aircraft and the flight path. Integrate the other two servomotors Angular movements as a function of control voltages derived from the three above-mentioned Servo systems are generated to reproduce the flight conditions of taxiing and pitch.

According to the known principles of aerodynamics

The airspeed ν is a function of the engine thrust T, which is always positive (with the exception of the propeller resistance when idling in flight with less than about 1200 revolutions per minute), as well as the force of gravity G, which can be either positive or negative, depending on the aircraft in

- Downward or upward flight, and the air resistance, which is of course negative. The air resistance generally has two components, namely a resistance component with a constant coefficient, which changes with the square of the airspeed v ² , and a further resistance _{component, which depends on a variable coefficient Cd (K} ), which changes with the angle of attack (α) changes, that is, the angle between the tendon of the wings and the airflow.

With reference to FIG. 2, it is assumed that a number of alternating voltages, which represent different values of thrust, gravity and air resistance 75 · corresponding to the instantaneous values of polarity and magnitude of the voltages concerned, are fed separately to a summing amplifier, which is indicated schematically at 25 and is part of a servo system called "air speed". Such amplifiers, which algebraically add a number of different alternating voltages of variable magnitude and polarity, are known per se, and a detailed circuit therefore does not need to be given. The output of the amplifier 25 controls an automatically adjusting transmission circuit which contains a two-phase motor 26, the control phase 27 of which is fed by the output voltage of the amplifier in the manner shown, while the other phase 28 is fed with a constant reference AC voltage + ^ ₁ . The mode of operation of such a motor is known per se; it rotates in one direction if the control and reference voltages in the relevant phases have the same instantaneous polarity values, and in the opposite direction if the instantaneous polarity value of the control voltage is opposite to that of the reference voltage, the speed of rotation in both cases depends on the size of the control voltage. The motor drives a two-phase feedback generator 29, one phase 30 of which is excited by a reference AC voltage - \ -: e ₂ , while the other phase winding 31 generates a feedback voltage E _n corresponding to the motor speed for a speed control described below. The feedback voltage E _n represents - = -, i.e. the acceleration

German

supply and is on the one hand the entrance of the Ver-.iio ' amplifier 25 and on the other hand via a line 24 to the device of FIG. 1 for a described below Purpose supplied. The motor also serves to have a reduction gear 32 and appropriate mechanical connecting members, which are indicated by the dashed line 33, the contacts a potentiometer system and the pointer of an airspeed indicator 34 to operate.

The individual resistance elements 35, 36 and 37 of the potentiometer are on in a known manner Forms wound and can be circular or band-shaped in practice, but are in the schematic representation indicated as planar resistors.

It can be seen that operation of the motor 26 in either direction is not just one

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Speaking change in the display of the own speed, but also a movement of the sliding contacts 35 ^a , 36 ^s and 37 "causes them to move on the relevant potentiometer elements in angular positions in which potentiometer voltages are derived, ie tapped or selected by the relevant Each potentiometer is shaped or has such a contour that the value of the tapped voltage at the potentiometer contact has a certain relationship to the linear movement of the sliding contact, which depends on the special function of the potentiometer, with one at the terminals of the potentiometer Voltage, whose instantaneous values of polarity and magnitude also depend on the function of the potentiometer linear course, while the Potentiometer 36 is contoured to represent the relationship χ = y ² , where χ is the linear movement of the contact and y is the derived potentiometer voltage, here

.25 case corresponds to the square of the airspeed.

More specifically, it is the outline or the change in width and therefore the resistance distribution of the various potentiometers used to derive voltages, the characteristics of the aircraft emulate, proportional to the derived value of the function of the property in question with reference to the variable represented by the setting of the potentiometer. Let it be B.

assume that the function is linear when e.g. B. a tapped voltage is directly proportional is the distance that the potentiometer contact actuated by the servomotor has from a zero position. the The slope of the function curve is then the constant

4.0 Ratio of the tapped voltage to the increase the independent variable represented by the contact path from the zero position. The derivation of this relationship is the same for all Kontakteinsteliungen, so that the width of the

1-5 outline is uniform and the resistor is rectangular in shape. Now if the function changes according to a quadratic law, e.g. B. χ = y ² , then the derivative of this equation f ( _X ) = 2 y determines the width of the potentiometer. The potentiometer therefore has a straight inclined edge so that it has a wedge shape.

In another case, if a cosine function is involved, the derivative or slope is the

Cosine curve equal to the expression - = —sin Θ, ^ ^

where θ is the angle measured in radians. The outline of the potentiometer resistance for corresponding values of θ is therefore sinusoidal, the negative values being taken into account by appropriate selection of the polarity applied to the potentiometer. Conversely, when a sine function is involved, the potentiometer resistors have a cosine outline for corresponding values of θ.

The potentiometer 36 of FIG. 2 is fed with a negative voltage - E from the reference voltage source at its upper terminal, which represents the maximum airspeed, and is grounded at its lower end, so that the voltage derived at the sliding contact 36 * has the value - v ² and therefore also represents the constant coefficient of air resistance mentioned above. The contact 36 "is connected by a line 38 to the input of the amplifier 25. The voltage of this contact is therefore used as an input voltage for the airspeed summing amplifier, which tends to counteract the positive voltage for the thrust T , the The arrangement is such that when all amplifier input voltages are in equilibrium, that is, during a period in which there is no change in airspeed, the output voltage of the amplifier is zero and the motor 26 is not energized Unbalancing in either a positive or negative direction, such as changing the throttle setting when the thrust and resistive voltages are unequal, causes the motor 26 to be actuated in a corresponding direction so that the potentiometer contacts are in a new equilibrium position be moved, causing the new from attempt to restore the equilibrium of the motor input voltages through the voltages conducted.

