FR2723191A1 - METHOD AND DEVICE FOR SELF-GUIDING A MILITARY HEAD MISSILE - Google Patents

Abstract

The missile autoguiding method according to the invention consists in that the missile (1), during its phase of approaching the objective, follows a curved trajectory on which the ratio of the signals reflected by the terrestrial objective the ground echo clutter is kept substantially constant and in that the trajectory vector of the missile is directed towards the ground. The autoguiding device comprises a group of antennas formed by a transmitting antenna (s) and receiving antennas (a, b, c, d) which is integral with the missile and which is connected to a transmitting and receiving unit or one or several projectors and detectors consisting of a transmitting and receiving unit with optical operation integral with the missile, the detectors being connected to an operating unit.

Description

The present invention relates to a method for homing a missile

  steerable and equipped with a military head as well as a device for the implementation

of this process.

  Devices of this type are used in particular in rockets, guided projectiles, small missiles carried and launched by another missile (submissiles), artillery ammunition and proximity weapons and in particular

in the missile field.

  In the field of current missiles, there are known methods of homing a guided missile provided with a military head which may be a projectile or a missile carried by another missile (submissile) and provided with an active search head. In this case, the projectile or submissile is used for the fight

  against resting or moving earthly targets.

  The object of the invention is to provide a method in which, during its approach phase to the objective, the missile uses for its guidance the background image of the ground (clutter). The invention also aims to provide a device by which it is possible to guide the missile to a terrestrial objective and which allows with inexpensive means to carry out a scanning or exploration of the

ground and earth goal.

  As regards the method, this object is achieved according to the invention by the fact that during its approach phase of the objective the missile follows a curved trajectory, that on this curved trajectory the S / C ratio of signals S reflected by the terrestrial objective at the ground echo C is kept constant or at least

  substantially constant and that the trajectory vector of the missile is directed towards the ground.

  As regards the device, this object is achieved according to the invention thanks to a group of antennas consisting of transmitting antennas and receiving antennas which constitute a fixed group connected to the missile and which is connected to a transmitting and receiving unit, or by means of a transmitting and receiving unit constituted by one or more projectors and by detectors and functioning optically, which constitutes a fixed group connected to the missile in which the detectors are connected to a

operating unit.

  The advantages of the method include in particular that the objective signal and the background image signal (echo signal) are converted simultaneously

  into a guidance signal with the greatest Doppler slip.

  The advantages of the device are that the scanning or exploration of the ground and the terrestrial objective is accomplished without the need for

  move parts mechanically.

  This allows inexpensive manufacture of the device according to the invention. It is also possible to use inexpensive manufacturing technologies such as those used for the lower GHz range for producing the device. In addition, the device resists strong accelerations and bad weather. Other characteristics and advantages of the invention will appear better

  in the detailed description which follows, which relates to the case of a missile and which is

  refers to the accompanying drawings, given solely by way of example, and in which: FIG. 1 represents the missile during its phase of approaching the objective; FIG. 2 represents the missile during its phase of search for the objective; FIG. 3 represents the concept of system of the method according to the invention which involves an integrated FLSAR method; FIG. 4 represents the decomposition into components of the movement of a mobile terrestrial objective; Figure 5 shows part of the g / lq plane of Figure 3; FIG. 6 represents a possible scenario for approaching the objective seen from the missile; FIG. 7 represents a modular block diagram of a preferred embodiment of the device according to the invention; and Figure 8 shows a preferred antenna configuration for the

  device according to the invention which is shown in FIG. 7.

  FIG. 1 represents the missile 1, which can for example be a guided rocket, during its phase of approaching the objective. The trajectory vector 21 of

  missile 1 is turned towards the ground 4 from the longitudinal axis of the missile. The extension

  ment of the trajectory vector 21 in the direction of the ground 4 meets the latter at a point Z which, at the instant t = to, takes the value Z0 on the abscissa t of a system of

  Cartesian coordinates of ordinate ii, which defines the plane of the trajectory (t, n).

  to represents an instant preferably earlier during the objective approach phase. At the latest at this instant to, during the approach of the objective, the missile 1 continuously transmits by a transmitting antenna (not shown) signals which are reflected on the one hand by the terrestrial objective 3 and d on the other hand by the ground 4 proper. These reflected signals S (by the terrestrial objective 3) and C (by the ground 4) are received by the receiving antennas (not shown) of the missile 1 and operated in an operating unit (not shown). Depending on the results of this operation, the missile 1 is guided so that its trajectory 2 is curved so that on average the ratio between the signal S reflected by the terrestrial objective 3, also called the terrestrial objective signal, and the signal C reflected by the ground 4, also called the clutter or background image signal, is kept constant or at least substantially constant. In the case of a terrestrial objective 3

  fixes the trajectory 2 is then substantially circular.

