DE1936406C1 - Method and device for monitoring a measuring device indicating the flight position of a missile - Google Patents

Description

Such inertial measuring devices give the Pilots now except for the respective; Position course and Speed values of his aircraft also for the alignment of his aircraft during maneuvers as well as important position reference values when the weapon is triggered. One way of being correct this indicated by the inertial measuring device To monitor position signals, there is also no so far, especially not when the occurring position error has such a size that he Already no longer portable for navigation and weapon triggering, but still when looking at the environment cannot be seen.

The invention is based on the object of a method and a device for performing the To create a method with which with the help of a relatively simple additional measuring device of respective monitoring time can be determined which the correctness of the position signals indicated by an inertial measuring device are monitored can.

Starting called by a process of the initially ge _fc type is B solved this problem according to the invention in that a comparison of the measured values is carried out both measuring means at a time at which the inclination angle φ and y of the missile with respect to its roll and pitch axis equal to zero and that this point in time is determined with the aid of the second measuring device in that the accelerations of the missile in the direction of the missile longitudinal axis, the missile transverse axis and the missile ^{vertical axis 1 are} determined and when the expression is determined

fold if it is assumed that the aircraft is is already in the horizontal position. The following then applies:

(b sin α. - b ₂ ) + (O ₃ ■: .- g)

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equals zero, where b = fb \ + b \, ^ _{1 is} the acceleration in the direction of the missile's longitudinal axis, b _{2 is} the acceleration in the direction of the missile's transverse axis, b _{3 is} the acceleration in the direction of the missile's vertical axis, λ is the missile's drift angle and g is the acceleration due to gravity .

In this way, the measured values of the inertial measuring device can be monitored easily and precisely, there is a missile, in particular a fighter aircraft, carried out during its mission flight Maneuvers relatively often, at least for a short time each, reaches an exact horizontal flight position in which the The angle of inclination with respect to its roll and pitch axes are equal to zero. Will be at this point the position reference signals displayed by the inertial measuring device with respect to these two axes read, it can be determined immediately whether the inertial measuring device is still working properly, or how big it is is a possibly existing display error.

Depending on the drift or slip angle, the first and the second accelerometer measure in the horizontal position of a missile, e.g. B. an aircraft, components of the aircraft acceleration. The drift angle must therefore be known with sufficient accuracy (max. Permissible error approx. ± 0.3 °). When using the Doppler radar as a speedometer, however, this condition is always met, since it can be used to determine the respective drift angle & very precisely.

The equations that can be used to calculate the aircraft's horizontal position are very simple

Vbi + b ₂ = b, (1) ' b ■ sin oi = b ₂ , (2) b ■ cos ot = K, ■ (3) with oi = Drift angle,

b%. = Acceleration in the direction of the aircraft longitudinal axis,

b ₂ = ■ acceleration in the direction of the aircraft transverse axis, - · ■

b ₃ = acceleration in the direction of the aircraft

vertical axis,
b = resulting aircraft acceleration.

As can be proven by investigations for all possible values of the angles and the angles of inclination γ and φ of an aircraft with respect to its roll and pitch axis, the aircraft is always in its horizontal position if the following system of equations is fulfilled:

- b ₂ - sin « U \ + b \ = 0, (4)

& 3-g = 0. (5)

In summary, the following must already be the expression mentioned earlier must be fulfilled:

sin CC-

The value of this equation is now calculated using the The above-mentioned device for carrying out the method is continuously determined and when it is reached the value zero is displayed.

If the value is zero, the aircraft is currently in its horizontal position. B. an inertial platform, displayed position reference signals z. B. in the form of the angle of inclination », γ and φ are given in relation to the yaw, roll and pitch axis, it can be determined whether this is still working with sufficient accuracy.

The invention is explained in more detail with reference to a block diagram shown in the drawing.

An accelerometer B ₁ , which is fixed to the cell in a missile (not shown here) in such a way that it measures the accelerations occurring in the direction of the missile's longitudinal axis, is via a first isolation amplifier V ₁ , a first low-pass filter T ₁ and a first squarer Q ₁ connected to one input of a first summing amplifier V ₂ . A second accelerometer B ₂ , which is firmly connected to the cell of the missile, not shown here, so that it measures the accelerations occurring in the direction of the missile transverse axis, is via a second transverse amplifier V ₃ , a second low-pass filter T ₂ and a second squarer Q ₂ connected to the other input of the first summing amplifier V ₂ . The output of the summing amplifier V ₂ is via a square root jR with a first input of a multiplier M; connected, the second input of which is supplied with a voltage signal which corresponds to the sine of the drift angle occurring in each case on the missile. This drift angle is determined by a Doppler radar D and emitted directly in the form of an electrical signal. The output of the multiplier M is

connected to one input of a second summing amplifier F ₄ , the other input of which is connected via an inverter / to the output of the second low-pass filter T ₂ .

