DE1029238B - Control device in aircraft for the prevention of buoyancy - Google Patents

Description

GERMAN

The invention relates to the altitude control of aircraft, in particular to a device through which a Aircraft that has wing flaps or other aerodynamic devices to change the aircraft's lift is equipped at a given angle of attack. This device is intended as a function of the deviation from a nominal longitudinal inclination alone or, combined with its vertical deviation, from one given trajectory can be controlled and the aircraft can be protected from the buoyancy that occurs as a result the operation of an aerodynamic device and as a result of changes in airspeed, the one accompany or follow such activity.

Usually, a pitch reference instrument is used to predict the deviation of an aircraft from a flight path of constant altitude or from a beacon, ζ. B. a slideway after the blind landing procedure, displayed. In this context, a signal indicating the deviation of the aircraft from a Specifies longitudinal inclination from a reference instrument, e.g. B. a plumb line, taken and algebraically with a signal can be combined, which the vertical deviation of the aircraft from a nominal altitude or a Indicates radio slide. In such a case, the altitude deviation signal is usually switched from one to the other Height appealing facility, e.g. B. a barometric or radio altimeter, mediated while the slideway deviation signal is taken from a radio receiver that is sent to the slideway transmitter is matched. The combination of one of the elevation or slideway signals with the pitch signal represents usually a combination of a deviation term with its first time derivative Aircraft controlled in such a way that this signal combination retains the value zero, either by hand control by a zero indicator controlled by the combined signal or by an automatic pilot control can happen, the aircraft asymptotically approaches the target height or glide path and maintains it after that a.

While the aircraft is being controlled manually or automatically for altitude and pitch, its desired pitch is usually set manually to match the angle of attack at which the aircraft must fly in order for its lift to be equal to its weight. With this nominal longitudinal inclination, the nominal height is maintained. W ^T hen over time changes the flight speed and the overall weight of the aircraft reduces fuel consumption, Sollängsneigung in accordance with a corresponding change of the required angle of attack is adjusted. Usually, however, when the airspeed changes during altitude control (travel condition), it is not necessary to adjust the nominal pitch

Control device in aircraft
to prevent buoyancy surges

Applicant:

Sperry Rand Corporation,
New York, NY (V. St. A.)

Representative: Dipl.-Ing. C. Wallach, patent attorney,
Munich 2, Kaufingerstr. 8th

Claimed priority:
V. St. v. America from 12/6/1954

Reuben P. Snodgrass, New York, NY (V. St. A.),
has been named as the inventor

is changed) because of small changes in cruising speed usually have little effect on the required angle of attack.

Assuming that the airspeed does not change when the aircraft controls move from altitude and pitch control is switched to slideway and pitch control, the target pitch becomes generally changed by an amount equal to the angle between the glide slope and the runway (usually 2.5 °). While the plane is still a few kilometers from the runway and his Flight is slowed down for the approach, the nominal pitch is often readjusted again by hand, see above that a constant buoyancy is maintained on the glide slope. This re-enactment is mainly used

to correct for the change in airspeed because the decrease in airspeed in this one Trap is so large that the lift of the aircraft is adversely affected. At the reduced or approach speed Even small changes in airspeed have a significant impact on aircraft lift.

Operated at reduced approach speed

the pilot the aerodynamic device z. B. by extending the wing flaps and changed thereby the lift and angle of attack characteristics of the aircraft. When extended, the aircraft can be flown at relatively low speeds while still maintaining a safe area between remains the actual angle of attack of the aircraft and the angle of attack at which the aircraft sags

809 508/35

and under certain circumstances the control can no longer be changed in such a way that the

reacted. Lift or approach angle characteristic depending on

With other aerodynamic devices, the extent of the adjustment of the aerodynamic

for this purpose the back inclination of the wing device changes. For adjusting the aerodynamics the fuselage longitudinal direction reduced or the 5 mixing device, the control device has a

Wing rotated around the transverse axes or even the device for measuring the pitch deviation

Wings are extended from the fuselage. on, which is a measured variable for the deviation of the aircraft

Such devices can moreover deliver together from the nominal longitudinal inclination and are designed in this way

be used with the flaps so that both is that the nominal longitudinal inclination simultaneously with one actions add to. Regardless of the shape of the io position of the aerodynamic device opposite

Device is set by their actuation of the aerodynamic a predetermined position and that

mixed condition of the aircraft changed and the lift the aircraft, if it is controlled in such a way that the measurement

enlarged on the wings. By using size for the pitch deviation remains zero before

Such devices therefore occur at a lower level of buoyancy as a result of the actuation of the aerodynamic angle of attack a greater lift on. * 5 mix facility is preserved.