  The S / C ratio between the signal from the terrestrial objective S and the clutter signal is given approximately by the following relation: S s C = const. - R (1.1) C Rz c Z 2f sint 2f VG in which S = signal reflected by the terrestrial objective 3, RZ = distance between the missile 1 and the terrestrial objective 3 ("distance from the objective"), y = angle between the distance RZ of the missile 1 to the terrestrial objective 3 and the distance RD between the missile 1 and the meeting point D of the proper speed vector VG of the missile 1 ("angle of the lines of sight"), c = speed light, fs = emission frequency of the missile 1 radar, B = operating bandwidth of the missile 1 radar,

VG = proper speed of missile 1.

  When the terrestrial objective 3 moves, this movement is detected by the missile 1 which carries out a corresponding correction of its trajectory. This correction can be carried out in that the missile 1 is guided towards an estimated point of impact which is distant by a distance AZ from the momentary position of the objective (momentary stopping position of the terrestrial objective 3). The estimation of this distance can be done for example according to the formula AZ = TI VZ in which T * = estimated instant of impact, VZ = speed of the terrestrial objective 3 which moves on the ground 4, AZ = distance traversed by the terrestrial objective 3 in the plane ,, q (see

figure 5).

  To determine this ratio and the transverse acceleration necessary for guiding missile 1, it determines the distance RZ between missile 1 and the terrestrial objective 3, the distance RD between missile 1 and the meeting point D on the ground 4 and the angle of the sight lines y between the two distances RZ and RD, by

  example using the FLSAR process (orward-Looking-, ynthetic-Aperture-

  Radar). The natural speed of the terrestrial objective 3 is determined for example from the ratio between the radial speed of the terrestrial objective 3 and the natural speed of the missile 1 multiplied by the cosine of the angle of the sight lines y, and this proper speed is taken into account for the correction of the trajectory. The correction of the trajectory can be calculated for example online or else it can be read

in a memorized table.

  The transverse acceleration bq which must be communicated to missile 1 is described approximately by the following relation: 2VG sin b = GO (1.2) zo in which VG = natural speed of missile 1, yo = angle of the lines of sight at the instant t = to of the acquisition of the objective,

  Rzo = distance from the objective at the instant t = to of the acquisition of the objective.

  The angle of the sight lines yo can be determined for example by

  selection of the Doppler frequency.

  FIG. 2 represents the missile 1 during its search phase for

the objective.

  All the objective points whose sighting directions make the same angle with the eigenvector vector VG of the missile 1 give echo signals of the same Doppler slip compared to the emission frequency of the radar. The curves formed by the intersection of these sighting directions with the ground 4 can be described by constant Doppler sliding curves also called "iso-Doppler". These curves are substantially elliptical and, for a constant frequency difference, the distances between these ellipses always become more

  small when the angle of the sight lines increases.

  The same distance curves, also called 'iso-

  Distances "('' Iso-Ranges"), are circles centered on the projection of the missile i on the ground 4, which, when observed at a sufficiently small solid angle, are separated by substantially constant distances. The lobe of radiation from the radar of the missile 1 radar cuts a specific part of this

family of curves.

  The FLSAR method consists in obtaining, by means of a narrow band exploitation of the echo signals of each distance domain, a resolution

  angular by Doppler effect rather than by the antenna radiation lobe.

  The resolution cells are therefore determined by the intersection of the "iso-ranges" and the "iso-Doppler" and not by the intersection of the "isoranges" and

of the antenna radiation lobe.

  The use of the FLSAR method in the context of the invention makes it possible to obtain essential advantages over the concepts of current millimeter wave radars. These advantages are particularly apparent when comparing the FLSAR method used within the framework of the invention and the current millimeter wave radar method which is used for transverse resolution.

of the antenna radiation lobe.