In order to derive a voltage which is proportional to the airspeed ν , the linear potentitiometer 35 is fed by a voltage - E , and the sliding contact 35 ° is adjusted according to the size of the airspeed. This derived voltage is used in another part of the system described below.

The shear stress is derived from the setting of the motor throttle potentiometer 15 (Figs. 1 and 2), the contact 15 'of which is set directly by the trainee pilot in order to mimic the throttle control. The potentiometer is fed by a voltage (line 39), which is ^{tapped from the contact 37 11 of} the potentiometer 37, which is connected at its lower end to a voltage - \ - E , while the upper terminal is grounded via a resistor R and also with the sliding contact 37 "is directly connected to derive a voltage, which is proportional to the reciprocal of the

Airspeed - is so that the relationship

T =

H ■ P

that is simply adhered to the basic equation

lake

lake

is equivalent to.

It can be seen that the thrust input voltage generally corresponds to the output of the motor, which is determined by the throttle setting and the airspeed.

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The input voltage C _{D of} the resistance coefficient changes, as indicated above, with the angle of attack a. Another servo system called the angle of attack is therefore provided to derive a set of voltages corresponding to various factors which depend on the angle of attack. A two-phase motor 40 (similar to motor 26) of the pitch system is energized by the output voltage of a summing amplifier 41 in the manner described above to drive a feedback generator 42 and the contacts 43 ", 44" and 45 "of the potentiometers 43, 44 and 45 These potentiometers are used to calculate the air resistance, the rate of change of α due to a change in lift and the pitch moment. The input voltages of the α amplifier 41 contain voltages that are a function of the force of gravity G, the lift Gl and the pitch moment Cm . These input voltages are briefly explained below: The air resistance can be expressed as a function of the angle of attack

D = C

QV ² p

There)

where D is the air resistance in kg, ρ is the air density, Ci> (, a) is the drag coefficient and s is the projected wing area. The air resistance is therefore a function of v ² , ie the square of the airspeed. To illustrate this relationship, the potentiometer 43 has a corresponding contour and is fed at its opposite terminals with a voltage - v ² , which is tapped from the potentiometer 36 of the airspeed system. The center of the potentiometer 43 is grounded via a resistor at an angle of incidence for which the drag coefficient Cd ^ is zero, and the sliding contact 43 "is connected via a line 46 to the ^{airspeed amplifier 25.} The voltage derived at the contact 43 ffl can therefore, because it changes with a change in the angle of attack, in general, according to the above relationship, can be used as an input voltage D (a) of the airspeed of the amplifier. the gravity input G, which depends on the longitudinal inclination of the aircraft, requires an additional servo system that will be described below.

The input voltages to the angle of attack (α) amplifier 41 are discussed below. The angle of attack servo system is an integrating servo system that derives the instantaneous value of the angle of attack from the temporal integration of velocities which, like ei ,, influence the angle of attack. The gravity factor that ; As mentioned above, is influenced by the climb or dive attitude, can be broken down into two components, which are fed to the angle of attack and airspeed amplifiers 41 and 25, respectively. In practice, these gravitational tensions are go ° components, ie the airspeed component acts along the flight path and the angle of attack component is perpendicular to it. In the present example, the v and α gravity components are derived from potentiometers 47 and 48 of the "pitch" servo system Θ , the pitch amplifier 50 being connected to drive the two-phase motor 51 etc. of a "pitch change system" (rate of pitch), which is described below. The tilt potentiometer 47 has a corresponding outline (cosine-shaped in the present case) and is grounded at points offset by 180 ° to represent both normal flight and plane return flight, and the potentiometer is provided with voltages - E and - \ - E , which represent the gravity values of the upward (negative) and downward flight (positive). The derived voltage at contact 47 ″ represents the gravity component G sin Θ , which at low angles of attack represents the effect of the aircraft weight with increasing or decreasing thrust and, as a result, airspeed fed to the ^ amplifier 25. The input voltage of the gravity component

for the α amplifier 41 can be considered as a relationship

be represented, where this expression has the dimension of a speed, like this of the integrating servo system of the angle of attack is required. To dissipate this tension, this will Inclination potentiometer 48, which has a cosine shape like the potentiometer 47, with potentials fed, which are derived from the airspeed potentiometer 37 and, + - and there-

put. The derived voltage at the contact 48 ″ of the potentiometer 48 (which is offset by go ° from the contact 47 ^a ) represents the gravity component, the counter being the component

which is to be generated by the lift derived from the angle of attack and which is generated by a Line 53 is fed to the α amplifier 41.

With reference to the angle of attack system, the lift L in kg can be given by the relationship

i- ¹ ~ - tsTj {a \

where Cz ( _a ) is the coefficient of lift. The lift is therefore also a function of the square of the airspeed and depends on the type of aircraft being simulated. The potentiometer 44 of the α system for determining the lift has a suitable shape to simulate the coefficient C1 («> of the aircraft in question, and is grounded via a resistor in the middle part at a value of the angle of attack at which the lift coefficient is equal to zero If the potentiometer were fed by a voltage v ² , the derived emf would represent the acceleration of the lift. If this acceleration is divided by the airspeed, then the factor has the dimension of a speed as determined by the integrating servo system of the angle of attack system for the Determination of the eye

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view value of the angle of attack is required. The potentiometer 44 is fed at its upper and lower terminals by voltages - -v and + ν derived from the airspeed potentiometer 35. The positive instantaneous value of ν can be obtained in a suitable manner by means of a 180 ° phase shifter. The contact 44 ″ of the potentiometer 44 therefore derives a voltage dependent on the lift, which is fed as an input voltage to the α-amplifier 41. The angle of attack system also has an input voltage which corresponds to the angular amount of pitch, and this input voltage is derived from the integrating servo system , which is designated with "change of inclination". co _y , namely on a potentiometer 59, whose sliding contact 59 ^s decreases a voltage that is directly proportional to the position on the ca ,, - shaft.