With this increase in lift, the pilot must have the Further details and advantages of the invention

Readjust the desired longitudinal inclination, so that under the result from the description of the execution

Influence of the combined pitch and glide slope examples on the basis of the drawing. In this shows

signals a flight path corresponding to the glide path with Fig. 1 in the block diagram of an aircraft control system, constant buoyancy is achieved. 20 to which the invention can be applied,

The choice of the point in time at which the longitudinal inclination Fig. 2 has a characteristic curve which readjusts the change in the pitch is compared with the time of winkeis for lift zero of a given aircraft pitch change However, it is now extremely significant as a result of the wing flaps pivoting out energetic. Far outside the airfield on which represents

Glide path and before reaching the glide path can at 25 Fig. 3 schematically an arrangement for generating

Altitude control in general the nominal longitudinal inclination after a lift correction according to the invention,

the change in lift. In the case of the one in FIG. 4, which is a curve for another aircraft

Near the runway, part of the glide path, where certain changes in the flap position and against

the altitude is low and is an accurate control of the time plotted lift control signal, and

For safety reasons, however, it is important to schematically show other embodiments in FIGS. 5 and 6

that the change in the nominal longitudinal inclination essentially to generate a lift correction according to the invention

at the same time as the change in lift.

is carried out. Otherwise it would normally start with the mathematical basics

rapid actuation of the aerodynamic device given for the invention and the sizes by which

an increase in buoyancy and thus a strong lift - the extent of the

cause impulses. Such a lift impulse, which is determined by the setting of the nominal longitudinal inclination,

the aircraft pulls up out of the nominal flight path, can be derived.

For the purposes of this explanation it should be assumed that the aircraft is steered on a glide slope determined by radio by reducing the engine power. The control can be achieved by or by the air resistance of the flaps swung out by the pilot by removing the altitude controls. In view of the low and steady tion of the aircraft so operated that an on the algebra-decreasing altitude at which these processes normally take place ische sum of signals for the deviation from, only a small period of time remains for the Beur glide path and the deviation from the Longitudinal slope division of the position by the pilot left when he 45 responsive display device indicates whether this sum should try zero, at the same time by changing the target is. Such an arrangement is known
longitudinal inclination by an estimated amount can prevent the lift on the other hand, the control by other known shock. automatic controls are carried out, by means of which a similar situation in which the aircraft is immediately controlled in such a way that the said algebraic intervention prevents a dangerous surge of lift, the sum of 50 is kept at the value zero. This must be occurs when the pilot initiates a scenic flight trap reads the equation for altitude control as is. In this case the aerodynamic entry becomes:

direction operated so that the buoyancy decreases, and it can ρ ρ _ £> m

a control for constant altitude can be switched on, R-

until a sufficient airspeed for a 55 where P = pitch,

Ascent is achieved. However, if the nominal longitudinal inclination p _R _ nominal longitudinal inclination,

not substantially simultaneously with such a £> - deviation from the slide.
Actuation (swiveling back the flaps) changed

and thus a corresponding loss of lift Equation (1) shows that with a vertical displacement

is compensated, a strong impact, namely a 60 of the aircraft against the glide slope, can reduce the pitch

Sagging, occurring in which the aircraft is at least the aircraft against its setpoint by the

falling in the direction of the runway until an altitude control is changed, until a signal that the

Altitude deviation signal is generated indicating an increase in the difference between the new pitch and the

the target height causes. Here, too, the pilot is only given the nominal longitudinal inclination, which is equal to a signal that the

is a dangerously small amount of time for the displacement to change from the glide slope.

Nominal longitudinal inclination by an estimated amount as long as the nominal longitudinal inclination represents the direction

Available to prevent the uplift. in which an airplane must be in order to go straight,

According to the invention, a control device is to fly parallel to the glide slope when there is no radio signal

Aircraft provided, which is available with an aerodynamic, is according to the track

Device is coupled through which the flight condition 70 chung (1) controlled aircraft asymptotically on the

5 6

Glide path, where the deviation from the assumption that the aerodynamic device

Glide path zero is: direction to change the lift of the aircraft at one

ρ _ ρ __ Q / 21 given angle of attack consists of wing flaps,

^R ' remains the inclination mla and the area S of the wing

where P = Pr. 5 constant. Of course, the weight W also remains im

essentially constant. Therefore, in this case, the

The basic quantity that simplifies the vertical flight direction equation (8) as follows:
but it is the flight path angle and not the.