  The angular resolution which can be obtained with an operating bandwidth B increases with the deviation from the natural velocity vector VG. Even when using an emission frequency situated for example in the Ku band, the most inaccurate "Doppler cell" which extends directly around the point of encounter of the speed vector with the ground 4 (surface of the earth) is not wider than the angular resolution of a radar finder head of the same distance having an antenna radiation lobe width which, in the case of an opening available in a missile, could only be obtained of

  emission frequencies located in the millimeter wave range.

  Unlike such a radar which, for the search for the objective, should accomplish a sweeping movement in the given domain, a FISAR head requires a significantly longer integration time (inverse of the Doppler operating bandwidth) , duration which it has however of the fact that it

  embraces the entire objective domain simultaneously.

  Although with a millimeter wave radar it is possible to improve the resolution outside the center of the research area by Doppler-Beam-Sharpening (DBS) (the same resolution could be obtained at the higher frequency even with a duration proportionately shorter integration), it is however still necessary to explore sequentially the

  research area, mechanically or by means of a "Steered-" antenna

  Phase-Array "which, however, is not yet feasible at present in the

millimeter wave domain.

  As a first approximation, there is no appreciable difference between the FISAR process and a millimeter wave radar for the detection or classification of the objective when the angular resolution and the resolution of distance are substantially identical in the two cases, because the mode of formation of the resolution cells has no influence on the choice of the criteria and on the classifiers. The additional exploitation of the echoes as a function of the co-polarization and the cross-polarization is also possible in the FLSAR method thanks to a corresponding construction of the antenna system. It exists

  however important differences with regard to the guidance towards the objective.

  In order to be able to follow the objective and determine the speed of rotation of the lines of sight which constitutes a measure of the necessary guidance orders, the current millimeter wave radar must constantly be directed towards the terrestrial objective 4 and be decoupled from all the movements of the missile 1, for example by a gimbal frame. For this purpose it is also necessary to have a

  reference system which can be produced for example in the form of a platform

  gyroscopic form. Regarding the guidance of the missile we usually use-

  the proportional navigation process.

  The gimbal frame and the gyroscopic platform constitute demanding and therefore costly fine mechanical organs which appreciably increase the cost, in particular the cost of the search head and which, concerning the possibility of producing in particular artillery ammunition guided by terminal phase, should be considered a factor of increased risk due to

  severe mechanical stresses occurring during firing.

  On the contrary, an antenna integral with the missile and with a large width of radiation lobe, an antenna which will be discussed more precisely in the following, makes it possible to give up not only the gimbal but also the systems of

  reference as will appear below.

  In the approach phase to land objectives 3, the trajectory which would be given by a proportional navigation guidance method would not be optimal for the missile 1 according to the invention because, when the speed of the land objective 3 is low at negligible compared to VG's own speed, the proportional navigation process gives rise to a trajectory which leads in a straight line, that is to say with an angle of sight lines close to zero, to the objective

terrestrial 3.

  The angle of sight lines offers, when acquiring the objective, a

  determined angular resolution which, together with the distance resolution

  ment, provides a sufficient S / C signal / clutter ratio. If the angular resolution decreases faster than the distance to the objective, the S / C ratio will deteriorate correspondingly, which could lead to a loss of

  the objective and an interruption in the pursuit of the objective.

  It is therefore provided according to a preferred embodiment of the method according to the invention that the missile 1 travels on a trajectory along which the S / C ratio is kept as constant as possible throughout the flight. A constant S / C ratio means that during the guidance phase the angle of sight lines must not decrease more than in proportion to the distance

of the objective.

  A circular trajectory to which the natural velocity vector VG is tangent at the time of detection and which passes through the position of the objective constitutes a very good approximation of this trajectory. Thus, in this method, the missile 1 is not guided according to the principle of the proportional navigation method over a constant angle of sight lines but over a radius of

constant trajectory.

  This is made possible by the fact that, depending on the distance from the objective and the line of sight angel at the time of detection, a constant line of sight angle for S / C is given beforehand. as a setpoint and is compared with the actual line of sight angle between the direction of

  the objective and the eigenvector vector VG. Guidance is accomplished through

  diary of a regulator according to the deviations which result from this comparison.