Finally, a third accelerometer B _{3 is} provided, which is fixedly arranged on the missile cell that it measures the accelerations occurring in the direction of the missile vertical axis. The accelerometer B ₃ is connected to one input of a third summing amplifier V ₆ via _{a third isolating amplifier F 5} and a third low-pass filter T _3. A fixed voltage is fed to the second input of this summing amplifier V ₆ , the magnitude of which corresponds to the negative acceleration due to gravity g.

In addition, a fourth summing amplifier V _{1 is} provided, one input of which is connected to the output of the second summing amplifier V _{4 via a first absolute value circuit S 1} and the second input of which is connected to the output of the third summing amplifier V _{6 via a second absolute value generating circuit S 2} . The output of the fourth summing amplifier F ₇ is connected to a zero indicator JV ₁ , which always provides an indication when the sum signal at the output of the summing amplifier V ₁ just reaches the value zero. The output signal is also fed to a comparison circuit which is only indicated schematically here and which has a NOR gate G and a further zero indicator JV ₂ . The position reference signals α ', β' and γ ' emitted by the first measuring device A are fed to the other inputs of the gate G, the signal indicating the drift angle or the angle of inclination <*' with respect to the yaw axis of the missile being inverted at the input of the gate G .

With the help of this relatively simple circuit, the point in time is determined and displayed on the zero indicator at which the missile, not shown here, and which has the three accelerometers B ₁ , B ₂ and B ₃ , has just reached its horizontal flight position or passes through it. The circuit arrangement can work both analog and digital, the switching speed of the components used is still so high that a reliable display of this position on the zero indicator JV _{1 is} possible even in the event of a fly through and thus only a very brief occurrence of the horizontal flight position during a maneuver is.

At this point in time, with the aid of the simple electrical comparison circuit G and JV _2, the data from the first z. B. an inertial measuring device A specified position reference signals z. B. in the form of the angles of inclination α, φ, γ with respect to the yaw, pitch and roll axes and any errors that occur are determined qualitatively and quantitatively. The circuit arrangement shown in the block diagram works in the following way:

The measured by the accelerometer ₁ B acceleration signal Zj ₁ is a function of the total acceleration b of the missile and the gravitational acceleration g. This signal Z) ₁ is amplified in the first isolating amplifier V ₁ , smoothed in the low-pass filter T ₁ and squared in the squarer Q ₁ so that a signal W ₁ appears at its output. Similarly, the temperature measured by the accelerometer ₂ B acceleration signal Zj ₂ is processed. These two squared signals Zj ₁ and b \ are summed in the first summing amplifier V ₂ and pass to the square root R, at whose output a signal corresponding to the square root of the output signal of the summing amplifier V _{1 appears.} This output signal of the root eraser R corresponds to the resulting acceleration b of the missile in the horizontal plane. This output signal b is multiplied by the respective value sin «with the aid of the multiplier M and sent to the second summing amplifier V ₄ . given. The second summing amplifier V _{4 also} receives the inverted acceleration signal b _{2 from} the accelerometer B ₂ , so that the signal (Zj · sin «- Zj ₂ ) appears at the output of the summing amplifier V _4. The amount is formed from this signal in the circuit S ₁ and this amount is fed to the fourth summing amplifier V _1.

_{The acceleration signal Zj 3} determined by the accelerometer B ₃ in the direction of the missile vertical axis passes via the _{isolating amplifier F 5} and via the low-pass filter T ₃ to the third summing amplifier F ₆ , in which the acceleration due to gravity g is subtracted from the signal b _3. The output of the summing amplifier V ₆ therefore carries the signal (Zj ₃ - g). From this signal the amount is again formed in the circuit S ₂ and this is then given to the fourth (summing amplifier F ₇ ... The summing amplifier F ₇ sums both signals given to its inputs and forms an output signal of the form I (i> · sin a - Zj ₂ ) | + | (Zj ₃ - g) |, which reaches the zero indicator JV _1. The zero indicator JV ₁ therefore always gives an indication when the output signal of the summing amplifier F ₇ , i.e. the aforementioned expression, is even If the first measuring device A is still working properly at this point in time, the position reference signals ψ and γ it emits with respect to the pitch and roll axes are also equal to zero is performed, the monitored output of the measuring device and ^ from the Doppler radar speed measuring D position reference signal is χ with respect to the yaw axis at the input of NOR gate G inverted. A signal therefore only appears at the output of the NOR gate G if the position reference signals φ ' and / and the output signal of the summing amplifier F _{7 are} equal to zero, but the position (reference signal λ' is not equal to zero.

As can be seen from the functional description, important signal parameters are represented by angle functions. For this reason, with certain values of the missile accelerations Z) ₁ , b ₂ and Z) ₃ zeros can occur even though the missile is not in its horizontal position. This undesirable effect is avoided, however, if the acceleration _signals 6 l3 Zj ₂ and Z) ₃ from the circuit arrangement shown in the drawing, as described above, are only processed when the drift angle is greater than approximately 2 degrees. This condition does not represent a restriction of the method specified according to the invention, since drift angles of more than 2 degrees, even without cross wind and without lateral thrust components given by the propulsion of the missile, occur very frequently due to flight maneuvers alone, so that sufficient Time intervals with the aid of the device shown, a safe and simple monitoring of the measuring device indicating the actual position reference signals of the missile is possible.