Pitch. The flight path angle is the difference p _R - q _r _ | _ / L _j_ A _L o). (9)

between the longitudinal inclination and the angle of attack of a to \ 1 I.

Aircraft: _w

η t> λ ii \ where K ₁ = = -.

G = PA, (3) ^τ mLA · S

where G = flight path angle, when applying the invention to an aircraft

P - pitch, ¹ O ^ ^m wing flaps is therefore zero buoyancy

A = angle of attack. ^em signal proportional to the angle of attack A lo

generated by a device that focuses on the position of the

From equation (3) it follows that if the pitch flap responds, since changes in A _L o in angle are not constant, the pitch must be changed primarily by changing the flap position so that the trajectory angle G is constant and 20 is produced . Thus, the expression can mean that the aircraft is kept on a straight path. the angle of attack at zero lift is written as follows

With a constant flight path angle, the opening must be:

_{propulsion of} the aircraft must be equal to its weight (under the a lq

Assumption that with airplanes usually small flights- ^ -lo = = 0 ° + f (F), (10)

path angle and a load factor 1 occur). Hence 25

applies where

W = L = qC _L -S, (4) ^ L = 0 ° = angle of attack during lift NuU

where W = weight, with flap deflection,

L = buoyancy, 30 zero = constans = K ₂ ,

q = dynamic pressure, f (F) = change in the angle of attack during lift

Cl = lift coefficient, zero for a flap position F. S = area of the wings.

Substituting equation (10) into equation (9), so

The lift coefficient Cl can also be obtained from 35 the following expression for the nominal longitudinal

The angle of attack can be expressed as follows: Inclination Pr :

Cl = MLa- (A-Al ₀ ), (5) _{Pr = Qr +} (K ± _{+ f {F)} \ ₊ ^ _ _(U)

where mla = slope of the lift coefficient or

Angle of attack curve of the aircraft, 40 It therefore follows from equation (11) that a signal,

A = angle of attack, which is a function of the flap position F , for

A lo = angle of attack with zero lift. an angle of attack A _L o is used with zero lift

can be.
Inserting equation (S) into equation (4), when the invention is applied to an aircraft

the following expression for the angle of attack: 45 .. _Tr1 . , ■, _A ,, K _{1 n,.} , _{/ n} ,

⁰ with flaps becomes the expression --- from equation (y)

w y

A = - + Alo. (6) and (11) one of three, depending on the circumstances

q mLA j basic methods:

_c ,, / -Ί · t. / λ \ 01 -u ιί \ ^a ) If the operational airspeed of the

Substituting equation (6) _m equation (3) s, so _{5o ste} 4 _{rten F} i _{aircra to} £ ° _{rch mor} e or less

the following expression is obtained for the flight path angle G: _{wdte pivoting of} the flaps by small amounts

/ W \ α / 7 \ changes or if at the same time with a valve

k ⁼ ~ \ q ~ ^ m _LA · S ) ° '"' actuation of the operational airspeed in principle

",,, ™. . jj _o ",. , _{/ o} > changes, the expression - ^L is treated as a constant.

Should the aircraft be controlled so that equation (2) 55?

is satisfied, and to the flight path angle G constant Therefore is achieved in this case by a fixed Einbleiben, so it is convenient to equation (7) as follows to position the signal strength for the expression A _L o or f (F) write, the required knowledge the influence of flight

speed to the level necessary for constant buoyancy '' 60 manoeuvrable setting of the nominal longitudinal inclination.

b) Stepping on a controlled aircraft, where Gr - constant target flight path angle. known changes in the given operational flight

speed after a flap has been actuated, so will

From equation (8) it follows that a change in Klam-, _A ■, ·, K ₁ ...,,, -ρ. , ^ ₁

,,. '·. · V jc 11 »· π, the expression --- during the short duration of the flaps-

More expressly a change in the nominal longitudinal inclination Pr 65?

required if the given trajectory angle G ^ a actuation is treated as a constant and is then to be behaved as a behavior. According to the invention, therefore, changes are made with known change with respect to time. Changes in the target pitch depending on Accordingly, in this case, a time device is used, changes in the parenthesized expression of the equation (8), which is used to carry out a given continuous change. 70 the signal strength of the expression A _L o or f (F) during

a given period of time after actuation of the flaps.

c) Finally, in the general case, which includes all control states, including those mentioned above, in which

TJ "

the printout - based on the operations

is estimated, this expression is supplied as an independent signal from a device responsive to the airspeed and is used together with the signal for A _L o or f (F) to change the nominal pitch.