  For the values which are taken into account for the opening angle of the antenna diagram of the antenna integral with the missile and which are for example between +/- 10 'and +/- 20', it is that is to say for initial deviations from the objective of the same order of magnitude or less, the course of the circular trajectory described results, in relation to the rectilinear trajectory resulting from the current proportional navigation method, by lengthening the duration flying only fractions of a second, which is insignificant given the speeds

  maximum possible for terrestrial objectives.

  Consequently, when for example: -the angles of incidence necessary to maneuver missile 1 are less than half the half-width of the antenna radiation lobe, -the approach path is so steep that the meeting point of the clean velocity vector with the ground 4 is still located in the range of the radar, the missile 1 has 4 receiving channels which can be combined according to a monopulse process in two planes in an azimuth difference signal, in a signal d 'elevation and in a sum signal, -the receiver channels have characteristics so stable that they make it possible to have in both planes a characteristic' error voltage as a function of the reproducible angular difference ', the missile 1 simultaneously has the following information simultaneously: 1. its own speed in the form of maximum measurable Doppler effect, 2. the distance to the ground in the form of distance to the point of maximum Doppler effect, 3. the speed re relative to the objective in the form of a Doppler effect of the objective, 4. the distance from the objective, and 5. the direction of the line of sight relative to the objective by comparison with the trajectory vector of the monopulse characteristic, 6. the component of the target's own speed in the direction of the lines of sight as the difference between the relative speed in the direction of the target and the missile's own speed multiplied by the cosine of l 'angle

lines of sight.

  Thanks to these values it is possible to directly measure the angle of sight lines between the direction of the objective and the proper speed, independently

  the angle of incidence of the projectile. Therefore, when the conditions mentioned above

  above are met, it is possible to forgo a separate reference system

  intended for the memorization of the trajectory vector.

  Figure 3 shows an example of a process system concept

  according to the invention in which the FISAR process is integrated.

  The bearing surfaces and possibly the control surface having been deployed, the missile 1 is placed on a trajectory with a pointing angle of approximately 45 ". Therefore, a righting maneuver to obtain a flight

horizontal is not necessary.

  The radar provided in missile 1 makes it possible to control this trajectory.

  The search for the objectives is done for example from an altitude of approximately 1500 m with a width of radiation lobe of the antenna of approximately

' for example.

  With an emission frequency of 18 GHz for example (which corresponds to a wavelength of 16.7 mm) it is possible to obtain this lobe width of 3dB as a product of the emission and

  reception with individual antennas of around 40 mm in diameter.

  From a geometric point of view, the situation as it is represented

in Figure 3 is as follows.

  In a system of terrestrial coordinates x, y, p, which is defined by the projection of the momentary position of the missile on the ground 4 (surface of the earth) and by the meeting point D of the eigen velocity vector VG, whose value results from the distance RD between the missile 1 and the meeting point D, the plane of the trajectory (tri) is defined by its abscissa t and its ordinate i. In this trajectory plane it is possible to draw a circular trajectory which is tangent to the trajectory vector and which meets the objective point Z insofar as this vector

  trajectory coincides with the eigenvector vector VG.

  As shown in Figure 3, there are also the following relationships between the distances RD, RW, RZ measured on the side of missile 1 and the solid angles ct, A, y and between the distances XZ, XD, with respect to the earth , between the momentary projection of the missile on the ground 4 and points Z, D. The missile is then at

the altitude PG-

  RD represents the distance between missile 1 and the meeting point D of the eigenspeed vector VG, RW represents the distance between missile 1 and a meeting point W, RZ represents the distance between missile 1 and point Z of the goal. The meeting point W is a point on the ground 4 which is defined by the extension of the longitudinal axis of the missile and which coincides with the trajectory vector 21 of FIG. 1. The solid angle ct is located between the line segment (GW) missile 1 / meeting point W and the line segment (GZ) missile 1 / meeting point Z. The solid angle f is located between the line segment (GD) missile 1 / meeting point D and the line segment (GW) missile 1 / meeting point W. In the system of line segments above, G denotes a point

missile 1.

  The solid angle y is the angle of sight lines which is decisive for

the Doppler slip.