1 sheet of drawings

Claims (3)

Patent claims: 1. A method for monitoring a first measuring device indicating the flight position of a missile for correct display with the aid of an additional second measuring device that works independently of the first measuring device, whereby a possible deviation from the actual flight position is determined and displayed by comparing the measured values of both measuring devices, characterized in that the comparison is carried out at a point in time at which the angles of inclination (φ, γ) of the missile with respect to its roll and pitch axes are equal to zero, and that this point in time with the aid of the second measuring device (S ₁ , B ₂ , B ₃ ) is determined in that the accelerations (Jb ₁ , b ₂ , b ₃ ) of the missile in the direction of the missile longitudinal axis, the missile transverse axis and the missile vertical axis are determined and it is determined when the expression (ft - sin α - b ₂ ) + (b ₃ - g) equals zero, where b = ] / 'b \ + b \, ^ ₁ the acceleration in the direction of the missile longitudinal axis, b ₂ the acceleration in the direction of the missile transverse axis, b ₃ the acceleration in the direction of the missile vertical axis, α the drift angle of the missile and g are the gravitational acceleration. 2. The method according to claim 1, characterized in that the accelerations (b _u b ₂ , b ₃ ) with the aid of individual accelerometers firmly connected to the missile cell (S ₁ , B ₂ , B ₃ ) and the drift angle (cc) with the aid of a Doppler radar speed measuring device (£>) can be determined. 3. Device for performing the method according to claims 1 to 2, characterized by a first accelerometer (B ₁ ) which measures the acceleration occurring in the direction of the missile longitudinal axis and whose signal output via a first isolating amplifier (F ₁ ), a first low-pass filter (T ₁ ) and a first quadrant (O ₁ ) is fed to one input of a first summing amplifier (V ₂ ) , through a second accelerometer (S ₂ ), which measures the acceleration occurring in the direction of the missile transverse axis and its signal output via a second isolating amplifier ( F ₃ ), a second low-pass filter (T ₂ ) and a second squarer (Q ₂ ) is fed to the other input of the first summing amplifier (V ₂ ) , through a square root (jR) whose input is connected to the output of the first summing amplifier (V ₂ ) and the output of which is connected to an input of a multiplier (M), the second input of which is supplied to the multiplicant sin λ, through a n second summing amplifier (V ₄ ), one input of which is connected to the output of the multiplier (M) and the other input of which is connected via an inverter (I) to the output of the second low-pass filter (T ₂ ), through a third accelerometer (B ₃ ) , which measures the acceleration occurring in the direction of the missile vertical axis and whose signal output via a third isolating amplifier (F ₃ ) and a third low-pass filter (T ₃ ) with one input of a Third summing amplifier (F ₆ ) is connected, the other input of which is supplied with a fixed voltage indicating the negative acceleration due to gravity (.—- g) as an addend, through a fourth summing amplifier (F ₇ ), one input of which is via a first circuit which forms the magnitude of a signal (S ₁ ) is connected to the output of the second summing amplifier (F ₄ ) and the other input of which is connected to the output of the third summing amplifier (F _{6 ) via a second circuit (S 2} ) which forms the amount of a signal, through one to the output of the fourth summing amplifier (F ₇ ) connected zero indicator (N ₁ ), which indicates the times at which the inclination angles (φ, γ) of the missile relative to its roll and pitch axes are simultaneously zero, further characterized by a Doppler radar speed measuring device ( D), which specifies the sine of the respective drift angle («), and by a comparison circuit (G, AT ₂ ) for qualitative and / or quantitative determination d er display error of the first measuring device (A). The invention relates to a method for monitoring a first, the flight position of a missile indicating measuring device on correct display with the help of an additional, from the first measuring device independently operating second measuring device, whereby by comparing the measured values of both "Measuring devices a possible deviation from the actual flight attitude is determined and displayed. For navigating modern aircraft, especially that of fighter planes, inertial measuring devices, so-called are increasing in recent times Inertial platforms, used that allow navigation of the aircraft without ground visibility, and also radio beacons Enable in-flight radar systems over longer periods of time. These inertial measuring devices have however, the disadvantage that they can be used even with the greatest possible effort are still afflicted with a certain error, the longer the operating time grows. Although it is known to support the inertial navigation system to carry out correction by a speed parade by the output signals of a Doppler radar, which is known to provide the ground speed, with the output signal the first integrator of the inertial navigation system are compared and a difference is attributed to the integrator. For a location correction, however, only those from the active or passive radio navigation used position values and / or astronomical position determinations, see TELDDC, paperback der Navigation, issue 1967, page 157. One such option, however, is a pilot not given during a mission flight over enemy territory. It is therefore not possible to date such a Pilots a possibly incorrect work of his carrier communication device so that he can still refer to the specified throwing today his inertial navigation device if they are already wrong.