In an airplane with flaps, the target longitudinal inclination is therefore changed as an estimated or measured function of flap position and airspeed.

Corresponds to the inclination that the aircraft must maintain in order to be parallel to the flight path in the absence of a deviation signal.

In addition to the hand control, which in accordance with the algebraic sum of the trajectory and des Pitch deviation signal acts, an automatic control is provided, which also the output signal of the addition amplifier 5 for controlling a

position alone, while the expression - the equation

An alternating current signal which is similar to the output signal of the height control device 8.

The output of the addition amplifier 5, which is the algebraic sum of the pitch error signal and is one of the trajectory deviation signals, via a rectifier 11 of the coil 12 is one Zero display device 13 supplied, which preferably has a horizontal pointer 14, which is through the Coil 12 is moved relative to its zero position, ίο the pilot controls the elevator 15 of the aircraft of the hand control 16 and holds the pointer 14 in the zero position and thus also the algebraic sum of the If the flight path and pitch deviation signal are zero, the aircraft becomes asymptotic

However, if the aircraft has, in addition to the flaps, an infeed 15 of the flight path in question and then this device for swiveling the wings about a transverse hold, provided that the nominal longitudinal inclination of that axis and the plumb line indicated for the longitudinal inclination is arranged in the fuselage, then the expression A10
from equation (9) represented by a signal that both
the position of the flaps and the inclination of the 20th
Wing depends on the fuselage, as one
Change in wing inclination a change in Α _τ ο
evokes. Does the airplane have flaps and also one
Device for changing the wing extension,
the expression Αχ, ο is used in the same way by an electrical servo device 17 and is supplied via a signal that depends on the flap position and an electrically operated clutch 18, the elevator 15 as well as on the wing extension, since one is set. Closing a single-pole switch 19 changes this expansion , causing a change in Ακ> the clutch 18 is fed from a power source. whereby the elevator 15 by the servo device

If, on the other hand, the aircraft has, in addition to the flaps, a 3 ° is automatically controlled so that the aircraft the Device for changing the wing inclination, adheres to the desired trajectory.

so the expression A ^ o becomes a function of the valve. The inventive features are therefore in the

Buoyancy control device 1, which supplies an output signal in a line 20, by means of which the (11) is set not only as a function of the estimated or desired longitudinal inclination of the system so that the measured dynamic pressure is q , but also protects an aircraft from buoyancy is, the function of the measured angle of inclination back, since the on the actuation of the lift-changing aerodynamic component Mla of K ₁ mix device and changes in flight as a function of this angle, follow speed, with the flight speed 1 with 40 speed during or after the actuation of this device with the exception of the buoyancy control 1 forms a known changes.

System by which the pitch of an aircraft was derived from equation (9) that the notin Depending on both its vertical deviation, agile setting of the nominal longitudinal inclination is a function towards a chosen trajectory as well as from its angle of attack for the lift zero and the under Deviation from a nominal longitudinal inclination is controlled 45 congestion resulting from the influence of the airspeed, pressure is.

A pitch deviation signal will be used for a flap-equipped aircraft

device 2 supplied by a signal generator 3, which is based on the changes in the angle of attack for the lift zero Transverse axis of a gyro 4 is arranged. This caused by changes in the flap position, Signal can be in an addition amplifier 5 with a 50 so that a signal which is a function of the flaps through the line 6 is supplied trajectory deviation position, can be combined algebraically to represent the angle of attack for the open signal. This zero drift can be used [cf. Equation (11)]. In Fig. 2, the change is for a specific aircraft

Wegempfänger 7 or from a height control device 8 f {F) of the angle of attack for the lift zero against the supplied, depending on the setting of the switch 9, the three 55 positions of the aircraft flaps shown. From the result possible positions. one can see that when the

If the desired trajectory is indicated by a glideslope transmitter, switch 9 is brought into its illustrated position "G" . If, on the other hand, the desired flight path is to have a constant altitude, switch 9 is set to »Hi. set. A position »Off & is also provided. This allows the system to be completely switched off from the beacon or the altitude-sensitive device if the pilot prefers to

The addition activity of the amplifier 5 itself is used by the winding 21 of a potentiometer, which is connected to a to lead. AC power source 22 is and one side is grounded

A modulator 10 is located at the output of the glide path.

receiver 7 in front of the switch 9 and forms the normal- The sliding arm 23 of the potentiometer is mechanical

wise consisting of direct current output signal via a non-linear transmission device 24 with the Receiver connected into a proportional phase reversing 70 shaft of a reversible electric motor 25,

swiveled flaps the angle of attack for the lift zero changes nonlinearly in the negative direction, i.e. with the pivoting of the flaps decreases.