The input voltages of the inclination change system contain a so-called inclination moment voltage, which is derived from the potentiometer 45 of the angle of attack system. This pitch moment corresponds to the relationship

and is a function of the square of the airspeed. The potentiometer 45 is grounded through a resistor in its central part for an angle of attack at which the tilting moment is equal to zero, and is fed by voltages - v ² and - \ - v ² as in the case of the height control potentiometer and has a contour such that the inclination moment voltage on the sliding contact 45® changes according to the desired characteristics of the aircraft in question. This voltage is fed to slope change amplifier 55 through line 54. The other main _{input voltage M 3 of} the amplifier 55 represents the moment of inclination in mkg, which is generated by the altitude control operated by the trainee pilot and causes a longitudinal inclination, and is derived from the altitude control potentiometer 16, which in turn is expressed as the square of the airspeed through voltages - \ - v ² and - v ^{2 is} excited. The middle part of the potentiometer is grounded to represent substantially level flight or zero pitch moment. The sliding contact 16 ′ of the height control potentiometer therefore picks up a voltage which can be represented as inclination moment M ₁ in mkg and which is fed to the inclination change amplifier 55. It should be noted that in the circles just described, a positive (+) signal indicates the vehicle's own speed.

speed, the change in the angle of attack, the change in pitch 'and the pitch in the normal positive direction is increased, as can be seen from FIG.

As with the previously described servo systems, the output voltage of the amplifier 55 feeds a two-phase motor 56 for driving a feedback generator 57 and actuates the contacts 58® and 59 ^{a of} the potentiometers 58 and 59 through a suitable reduction gear 60. The linear potentiometer 58 is used to control a To derive input voltage, which represents the centrifugal force for the ß- amplifier 100 of Fig. 1, and is therefore grounded in the middle part and is fed according to the own speed by voltages - \ - v and - ν , so that the derived voltage as _y v on the sliding contact 58 ^{s is} determined by the factors ν and ω _ν and is fed through the line 61 to the device shown in Fig. ι. The linear potentiometer 59 for supplying an input voltage to the inclination (ö) integration system, which was mentioned above, is supplied with voltages - \ - E and - E , so that the derived voltage at the sliding contact 59 ⁰ , which is carried by a line 62 is fed to the Q amplifier, is proportional to the change in inclination, the value 'integrated over time representing the inclination angle θ of the aircraft. This integration is carried out with the aid of the pitch motor 51 and the feedback generator 63, the aforementioned potentiometers 47 and 48 ^{supplying voltages to the sliding contacts 47 01} and 48 ″ which not only reflect the two aforementioned gravity components, but also the instantaneous value of the pitch angle a flight condition gyro can be operated directly by the inclination motor 51 if necessary.

It should also be noted that the changes in the various angular velocities, forces and moments, such as B. gravity, buoyancy, centrifugal force, thrust, air resistance, pitch moment and the like, by changing the sliding contact position of the relevant potentiometer can be achieved with changes in the potentiometer supply voltages, while the relative magnitude or the effect of the mentioned changes, forces and moments due to the value of the input resistance the different amplifiers is determined. As a specific example is the relative size of the lift on the values of the air density ρ and the

constant surface factor - dependent. In the present example, ρ is also regarded as a constant, and these expressions therefore determine the resistance value of the input circuit C1 at the amplifier 41. A decrease in the value of the resistance increases the relative size of the constant. The use of feedback generators is particularly important for change control, with the pitch integration system serving as an important example. If the motor 51 were to do the pitch integrating process alone, the inherent inertia of the drive mechanism would introduce such a large error that the system would not be useful for practical purposes. With the feedback generator, however, which is switched on in the illustrated manner, the generated feedback savings E _{n forms} an input voltage for the pitch amplifier and has such a phase position with respect to the summed or resultant input signal that it

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against, d. H. in the manner of a negative feedback, acts. With a large gain in the control amplifier the speed of the motor has a linear dependence according to principles known per se the speed on the size of the input signal, d. H. on the slope change voltage, without lagging or overshooting, allowing both large and small incline changes with can be integrated with the same accuracy. It can be seen that when the main input signal is reversed and operating the motor and generator in reverse direction of the phase of the generated feedback voltage is also reversed to counteract the input signal as before.

If, according to Fig. 2, the throttle potentiometer contact 15 'downwards, for. B. is moved into the open throttle position, then the derived input thrust voltage T of the amplifier 25 increases the disturbance of the equilibrium on the airspeed servo system

ao. and causes the servo motor 26 to run in a direction in which it ^{moves the potentiometer contacts 35 ", 36" and 37 a} upwards in the sense of an increase in airspeed, so that the following processes take place in the airspeed potentiometer system: 1. the derived airspeed voltage ν increases; 2. the derived w ² stress increases with the square of the airspeed; 3. the derived voltage,

which is the reciprocal of the airspeed -

represents becomes smaller, and 4. the airspeed indicator 34 shows a higher airspeed. However, the airspeed cannot increase infinitely because the constant drag coefficient increases with v ² · , as does the air drag CD ( _K ). At the same time, the thrust, which changes with the reciprocal of the airspeed, decreases when the new equilibrium is reached. If the airspeed increases, the angle of attack system gets out of balance, since the input voltages derived from the potentiometer 59 of the inclination change system and from the potentiometer 44 of the angle of attack system, both of which are indirectly or directly dependent on the airspeed ν , are now increasing. The pitch system gravity input voltage is also changed as described below. The α servo motor 40 begins to run in one direction so that it seeks a new equilibrium position, and moves the potentiometer contacts 43 ″, 44 ^a and 45 ^a downwards in the sense of a decreasing angle of attack the derived voltages of the three α-potentiometers 43, 44 and 45 are used as follows:

i. The derived air resistance voltage (negative) of potentiometer 43 is used as input voltage D (α) for the airspeed booster and increases in size so that it counteracts the increasing shear stress (positive) resulting from the higher throttle setting.