  The scalar products of the distance vectors GD. GZ, GW. GZ and GW. GD give, at the time of the detection of the objective, the components being represented in the usual way: GD = xD; 0; (PG3.0.1) GZ = xz; Y; PG (3.0.2) GW = xz; 0; -PG (3.0-3) cose = XD xz + P G1) cosd = Z DP G (3-1) Rz 'RD

2 2

x Z + PG cosO = (3-.2)

RZ RD

x. x À2P cosfl, 'G (3-.3)

R. RD

  For the plane angle e and the momentary components above we obtain the following relations:

Y

  Z> XD = arctan Y (3.4.) ÀjX9 X1 = -4 (3.4.2) 'IXz = XD2 XZ <' D = arctan +) (3.4.3) IXz <XD Xz - XD and for the solid angle t we have: = m + y (3.5) The solid angles t, o and O are located respectively between the

  line segments GD and DU, GZ and DZ, and DZ and DW.

  it These segments are limited by the points G, D; OF; G, Z; D, Z; and D, W. The points G, U, D, Z are located in the plane of the trajectory, (, i). U is the origin of the imaginary two-dimensional Cartesian coordinate system of the

trajectory plane, (g,).

  The line segment GU has a length i1G, the line segment UD has a length gD and the line segment UZ has a

length gz.

  Furthermore, the solid angles E and e are located between the segments (O, O, O) D and DG and between the segments DW and DZ. (O, O, O) is the origin of the Cartesian terrestrial coordinate system which contains the components x, y, p.

  Figure 4 represents the decomposition into components of the movement

  ment of the terrestrial objective 3 according to figure 3 insofar as it moves. The reference elements x, y, g, U, D and O of Figure 4 are those of Figure 3 which

  are necessary to represent the movement of the Earth's objective.

  If it is assumed that the terrestrial objective 3 moves in a spatial direction DZ, the speed of the mobile terrestrial objective 3 is broken down into

components in missile 1.

  A component DIZ, which is located in the plane of the trajectory (, q), contributes to the relative speed in the direction of the objective and thus causes a difference between the product of the natural speed of missile 1 and cozy and the

  measured value of the relative speed in the direction of the objective.

  The angle between the direction of the objective and the ground 4 can be obtained from the distance RZ between the missile 1 and the objective 3 according to FIG. 3, the distance RD between the meeting point D and l line of sight angle y so that it is possible to also obtain the lens speed component in

the plane of the trajectory (t, q).

  Another component DQZ is located perpendicular to the plane of the trajectory and appears in the missile 1 in the form of speed of rotation of the plane

  of the trajectory in missile 1.

  From the measured values described above, it is possible to determine unequivocally in the missile 1 the amplitude and the direction of the maneuver acceleration necessary to obtain the correct circular trajectory. As Figure 4 shows, the following relationships exist between the different components: Dz À / DX2 2 (4.1) DZ = / 2r z xz YZ DZ = D cos (If-,) (4.2) DQ = D sin (' z - ') (4.3)

QZ Z

  with Dyz = arctan D (4.4) DXz the general motion vector DZ of the terrestrial objective 3 decomposes into a component Dz located in the plane of the trajectory and into a component DOZ perpendicular thereto. The DZ component generates a radial component in the direction of missile 1, which makes it possible to measure it. The component DQZ causes a rotation of the plane of the trajectory around the own vector VG VG. qp is solid angle which results from the decomposition into components of the motion vector Dz, a component Dyz being parallel to the axis y and a component DXZ being parallel to the axis x of the coordinate system x, y. FIG. 5 also represents part of the plane of the trajectory (nWi) of FIG. 3. The reference elements of this figure also come from FIG. 3. These are the reference elements 1, A, 1, VG, 3, RD, RZ, G, U, D, Z, oD and The reference element AZ is identical to that which has been described in connection with FIG. 1. As shown in FIG. 5, the missile 1 describes substantially an arc of a tangent to the initial eigenvector vector, represented here in the form of a vector VG shifted in parallel, and which passes through the objective point Z. As an example, we will now consider the case where the speed VZ of the terrestrial objective 3 is equal to zero at the objective point Z. It follows that: Rz 1 sin o = 2 '- (5. 1.0) Rz r0 RZ (5.1.1) - 2s in 7o O y0 and RZ0 represent the values measured for y and RZ at the instant t = to = 0. The transverse guide acceleration bq which must be communicated to missile 1 obeys the rela tion: b = (5.2.0) q ro or, taking equation (5.1.1): b = G sin0 (5.2.1) q R O

ZO

Ro

  and gives, when bq is constant, the circular trajectory which must be followed.