Fig. 3 shows an embodiment of the buoyancy control 1 (Fig. 1) according to the invention, in which an output signal which changes with the flap position when there is a change in the angle of attack for lift zero occurs. To generate the signal

which is also used to actuate the flaps 26 and a device 27 indicating the flap position.

A pole reversing switch 28, which is switched on between the motor 25 and the battery 29, controls the swinging out or swinging back of the flaps 26, depending on which position the pilot has switched on. The non-linear transmission 24 has, for example, a suitable curve guide through which the grinding arm 23 is guided non-linearly with respect to the position of the flaps 26 in accordance with the curve shown in FIG. Assuming that the potentiometer itself is non-linear, a signal is therefore generated between the grinding arm 23 and earth which changes with the flap position when there is a change f (F) in the angle of attack with respect to zero lift.

The signal of the sleeping arm 23 is via a line 20 to the terminal 31 of a winding 32 of the changeable Transformer supplied, the other winding 33, which from an AC source is fed, is rotatably connected to the transverse axis of the plumb line 4. For manual adjustment of the nominal longitudinal inclination a button 35 is provided through which the winding 32 can be pivoted. The other terminal 36 of the Winding 32 is connected to addition amplifier 5 (FIG. 1) via line 2. This will make the Potentiometer and encoder signals occurring at the amplifier input are superimposed on one another.

Between the potentiometer signal derived from the flap position and the transmitter signal there is a such a phase relationship that the latter always decreases by an amount equal to the former will. As a result, when the flaps 26 are pivoted out, the aircraft is flying with its nominal longitudinal inclination and the pitch signal is therefore zero, despite having a net signal between line 2 and ground. It has so it appears as if the nominal longitudinal inclination has been reduced. The nominal longitudinal inclination is thus electrical is set as a function of the damper position, and the net signal in line 2 is reduced to zero when the The aircraft is brought into the set pitch.

The lift control 1 described in connection with FIG. 3 is particularly suitable for an aircraft, whose flight speed is already known from operating procedures when the flaps are swiveled out and changes little either during or after the flaps are pivoted out or changes during the pivoting out the flaps fundamentally changes. As above

mentioned, the size --- of equations (9) and (11)

in such a case are treated as a constant, in which case no special facility is required, through which a further nominal longitudinal inclination setting for

—'- ■ is supplied, except that produced by potentiometer 21 for f (F) or A _w .

However, there are aircraft in which the airspeed has changed from a known operating value to one another known value changes significantly during a certain period of time after the flaps are in their desired setting are brought. One for setting the nominal pitch of such an aircraft a suitable lift control signal would take the form of FIG. 4 as functions of flap position and time signal. It is assumed here that the nominal longitudinal inclination to be set is the inclination in which a constant lift is generated with flaps swung out to 20 °. This flap position is usually used for the approach condition as opposed to the fully deflected flap position in the Landing.

FIG. 5 shows an embodiment of the lift control of FIG. 1, which represents an output signal for setting the nominal longitudinal inclination in accordance with the signal curve of FIG.
In FIG. 5 and in FIG. 3, a signal transmitter for setting the nominal longitudinal inclination of the system is connected in series with the winding 32 of the longitudinal inclination transmitter. Instead of a non-linear transmission device in the positive connection to a linear potentiometer, however, an otherwise equivalent non-linear potentiometer 37 can preferably be used, which is wound so that the desired signal is generated in accordance with the position of the grinding arm 38.

With flap positions of 0 to 20 °, the grinding arm 38 remains in its zero position, in which it has no output signal supplies. However, as soon as the flaps are swiveled out over 20 °, one of the flap motor 25 closes driven cam 39 at the same instant

such a normally open switch 40, one of which Terminal is connected to the positive side of a battery 41 and the other terminal via the winding a relay 42 and a normally closed switch 43 to the negative side of the battery 41 and earth is connected. In this way, relay 42 is excited and its holding contact on switch 43 closed, whereby the switch 40 actuated by the cam disc bypasses and thereby energizes the relay is held.