2. Since the wing lift of an aircraft must balance the centrifugal force and weight components that act perpendicular to the wing, the derived lift velocity voltage Ci _{1 of} the potentiometer 44 must be both the gravitational

factor and the incline speed

Factor (o _y compensate. Assuming the aircraft was originally in level flight, then the pitch angle speed is zero, and the increasing airspeed tends to reduce the angle of attack, which is therefore negative. This effort is counteracted by changing the pitch moment which influences the position of the contact 59® through the amplifier 55, as described below.

3. The derived pitch torque voltage from potentiometer 45 applied to input Cm of pitch change amplifier 55 becomes more positive as the angle of attack decreases and therefore causes the pitch change servo system to become unbalanced, moving contacts 58 ° and 59 "up. The contact 59 "sets a pitch angular velocity voltage ω y for the amplifier 41 which tends to restore balance to the α servomotor. The upward movement of the contact 59 ° also generates an increased input voltage for the amplifier 50 of the tilt integrating servo system. All four servo systems therefore work together in a combined calculation and integration process which is necessary to determine the new airspeed and the new pitch state to determine.

When the pitch system is moved to a more positive pitch position, ie an increase, the derived voltages at contacts 47 ^a and 48 ″ of potentiometers 47 and 48 represent the gravity input component for the υ and α amplifiers, which vary in size , with the D component increasing and the α component decreasing in the present example It can be seen that if the tip of the aircraft were directed to the zenith, the gravity component in the direction of the aircraft's motion then - G and the gravity component perpendicular to the Wings, ie the a-servo component, would be equal to zero.

The negative gravity component (- G sin Q) of the airspeed servo system tends to reduce the maximum speed that the aircraft can reach at a higher throttle setting. At the same time, the required wing lift due to the decrease in the input voltage G cos θ to the α amplifier 41 is reduced. Hereby is a further decrease in the angle of attack and a further reduction of the negative pitching moment of voltage Cm on the change in inclination of amplifier 55 instead, which causes in turn a greater upward movement of the contacts 58® and 59 ^a, whereby the effect of increases in the pitch and Anstellwinkelservomotoren until finally these servomotors have produced too great a change in the weight component for balance and are overriding it. Consequently

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there is a drop in airspeed. This in turn causes a decrease in the lift voltage Cl at the α-amplifier 41, so that the angle of attack is increased and a larger negative inclination torque voltage is generated at the potentiometer 45 for the ω "amplifier 55. The "" contacts 58 ° and 59Λ will now move down to control the pitch integrator servo so that pitch will decrease until it eventually becomes negative. The G-sin-0 component of the airspeed servomotor has now become positive, so that the thrust is supported and the airspeed is greater, and this process is repeated, finally with damping a final equilibrium position is reached which corresponds to the new throttle setting. The tilt servo can be used for this. are to operate the tilt moment of a state gyro. In the way described above, the real damped wave-shaped path of the vertical oscillation or the "fugoids" of an aircraft is reproduced exactly, so that the simulation is very realistic. The extent of the damping of the undulating path depends on the choice of the constants of the circuits including the percentage of the speed feedback, the gear ratios, the relative input quantities and the position of the potentiometer's center taps. The excessive flight condition is also determined by these factors.

In the above discussion it was assumed that only the throttle setting was changed and that the altitude control remained in the normal plane attitude or in the neutral position. When the elevator control is adjusted, a derived voltage corresponding to the turning moment is used to control a force integrating servo system, ie pitch change servo system a> _y , from which a voltage corresponding to the pitch angular velocity is derived. This pitch rate voltage eo provides an input voltage to the pitch servo system for deriving a pitch torque input voltage which is in the opposite direction but has the same magnitude as the first or altitude torque voltage. This same integrated moment voltage or inclination change voltage co _v controls the derivation of another change voltage, which represents the effect of the lift on the α change and whose polarity is opposite

.50 and which increases to offset the effect of the original a-change voltage. A damping quantity for the change in inclination, namely - co _v ■ ν , is determined by the. Contact 94 ^s of potentiometer 94 of the ("" servo system tapped. This

.55 generally illustrates how an equilibrium between the change in inclination and the angle of attack is achieved.

During the pitch control described above, the α system seeks a state of equilibrium which depends on the input voltages which determine the slope change of the slope change system and the gravity component of the tilt system on the one hand and the lift velocity stress represent the changed angle of attack on the other hand, the resultant of these input voltages the α-motor 40 in the positive or negative direction actuated and comes into equilibrium when the inclination change and the incline system .be stabilized.

From the above discussion it can be seen that two integrations are involved in order to either determine the angle of attack of the aircraft or the pitch position of the aircraft; the first Integration refers to the acceleration (force) on the speed and the second on the speed over the angle.