  ro is the radius r at the instant t = to = 0 with the center Mo of the trajectory of the missile 1. The center M0 is offset by a distance Ar on the segment GM0 or GM, the point M being located between the points G and M0 of the GM0 segment. M is located at the point of intersection of the segment GM0 and a line which intersects the distance RZ at right angles and which presents substantially the minimum distance to the objective Z. The transverse resolution Q in the range of a detected terrestrial objective is therefore given by: Rz Q = a1 1 (5.1) V G'- s in c B with: RZ = direct distance between the missile 1 and the terrestrial objective 3 at the objective point Z, fs = frequency of the signals d 'emission, c = speed of light, VG = natural speed of missile 1, B = operating bandwidth of radar of missile 1,

y = line of sight angle.

  nl results for the resolution of the objective at the time t = to = 0 at which the terrestrial objective 3 is identified by the missile 1: Rz 2f 1 Cs. V (5.2) sin ro c B the index 0 added to the variables of equation (5.2) indicates that these are the variables mentioned above at time t = t0 = 0. For example, yo represents the angle

  line of sight at time t = 0.

  From 5.2 it follows in general: R (t) X (t) 'si (t) -t) = (5-3) r0 and from Rz (t) = r0 = Constant (54) s in (t ) it results c (t) = CO (5.5) The variables that appear in equations (5.3) to (5.5) are identical to the variables mentioned above. In other words, the above result means that C and therefore S / C remain constant over the entire trajectory in

  measurement o missile 1 describes a circular trajectory.

  Figure 6 shows an example of a possible approach scenario

objective seen from the missile.

  In the context of this scenario, suppose that the objective is acquired, that the objective moves and that a component of the speed of the objective is located in the plane of the trajectory, that the eigenvector vector of the missile is turned in the direction of its meeting point and a guiding maneuver is initiated so that, momentarily, the main axis of the missile is no longer turned in the direction of the meeting point. On the basis of this scenario, we recognize in the figure a Cartesian coordinate system specific to the missile comprising a vertical axis Ver. and a horizontal axis Hor. as well as an axis FKH perpendicular to the previous ones and which preferably coincides with the main axis of the missile. The natural velocity vector V G is directed towards its meeting point D around which the constant Doppler sliding curves are represented. The meeting point D and the objective Z mentioned above are therefore located on the axis t of the plane of the

  trajectory (z, È) according to Figures 1 and 3 to 5, plane of which only the axis t is shown here.

  When the objective Z moves, a component of the speed of the objective is located in the plane of the trajectory and then contributes to the relative speed in the direction of the objective, as also shown in FIG. 4 for example. If we now transpose this representation scenario to the generally known monopulse process (identification of a representation mode for example for the exploitation of radar reception signals, which should not be confused with the monopulse process belonging to the field of radar signal production), the representation described in FIG. 6 is enriched by two coordinate systems into which a horizontal or vertical error signal is introduced which is a function of the horizontal or vertical deviation in this two-dimensional Cartesian coordinate system, the objective Z and the meeting point D being represented in the two coordinate systems in a known manner in the form of points Hor.Z and Hor.D or Ver.Z and Ver.D, as this is shown in Figure 6 for the scenario represented there, in which case, when the approach to the objective continues to unfold, the points Hor.z, Hor.D, Ver.Z and Ver.D may see their posit modified ion and be located at the origin of the coordinate systems cited in the event of direct alignment of the main axis of the FKH missile with the Z objective (towards the

  end of the objective approach phase).

  FIG. 7 shows a preferred embodiment of the device according to the invention for implementing the method according to the invention. This device consists of a radar target finder head, also called a radar finder head. This contains a group of antennas 77 which consists of a transmitting antenna s and four receiving antennas a-d. The transmitting antenna s and the receiving antennas ad are substantially of the same size and are mounted in the missile 1 as shown for example in FIG. 7. Another embodiment (not shown) consists in grouping the receiving antennas ad around the transmitter antenna outside the missile 1. The choice of the transmission frequency is not subject in principle to any limitation, but it is advantageous that the transmission frequency is located in the lower part of the gigahertz domain ,

for example at around 18 GHz.

  The group of antennas 77 is connected to a transmitting and receiving unit 79.