In this state, the switch 44 is also closed, which connects the positive side of the battery 41 to one of the Connects winding terminals of an electromagnetic clutch 45, the other terminal of which is grounded. is the coupling 45 is excited in this way at a flap position of 20 °, so it connects the shaft of the flap motor 25 mechanically with one input side 46 of a mechanical differential actuator 47, the other input side 48 is prevented from rotating at this time by friction from wheels. The starting page of the differential actuator 47 is mechanically connected to the wiper arm 38 of the potentiometer 37. on in this way, the grinding arm 38 is only set in motion by the flap motor when the flaps are over their 20 ° position are pivoted out.

Furthermore, the grinding arm 38 is rotated only in accordance with the flap position until the Flaps reach their full deflection of e.g. 52 °. In the meantime, in which the flaps are out of their approach position are brought into their landing position, so a potentiometer signal is generated, which is non-linear changes with the flap position. The potentiometer 37 is wound so that it is in Fig. 4 against the flap position applied, non-linearly changing lift control signal generated. With full flap deflection

includes one in the transmission between the differential actuator 47 and the grinding arm 38 lying cam 50 is currently a normally open Switch 51, one terminal of which is connected to the positive side of the battery 41 and the other terminal of which is connected

bo conclusion about the winding of a relay 52 and a normally closed switch 53 is grounded. The relay 52 is energized in this way, its holding contact closed at switch 54, whereby the switch 53 actuated by the cam disk is bridged and thereby the relay 52 is kept energized.

By energizing the relay 52, the contact on the switch 55 is also closed, through which the positive side of the battery 41 is connected to one of the terminals of a time motor 56, the other terminal of which via a speed control

809 508/35

Resistor 57 is connected to earth. The shaft of the motor 56 drives the other input side 48 of the differential 47 via gears. At the moment in which the flaps reach their full deflection, the motor 56 drives the differential actuator 47 Grinder arm 38 returns to its neutral position at a speed proportional to that by setting of the resistor 57 is a certain motor speed.

The terms occurring in these equations, _{m is} ¹ A and f (F) as shown in FIG 6 by a ^nonlinear>:. Potentiometer 37 (see also Fig. 5.) Complied with the "is operated in · this case, directly by the flap motor 25. The winding ends of the potentiometer 37 are i; in turn connected to an alternating current source for supply ·:. However, one end is not grounded! but connected via a line 65 to a rotor 66 which is connected to a fixed contact sector 67 in a

When the motor 56 stops the grinding arm 38 by a distance. This sector is by means of a line 68

K ₁

with the sliding arm 61 of the - ^{L -potentiometer 60}

bound. The rotor 66 is moved by the flap motor 25 from an insulating sector 69 to the contact sector 67

has adjusted the correct amount, the cam disk 50 opens
momentarily the switch 53, whereby the relay 52 is de-energized and as a result the motor 56 is stopped.
The cam disk 50 is designed in such a way and the resistor 57 used for speed control is turned in as soon as the flaps 26 are moved from their zero position, so that the potentiometer 37 supplied by the potentiometer 37 is swiveled open. In this way they are

Output sides of the potentiometer 60, the potentiometer 37 and the winding 32 of the solder circuit 4 with one another connected in series, whereby the longitudinal inclination in accordance with the functions of the dynamic pressure and the flap position is adjusted, both by

drive control signal is reduced by a given value within a predetermined period of time, whereby the signal is given the form plotted against time in FIG. 4.

To reset the device according to FIG. 5, the pilot actuates the polarity reversal switch 28 by hand, whereby the Flaps 26 are pivoted back. With a flap position of less than 20 °, z. B. 15 °, the cam opens 39 momentarily the switch 43, whereupon the relay 42 drops out and the electromagnetic clutch 45 disengages. During the pivoting back of the flaps, the potentiometer 37 is automatically in its zero position brought back.

Potentiometer signals are shown. The potentiometer signals have such mutual phase relationships

Draws that the --- signal from the A w or f (F) -

Signal is withdrawn.