The above description of the actuation of the servo systems deals mainly with a "vertical system", ie a movement around the y-axis of the aircraft. Movement around the z-axis (yaw) and around the * -axis (roll) will now be discussed. A yaw change servo system is fed with a number of input voltages which contain an alternating voltage which represents the turning moment M _{t of} the rudder potentiometer 18. This potentiometer (Figs. 1 and 2) is grounded in its middle section to represent straight flight and is fed with voltages at the upper and lower terminals which ^{correspond to the squares + u 2} and - v ^{2 of} the airspeed. The derived voltage at the rudder-controlled contact 18 'corresponds to the yaw effect or turning moment generated by the right or left rudder and represents the input M _t for the co _z amplifier 65. The ftv system changes the rudder voltage out of equilibrium, provided that there is no compensating change in the other input voltages, so that the output of the amplifier 65 moves the two-phase motor 66 to actuate the potentiometer contacts 67 "and 68 ^a in a new equilibrium position on the potentiometers 67 and 68 in question, ie for right The potentiometer 68 has a grounded center tap and is fed at its opposite ends by voltages which represent - {- v and - ν , so that the voltage ω _ζ · ν, ^{derived at the contact 68 a,} ie the centrifugal force, - This voltage is fed through a line 64 to the device of Fig. 1. The Po tentiometer 76 derives a voltage at contact 76 ″ which corresponds to the action of yaw on the roll, and supplies an input voltage on line 76 * for this purpose, which is fed to roll servo amplifier 77. As with the servo systems previously described, the motor drives a feedback generator 69 and operates the potentiometer contacts via a gear box 69 ° and mechanical links bg ^h .

The remaining input voltages for the amplifier 65 contain a voltage which represents a damping factor, ie the reaction force ßv ^{2 of} the chassis on the yaw, which is derived from a / 3 servo system for the slip angle, and the feedback voltage E _n from the generator 69 for Ensuring the operation of the engine 66 with

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the correct speed for the integration of the rudder and side force voltages over the Yaw angular velocity.

The / 3 system contains an integrating servo amplifier 70, the output voltages of which are fed to the two-phase motor 71 and the speed-controlling feedback generator 72 in order to adjust the sliding contact 74 ° of the potentiometer 74 via the gear box 73 and mechanical connections. The potentiometer 74 is grounded in the middle and is fed with voltages - v ² and + u ² at its upper and lower terminals, so that when β increases or decreases, the derived voltage, which represents the reaction force of the chassis or the damping factor, changes accordingly. This voltage is fed to the ω _ζ amplifier 65 through a line 75.

The input voltages for the β amplifier 70 contain the feedback voltage E _n , a voltage

ao voltage, which represents a factor of gravity and is taken from the roll (99) servo system, which is described below, and a voltage which represents co _z and is derived from the potentiometer 67 of the ω _z system. This potentiometer is grounded in its middle section and is fed at its upper and lower ends with voltages - E and - \ - E , so that the voltage at the contact 67®, which is fed through a line 78 to the β amplifier, the Value co _z .

The gravitational acceleration stress, which acts in the usual -y plane of the aircraft (Fig. 4), namely in the direction of slipping over time, is broken down as G sin φ cos 0, where ψ is the roll angle and θ is the pitch angle. Since the lateral slip angle /? is derived by a time integration of the angular velocity, the gravitational acceleration factor G sin φ cos θ is divided by the airspeed in order to obtain a factor which has the dimension of the change.

To derive the factor, the feed

Contact voltages of the pitch servo potentiometer 48 described above determine the cosine resistance 90 of the roll servo system to tap the factor at contact 90 ".

It can be seen that the /? System as in the Reality the essential factors of gravity and the speed of rotation around the vertical Axis of the aircraft contains and that the "^ system the factors" of the turning moment due to the rudder control and the chassis reaction or damping component is taken into account. The /S-S.ervo system can be conveniently used as shown to provide an inclinometer 79 to operate.

The integrating servo system for indicating roll (9j) status includes summing amplifier 77 which is responsive to the derived aileron voltage of potentiometer 16 (FIGS. 1 and 2). The voltage of the potentiometer contact 16 'is a speed voltage and is fed through a line 20 to the amplifier yj in order to represent the change in the rolling. The other input voltage is derived at the yaw system and represents the coupling factor between yaw and roll. The roll servo motor 80 and the generator 81 are connected to a reduction gear 82 in order ^{to actuate the three contacts 83 ″, 83 b} and 83 ° of a cosine potentiometer 83, which has the same general shape as the potentiometers used in the tilt servo system. Opposite voltages - \ - E and '- E are applied to the terminals of potentiometer 83 to represent gravity G. The contacts are all connected to servo shaft 84 for actuation, wherein the contacts 83 * and are relative to the contact 83 ⁰ at 90 83 ° or i8o °, so that the voltages corresponding to the component of gravity along the z-axis during rolling can be derived at the contacts 83 'and 83 ° , which represent cosine functions of opposite signs, namely + G cos φ and - G cos φ, and the voltages that result at contacts 83 ^ and 83 ^{& δ} are sine functions of opposite signs, namely + G sin φ and - G sin φ . These sine function voltages feed the pitch potentiometer 130 via lines 131 and 132, so that the derived voltage at the sliding contact Ι3θ ^α corresponds to the value G cos θ sin φ , ie along the gravity component. the y-axis when rolling. This voltage from contact 130 ⁰ is fed through a line 85 to the device shown in FIG. 1 for calculating the force reaction.

In order to introduce the pitch correction factor, the pitch system potentiometer 86 is fed by the roll system, and for this purpose the derived voltages + G cos φ and -G cos φ of contacts 83 "and 83" of the <p-potentiometer 83 are used to adjust the To feed lines 87 and 88 and the corresponding terminals of the 0-cos potentiometer 86. The derived voltage at the contact 86 ^{a of} this potentiometer represents the function G cos φ cos Θ, ie the gravity component along the z-axis of the aircraft, which is broken down for roll and pitch. The contact 86 ^a is connected to the device of FIG. 1 by a line 89.