  This transmitting and receiving unit 79 is constituted by other elements 78, by an objective identification logic 7, by a geometry calculator 8 and by a regulator 9. The different antennas s and ad are connected to the logic d objective 7 identification via the elements 78. From the electrical point of view, this objective identification logic 7 is followed by the

  geometry 8 and regulator 9 for guiding missile 1.

  The transmitting antenna is controlled by a control unit 64 by means of a modulator 63, of a control stage 62 connected in

  downstream and a power stage 61.

  The control unit 64 also controls the identification logic.

tion of objective 7.

  The receiving antennas ad are each connected to the objective identification logic 7 via an amplifier 51, a filter 52, an adjustable amplifier 53, an A / D converter 54, d '' a series of distance filters, a series of speed filters, preferably made in the form of

  fast Fourier transform, possibly with focusing, to compensate

  the non-linearities of the echo frequencies, in the R-FFT 55 elements and

V-FFT 56.

  As shown in this figure, the elements mentioned above are connected in series. The function of the device according to the invention is therefore as follows: The transmission power is produced in two stages and is radiated by the transmitting antenna s. The modulator 63, controlled by the control unit 64 of the search head of the missile 1, then produces a frequency modulation in

sawtooth shape.

  This radar operates according to the FM-CW process. It is also possible

  instead of using a radar using a pulse method

  with or without additional modulation to increase the distance resolution

  is lying. In the FM-CW method, the spectrum of the echo signals is significantly narrower than in a pulse method, which allows homodyne mixing, and the digitization of the echo signals can take place very far forward in the signal train. The four parallel analog circuits

  each consist only of a receiving mixer (possibly preceded

  using an HF preamplifier), a filter whose damping pattern is adapted to the proportionality between the echo frequency and the distance damping (SRC filter also called sensitive-range-control filter) and by

  example in an amplifier with automatic gain control.

  The simplicity and narrowness of the strip facilitate compliance with the conditions of stability and uniformity of the signal circuits and make it possible to carry out the formation of the sum and difference diagrams only in the geometry calculator 8 instead of carrying them out. in a comparator arranged in the HF plane, the choice of the correct coefficients allowing also to develop non-symmetrical diagrams comprising zero positions in a direction

wanted ("Adaptive-Null-Stecring").

  This provides a significant improvement in resistance to

  radar finder disturbance.

  In the objective identification logic 7 the sum signals are formed for each resolution cell and the objective identification and objective classification algorithms are used for the groups of cells adapted to the objective type. Algorithms for "Adaptive-Null-Steering"

  possibly incorporated can also be integrated into this element.

  After acquisition of the objective, the signals available in the radar finder head (for example own speed, distance from the ground, radial speed in the direction of the objective, distance from the objective, deviation from azimuth and elevation of

  the objective and meeting points) are transmitted to the geometry calculator 8.

  During the pursuit of the objective, the objective identification logic 7 operates in parallel so that the correct resolution cells can be exploited

as a goal.

  The geometry calculator 8 corresponds to the autopilot for common missile search heads. The geometry calculator 8 converts the measured data into control signals for the guidance regulator 9 into

  depending on the guidance process to be implemented.

  We will describe in the following an example of a particularly favorable mission sequence for the missile: - launching and ballistic flight in the area of the objective, - reconfiguration of the missile into a guided missile, pivoting in a trajectory suitable for research, - search for the objective, identification of the objective and acquisition of the objective,

- guidance in the objective.

  Launching and ballistic flight concerns the search head of the missile insofar as it must withstand the extreme mechanical and thermal stresses which then appear. To this end, it is necessary to take appropriate measures with regard to the construction, measures which, as already indicated, are however appreciably reduced thanks to the abandonment of the common fine mechanical organs which are the cardan and the gyroscopic reference. The guided missile reconfiguration includes the release of the radome which must be adequately protected during storage, launch and ballistic flight against mechanical and thermal deterioration, as well as the deployment of the bearing surfaces and the guiding devices (control surfaces) . The conditions concerning the possibility of maneuvering in the case of a missile according to the invention are less severe than in the case of a homing missile in the millimeter wave domain, given that - the condition concerning a research flight horizontal disappears, - the detection ranges are greater so that there is a longer time for maneuvers aimed at reaching objectives located at the edge

of the field of vision.