The arrangements shown in FIGS. 3 and 6 are not only for changing the longitudinal inclination when pivoting out the flaps, e.g. B. during the approach that the arrangements according to FIGS. 3 and 5 are the best runway, but are also suitable then

possible adjustment of the longitudinal inclination bring about effective when the flaps are swiveled back,

z. B. missed landing or turning. In any case, the slope is set correctly, see above that no lift jolts are exerted on the aircraft,

speed, so there is neither jumping nor sagging.

The arrangement shown in FIG. 5 is, however, the

For simplicity as a non-reversible system only for setting the slope during an approach to a runway. Nonetheless, it can obviously by replacing the time motor 56 (Fig. 5) with a reversible time motor and by

and thereby prevent buoyancy surges caused by both the flap actuation and during or changes in flight compliance after this activation. For best results, the Pilot therefore swing out the flaps at the speeds which are the basis for the estimated A ',

Values - 'form when setting these arrangements can be used on the slope. corresponding change of the associated switch the In the arrangement according to FIG. 6, the pilot needs longitudinal inclination as a function of the flap position and on the other hand, there are no specific operating procedures at the time both when swinging out and when aiming back to achieve the best results, as the flaps will stop swiveling out.

j ι ΑΙ. ,,,, ..., "While the extent described exemplary pressure - - does not have to be estimated in advance, but, · ■,.. ,. . ,,. , _A °

q ° shapes to change the lift with a given propulsion

however, flaps can be used by means of an adjusting angle that is responsive to the airspeed Facility is measured. Accordingly, other aerodynamic devices are of course shown in FIG a non-linear potentiometer 60 with a sliding arm 50 alone or together with flaps to achieve an im 61 provided, which can be used by the movable end of a substantially equal result. Pressure cell 62 can be adjusted. The interior of the capsule becomes such other aerodynamic devices is used above by means of a Pilot's tube 63, so they can easily be used in any of the sustained dynamic pressure. The potentiometer 60 and FIGS. 3, 4 and 6 are set in place of the flaps shown Pressurized can 62 can be accommodated 55 in a container 64. The motor 25 would then be in place of the and attached, the interior of which by a pipe 65 to the flaps 26 either the back inclination of the wings

or change the angle of inclination of the wings when pivoting about a transverse axis or the lateral extent of the wings relative to the fuselage. If the 60 engine is used to make one of these changes, e.g. B. the potentiometer 37 according to FIG. 6 only to be wound so that it supplies a signal which is proportional to the change in the grinding arm caused by the aerodynamic device used. If the potentiometer is 60 so 65 is the bracket expression of equation (8).
wound that its output signal of the position of its knows on the other hand the device both flaps as

Grinding arm is inversely proportional, the off is also a device for performing one of the above If there are changes, a special actuating motor is preferable. In this case a (11) directly proportional. 70 special signal generators in the manner of a potentiometer

static pressure. With this arrangement, the grinding arm 61 is adjusted as a function of the dynamic pressure q and thus in accordance with the square of the airspeed.

The winding of the potentiometer 60 is connected to an AC power source, one side of which is grounded, so that an output signal occurs between the grinding arm 61 and the ground, the size of which depends on the position of the

output signal corresponds to the expression - - of equations (9) and

driven by this motor and is electrically in series with the flap potentiometer and the encoder winding 32 of the gyro 4, so that the changes in the lift or angle of attack characteristic curve caused by actuations this device and the flaps are caused by a corresponding change in the the pitch signal fed to the addition amplifier 5 are accompanied.

Claims (18)