The longitudinal inclination potentiometer 48, which is excited according to the reciprocal of the vehicle's own speed, is provided with a second contact 48 ^b which is offset by 180 ° from contact 48 ″, so that the wiper contacts in question are tapped

Tensions the values - | cos θ and cos θ '■

represent, the first value also an input voltage for the a-servo system described above forms. These voltages are applied by leads 91, 92 to the input terminals of the roll cosine potentiometer 90 supplied so that the voltage derived therefrom at the sliding contact 90 "the gravitational

factor - sin φ represents cos θ , where the voltage

is fed through a line 93 to the aforementioned ^ system.

If it is now assumed that the student pilot the rudder z. B. lays hard to the right, then

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increases the rudder _{torque tension for the co z} amplifier to the maximum amount, with the <M _Z system coming out of balance, so that the motor 66 ^{moves the potentiometer contacts 6y a} and 68 ^a upwards when the system seeks a new equilibrium position corresponding to the change in yaw. Since the derived potentiometer voltages fluctuate in size, the / S system is brought out of balance by the increasing angle change factor ω _ζ. The ß-system works like this, if one assumes that the roll angle φ has not changed, that a lateral slipping is indicated, whereby the motor 71 in this case ^{moves the contact 74 a} downwards in a direction corresponding to the position » ball to the left "corresponds to the known cross inclinometer, ie a" slippage "due to insufficient ^one bank. This condition can be corrected by increasing the roll angle φ , with the magnitude of the gravity component

- sin φ cos θ is increased in order to

counteracting input voltage ω _ζ . When these two input voltages are properly related; then the / - system is in equilibrium for a zero angle of the side slipping, wherein the contact 74 ⁰ is in the middle or zero voltage level and the ball of the transverse inclinometer is centered. The voltage derived at contact 74 ″ is, as mentioned above, used as the input voltage for the ω _ζ amplifier for damping in order to dampen the lateral slipping and to move the contact 153 to the center Pitch and roll angles are shown schematically in Fig. 4, in which an aircraft is shown in a combined pitch and roll position to the right. The pitch angle θ and the roll angle φ are between a horizontal plane defined by the fixed axes X ₀ and y _{0 is} defined and measured along the aircraft's x and y axes The vertical gravity or weight vector G from the center of gravity can therefore be projected onto the aircraft's yz plane as G cos θ, and this component in turn can be on the »-y plane can be projected as G cos θ sin φ This last component therefore represents the component of gravity acting along the wing and tends to cause the side slip. If this component exceeds the centrifugal force component, then there is a lateral slip; if it is smaller, then there is a j "push". It should be noted that the centrifugal force ω _ζ ■ ν for a certain turning speed becomes a direct function of the vehicle's own speed. Therefore, in order to make a curve without slipping sideways, the bank angle must be increased with increasing airspeed, so that the expression G cos θ sin φ = ω _ζ ν or vice versa

Expression - cos θ sin φ = ω »is. This is an important

actual simulation of the actual behavior of an aircraft.

With a suitable choice of the damping input resistances, the feedback voltages on the ß- and w. Servo systems and the gear ratios, the exercise system can be made to vibrate in the desired manner in order to reproduce the lateral vibration and damping of an aircraft.

In the vector diagram shown in FIG. 3, various centrifugal force and gravity component stresses are shown as vectors with reference to the x, y and z axes of the aircraft. These axes should be fixed with respect to the aircraft, so that they rotate with it and assume different positions of pitch θ and roll φ with respect to fixed horizontal and vertical reference planes. As already mentioned, the # -axis is the longitudinal axis of the aircraft, the y-axis is the transverse axis in the direction of the wings, and the x-axis is the axis perpendicular to the x- and y -axes. The force of gravity G, which acts on the center of gravity CG of the aircraft, ie at the intersection of the axes, can be broken down into vector components along the x, y and 2 axes, so that the ^ component G sin Θ, y component G cos θ sin φ and the ^ component G cos φ cos θ . The vector that represents the centrifugal force due to yaw (ω _ζ ν) acts along the y-axis additively to the gravity component G cos θ sin φ, and the centrifugal force vector as a result of the longitudinal inclination (coy ν) acts additively along the 2-axis to the gravity component G cos φ cos Θ. Also a vector that does the

Self-acceleration -7- acts along the

Ut

# -Axis additive to the gravity component G sin Θ. The resultant of the vectors in the% -2 plane can therefore be represented as

ill dv \ ²

R- = I / iGsinfl + - I + {ω _ν · ν + Gcosycosö) ² .

The resulting vector R _xz forms an angle Δ with the z- axis.

It can be seen that the resultant R of all vectors along the x, y and z axes and their angle with respect to the% -2 plane by combining the resultant R _xz with the sum of the y-axis components G cos θ sin φ and ω _ζ ■ υ can be obtained. . It surrenders

R = V (R _XZ ) ² + (G cos θ sin <ρ ^ ω _ζ - ν) ² .

From Fig. 3 it can be seen that the final resultant R includes an angle δ with the # - £ plane. For proper "Sitzgefühk the cabin of the airplane pupil should be inclined at an angle Δ longitudinally and at an angle δ to be inclined transversely.

In order to incline the car S according to these angles Δ and δ , two rotary transformers 95 and 96 are provided in Fig. 1, the secondary windings are in the neutral position, while the primary windings are fed by certain vector voltages of acceleration and gravity and the secondary windings Servomotors are controlled. The rotary transformer 95 is used for

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Used longitudinal inclination and has two fixed primary windings 97 and 98 which are arranged crosswise. The winding 97 is excited by a voltage,

that of the combined -axis vector I G sin 0 + -j-) V dt J

of the summing amplifier 99 corresponds, whose input voltages G sin θ and - from the flight computer the Eig. 2 through lines 24 and 49 are fed. The other winding 98 is excited by a voltage of the summing amplifier 100 which represents the sum of ω _ν ■ »and G cos φ cos θ , ie the combined centrifugal force and gravity vectors along the z- axis according to FIG. Axle voltages are fed to amplifier 100 through lines 61 and 89 from the flight computer of FIG. The gain of amplifier 100 can be adjusted by feedback in a known manner, so that the output voltage across coil 98 has the same relationship of magnitude as the voltage across winding 97 for equal accelerations.

The primary windings are therefore excited by voltages corresponding to the crossed vectors along the x and y axes. From this relationship it can easily be deduced that when the crossed coils 101 and 102 of the rotatable secondary winding are rotated relative to the primary winding by the shaft 103 so that the induced voltage in the secondary winding 102 is zero, the induced voltage in the coil 101 represents the resultant of the crossed voltages of the primary winding, ie the resultant R _xz ; In this position, the shaft 103 assumes the aforementioned angle A , which is measured from a reference position which corresponds to the plane position of the car S.

The terminals of the secondary windings are with

the external circuit via slip rings 104, 105 and 106 connected as follows:

The coils 101 and 102 have a common one Terminal grounded through slip ring 104; the coil 101 is through a slip ring 105 through a Line 107 is connected to the rotary transformer 96, and the zero voltage coil 102 is via a slip ring 106 connected to an amplifier 108 to a zero servo motor 109 to energize. The motor 109 is connected by a reduction gear 110 of the shaft 103 and is polarized so that the secondary winding 101-102 in such a direction is rotated so that the induced voltage in the coil 102 decreases. One induced in coil 102 Voltage therefore causes the motor 109 to be energized in a direction which reduces the coil voltage seeks. The motor 109 can be a two-phase motor of the type described above, with the connections for the reference voltage are omitted to simplify the drawing.

As a result, the angular position of the shaft 103, the longitudinal inclination of the cabin of the student pilot represents, this shaft can be coupled to the shaft 12 of the cabin bracket, so that the tilting movement of the cabin by suitable means, such as. B. a mechanical connection in, which is indicated by dashed lines, is effected as well as by a trailing mechanism or torque amplifier 112.

The rotary transformer 96, which is used for the transverse inclination, has a fixed primary winding, which consists of two crossed coils 113 and 114, and a relatively rotatable secondary winding, which has two crossed coils 115 and 116, which is driven by a shaft 117 with a servo motor 118 are connected in a similar manner as the rotary transformer 95. The primary winding 113 is fed from a summing amplifier 120 by a voltage which represents the vector sum ω _ζ · ν and G cos θ sin φ along the y-axis according to FIG. The input voltages for amplifier 120, which represent said vector components, will be '.' from the flight computer of FIG. 2 through lines 64 and 85. The other primary winding 114 is fed by a voltage R _xz from the secondary winding 101 of the rotary transformer 95 via a slip ring 105 and a line 107. The gain of the amplifier 120 is adjusted by feedback in a manner known per se, so that the output voltage at the coil 113 has the same magnitude ratio as the voltage supplied to the winding 114 for the same accelerations.

The crossed primary coils 113 and 114 are fed with voltages corresponding to the combined force vectors along the y-axis and the aforementioned resulting vector R _xz. As in the case of the transformer 95, it can be shown that when the crossed coils 115 and 116 of the rotatable secondary winding are adjusted by the servo-operated shaft 117 so that the voltage induced in the winding unit is zero, the induced voltage in the coil 115 is the same Represents resultant of the mentioned primary stresses; the shaft 117 assumes the angle δ , measured from a reference position which corresponds to the plane position of the car S. The servomotor 118 has a similar structure to the servomotor 109 of the transformer 95 and is connected to the transformer shaft 117 via a reduction gear 121 so that it rotates the secondary winding into the zero position, as just described. The motor 118 is fed by the zero setting coil 116 via a slip ring connection 122 and an amplifier 123. The coils 115 and 116 have a common terminal which is grounded via a slip ring 124, and the coil 115 can be connected via a slip ring connection 125 to an indicating voltmeter 126, e.g. B. is calibrated so that it shows the acceleration voltage acting on the student pilot and gives the instructor a visual display of the flight maneuver.

Since the transformer shaft 117 assumes a position corresponding to the angle δ , this shaft can be connected to the shaft 11 of the cabin bracket 9 by a suitable device, e.g. B. a mechanical connection 127, which is indicated by dashed lines, can be operationally connected and by a follower device or torque amplifier 128 so that the cabin is inclined by the angle δ transversely '. This bank control δ together with the above-mentioned one

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The longitudinal inclination control Δ of the shaft I2 tilts the cabin S of the trainee pilot in such a way that the real force effects on the trainee wing are simulated during the usual flight maneuvers.