Patent claims: 1. control device to prevent uplift, in aircraft that are equipped with an aerodynamic device which is opposite * 5 the rest of the aircraft can be adjusted so that the state of the aircraft is changed in that way is that the lift or angle of attack characteristic curve as a function of the adjustment of the aerodynamic Device is changed, which are also equipped with a device through which the aerodynamic Device can be adjusted as desired, wherein the control device has a reference instrument, which the deviation of the Aircraft determined by a nominal longitudinal inclination, characterized by a device through which automatically the nominal longitudinal inclination set and at the same time the aerodynamic device compared to a predetermined one Position is adjusted, the aircraft by actuating the aerodynamic device in front Buoyancy is preserved when it is controlled so that the measure of pitch deviation Remains zero. 2. Control device according to claim 1, characterized by an arrangement which ensures that the set nominal longitudinal inclination is a predetermined function of the at a given angle of attack change in lift generated by actuation of the aerodynamic device. 3. Control device according to claim 1, characterized by an arrangement through which ensures that the set nominal longitudinal inclination has a predetermined function for adjusting the aerodynamic Establishment is. 4. Control device according to any one of the preceding claims, characterized by a reference instrument, which the deviation from the longitudinal slope measures and which has an encoder which provides a pitch deviation signal which is a measure for the deviation of the aircraft from the nominal pitch, and by a display device or a Servo amplifier for the elevator, to which the pitch deviation signal is fed, and through means for generating a bias signal the magnitude of which has a predetermined relationship has to adjust the aerodynamic device, and in that the nominal longitudinal inclination is set by sending the bias signal to the display device or servo amplifier in the opposite direction to the pitch deviation signal is fed. 5. Control device according to any one of the preceding claims, characterized by one according to Drive device to be actuated and to adjust the aerodynamic device and thereby, that the nominal longitudinal inclination is set by a device that is operated by the drive device is operated while this is adjusting the aerodynamic device. 6. Control device according to each of the preceding Claims, characterized in that there? 70 Bias signal as output of a potentiometer is generated, the grinding arm of which by means of a non-linear measuring device simultaneously with the adjustment of the aerodynamic device is set. 7. Control device according to any one of the preceding claims, characterized in that the Bias signal is generated by a non-linear potentiometer, the sliding arm of which is simultaneously with the adjustment of the aerodynamic device is set. 8. Control device according to any one of the preceding claims, characterized in that the nominal longitudinal inclination is adjusted when the aerodynamic device from its zero position in which it does not Has an influence on the lift of the aircraft, is moved out. 9. Control device according to any one of the preceding claims, characterized in that the nominal longitudinal inclination is also set as a function of the airspeed. 10. Control device according to any one of claims 1 to 8, characterized in that the nominal longitudinal inclination is also set inversely proportional to the square of the airspeed. 11. Control device according to any one of claims 1 to 8, characterized in that the nominal longitudinal inclination depending on the adjustment of the aerodynamic device, regardless of the Airspeed is set and that after adjusting the aerodynamic device more than a predetermined amount of airspeed changes by a subsequent, one Change of this setting of the nominal longitudinal inclination which lasts for a certain time must be taken into account. 12. Control device according to any one of claims 4 to 8 and 11, characterized in that the following Change is caused by a signal, the magnitude of which changes over time with a predetermined function changes, and that this signal in the opposite direction to the mixed signal is fed. 13. Control device according to any of the preceding Claims, characterized in that the aerodynamic device has wing flaps, which can protrude from the trailing edges of the wings. 14. Control device according to any one of the preceding claims, characterized by a device which produces a signal that is a measure of the vertical deviation of the aircraft from a predetermined trajectory, e.g. B. a radio glide path, by a device for setting the nominal longitudinal inclination in accordance with the vertical deviation signal and by a display device or a servo amplifier for the elevator operated by the pitch deviation signal, whereby the aircraft follows the predetermined glide path under manual or automatic control. 15. Control device according to any one of the preceding claims, characterized in that the device to measure the deviation from the longitudinal slope is a plumb line. 16. Control device for aircraft that is equipped with an aerodynamic device, e.g. B. wing flaps, are equipped, which can be adjusted with respect to the rest of the aircraft, whereby the condition of the aircraft is changed in such a way that the lift or angle of attack characteristic depending on the adjustment of the aerodynamic see device is changed, and which are equipped with a drive device through which after If desired, the aerodynamic device can be adjusted in this way, with the navigation device a gyroscope reference device, e.g. B. soldering gyro, on which a transducer is arranged, the one Signal generated that is a measure of the deviation of the aircraft from a nominal pitch, characterized by means actuated by the drive device, issuing a first bias signal generated simultaneously with the setting of the aerodynamic device, the signal being a predetermined has functional relationship to the setting of the aerodynamic device, and by a Display device or a servo amplifier for the elevator and by a device that the Pitch deviation signal and the bias signal to the indicator or servo amplifier feeds in opposite directions. 17. Control device according to claim 16, characterized by a device which generates a second bias signal generated that has a predetermined functional relationship to airspeed, and by a device that sends this second bias signal to the display device or the servo amplifier, opposite to the first bias signal, supplies. 18. Control device according to claim 16, characterized by a device which the size of the first bias signal in accordance with a predetermined function during a Reduced predetermined period of time after the aerodynamic device has a predetermined Position. Considered publications:
British Patent No. 690,985;
U.S. Patent No. 2,613,050. 1 sheet of drawings © 809 508/35 4.58