Claims (12)

Patent claims: 1. Device to simulate the effects of flight forces on 'a student pilot who has the Control units of an aircraft located on earth: ίο exercise device operated, which is a movable Cabin for the student pilot and flight computer for simulating flight conditions as a function of the operation of the control devices as well as devices for calculating forces, which respond to the simulated flight conditions, characterized in that the last-mentioned Facilities' are appropriate to the direction of the. Resulting of simulated flight acceleration forces that act on the student pilot determine, and that drive means are provided which are based on the force calculation device address to move the student's cabin with respect to the direction of the resultant, whereby the trainee pilot is subjected to forces that correspond to those that he would perceive during real flight maneuvers. 2. Apparatus according to claim 1, characterized. shows that the force calculation device is a known conversion device that responds to components of the simulated gravity and inertial forces in order to the direction of the resultant of these forces with respect to the axes of rotation of the simulated Determine aircraft while the drive means a movement of the cabin of the student pilot in a corresponding direction with respect to the vertical axis of the student's cabin. . 3. Apparatus according to claim 2, characterized in that the force calculation device is suitable to generate control variables which. Components of the simulated centrifugal force and represent gravity acting in the direction of the reference axes of the modeled aircraft, wherein the conversion device is responsive to the control quantities of the components in order to obtain the Angle of the resultant of the forces with respect to the horizontal axes of the student's cabin to be determined, the drive means being controlled according to these angles, to tilt the student's cabin. 4. Apparatus according to claim 2, characterized in that the force calculation device is suitable to generate control variables, which the size of the components of the simulated Main acceleration forces including. represent the force of gravity, which along the pitch, Roll and yaw axes of the simulated aircraft act that the conversion device responds to these control variables of the components in order to determine the angles of the resultants of these Forces against the pitch and roll axes of the aircraft to determine, and that the drive means on the conversion device address to the cabin of the student pilot by the relevant angular amounts in the longitudinal direction to tilt, and cross direction. 5 · Device according to claim 3 or 4, characterized in that electrical Sizes, e.g. B. voltages are used and that electromechanical drive means are provided are. 6. Apparatus according to claim 5, characterized in that the conversion device has two Contains servo systems that respond to voltages, which components of the simulated Centrifugal force and gravity represent the angle of the resultant of these forces using In relation to the longitudinal and transverse inclination axes of the student's cabin. 7. Apparatus according to claim 6, characterized in that alternating voltages are used as control variables are used and that each servo system of the conversion device has a rotatable magnetic Includes converter that has two crossed primary windings, each on one AC voltage responds, which represents one of two perpendicular forces standing / and which has a secondary winding which is rotatable with respect to the primary winding, and a servo motor responsive to the voltage induced in the secondary winding to drive the To set the angle between the primary and the secondary winding to a zero value, which represents the angle the student pilot's cabin is tilted. 8. Apparatus according to claim 7, characterized in that one of the rotatable magnetic Converter has a further secondary winding, which is opposite to the zero position secondary winding is crossed, and that one primary winding of the second rotatable magnetic transducer with a voltage is fed which is induced in this second secondary winding. 9. Flight training device with simulated flight control devices and with flight computing devices that focus on the operation of the control devices in simulated flight conditions, e.g. B. airspeed, pitch change, angle of attack 'and pitch, respond, characterized by a device (17) which is responsive to the operation of one of the flight controls to derive a first voltage representing a pitch moment M _v , further by a computing device (55, 56) for the change in inclination, which responds to changes in the value of the torque voltage, as well as an equilibrium voltage, which represents the reaction inclination moment C _m , in order to derive a voltage which reflects the change in longitudinal inclination ω _υ , furthermore by means of a device (25, 26) for deriving Control quantities which represent various functions of the airspeed, further by a device (48, 44) which is jointly responsive to airspeed control quantities and other quantities 656/316 D 18685XI / 62c to derive additional stresses that affect the pitch angular velocity due to gravity or as a result of the buoyancy - set, further by a device (41) to algebraic summation of pitch change and pitch velocity stresses, furthermore by an integrating device (40) which, according to the algebraic sum of these voltages is operated to determine the angle of attack α and also to determine the pitch rate mentioned due to the buoyancy - to determine, and finally by an integrating device (50, 51), which is responsive to the pitch calculating device (55, 56) to determine the pitch condition θ of the aircraft and also to determine the pitch velocity stress due to gravity G cos θ ■ ...,
-; to determine. 10. Apparatus according to claim 9, characterized by a feedback device (94) which responds jointly to longitudinal inclination change and airspeed variables in order to change the response of the integrating device for simulating a damping ω _ν ν. 11. Apparatus according to claim 10, characterized in that the computing device (55, 56) of the longitudinal inclination change contains a summing amplifier (55) for algebraic addition of the voltages, the longitudinal _{inclination angle moment M 3} ,, the retroactive longitudinal inclination moment C 7n and the damping ω _ν ν represent, as well as a servo motor (56) which is controlled by the summing amplifier, and that the feedback device (94) contains a potentiometer (94) which is fed by the airspeed voltage and adjusted by the servo motor (56) in order to derive the voltage, which reproduces the damping ω _υ ν. 12. Flight training device with simulated flight control devices and with flight computing devices that respond to the operation of the controls in order to avoid flight conditions, e.g. B. to imitate yaw change, lateral slipping and rolling, characterized in that a device (18) is provided which is responsive to actuation of one of the controls in order to derive a voltage which represents a yaw angle moment M _t , furthermore a device (16) which is responsive to the actuation of another control to derive a voltage ¹ representing a roll angular velocity M _r , further a computing device (65, 66), which is responsive to the yaw moment voltage, in order to derive other control variables which the yaw change ω _ζ . furthermore an integrating device (77, 80) which is responsive to the roller speed tension and one of the yaw change control variables in order to indicate a relative position which indicates the roll angle ψ , furthermore means which can be set by the integrating device [jj, 80) for deriving a tension, one component of gravity in the direction of the lateral slipping G sin ψ cos 0 represents, and a second Computing device (70, 71) responsive to both another of the yaw change quantities and the gravitational component tension to indicate a relative position which reflects the angle β of the lateral slip. 1 sheet of drawings