  Pivoting in a trajectory suitable for research means the course, from the ballistic curve, of a rectilinear trajectory forming a pointing angle of about 45 'relative to the surface of the earth. As soon as a sufficiently low altitude (for example between 2000 and 3000 m depending on the reflectivity of the ground) is reached, the search head of the missile can acquire the ground and determine the angle of the trajectory with the axes of the missile. From the distribution of the distances over a determined angle of line of sight (Doppler effect) it is possible to determine and correct the aiming angle: the curve of a Doppler effect determined on the ground is an ellipse along which distance varies more or less strongly depending on the pointing angle. In the distance-speed diagram (R-V diagram) this ellipse takes the form of a line of distance cells shifted down by the factor

  cozy compared to the maximum speed.

  Because of the long integration time which is necessary for Doppler analysis, the search for the objective, the identification of the objective and the acquisition of the objective require proper motion compensation because for a period of time integration of 64 ms for example, the projectile typically moves from 12 to 19 m, that is to say a multiple of the depth of the remote cells. This self-motion compensation is accomplished, in simplified representation, by the modulation of the signals received by a "chirp" signal (linear frequency modulation signal) which opposes the modulation.

caused by own movement.

  The methods of objective identification and rejection of false objective which are already used in the millimeter wave search head of the missile can be used here also in a similar manner since the dimensions of the

  resolution cells are comparable.

  FIG. 8 shows by way of example a possible configuration of the group of antennas 77 which consists of four receiving antennas a-d and one transmitting antenna s and which has a total diameter of approximately 120 mm for example. The Fraunhofer area diagrams of the different antennas are not offset from each other and the angular error signals (offset) are obtained on the basis of the phase differences of the signals (single phase pulses). At the low emission frequency fs and the fact that the diagrams do not rotate relative to the axes of the missile, the structure of the radome can be appreciably better adapted to aerodynamic constraints than in the case of a

  current millimeter wave radar finder head.

  Because the different receiving antennas ad are arranged in missile 1, the azimuth and elevation are given by adding the reception signals in the following ways: Azimuth = a + c - b - d (7.1) a + c + b + d Elevation = a + b - c - d (7.2) a + b + c + d In the case of mobile objectives, it is necessary to make a correction of the flight time and therefore of the trajectory. For this purpose, as shown

  In Figure 1, the flight time must be assessed according to the mobile objective.

  The duration of flight as a function of a mobile objective can be evaluated for example in the following way: the estimated relative speed between the missile and the mobile objective is: V * = VG 'cozy * + V cosô * (8.1) r Z in which VG represents the speed of the missile 1 measured for example in the direction of the meeting point D according to FIG. 3. y is the angle of sight lines measured and o is the solid angle according to FIG. 3. VZ is the lens speed

  terrestrial 3 which moves on the ground according to figure 1 and figure 5.

  The estimated values of the different variables that are not estimated

  in the case of a stationary terrestrial objective are provided with an asterisk.

  Given Figure 5, it follows from equation (8.1) that: v * - v cos * V r G (8.2) Z Posai - "'C OSCtO * From the measured values or the estimated values RZ *, RD, y * it is possible to calculate a trajectory directed towards Z *. This is characterized by y *, t *, o) *, t *, (t = derived from t with respect to T *) and by the internal relation between y and t according to equation (3.5) which, in the case of the movement of the terrestrial objective, is transformed into an equation of the form, * = to * + y *. Concerning this trajectory, the speed of the objective can be broken down into a tangential component VtZ

  located on the trajectory and in a radial component VZ.

  Vtz = Vz. cos (o * - y *) (8.3) To this tangential component corresponds an angular velocity) * (around M *): * = V- = tZ (8. 4) tï r * the estimated time is given by:

T *. + 2) = 2 (8.5)

  where: 2_ z (8.6) TF v = if4 - V The position estimated at the instant dc impact is then shifted by AZ = T * .VZ according to equation (8.6). For the target position, 0 * is calculated again in a later step and is used as the setpoint. This setpoint is corrected by AZ so that the line of sight is not directed towards the point of impact but towards the terrestrial objective 3. The deviation of lines of sight Ay which then results from the movement of the terrestrial objective 3 over the distance AZ according to FIG. 5 results from:

2 2 2

  FZ = RD + R - 2RD Rz - cos'a (8.7) o 2 2 dZ2 r = arcos D + RZ (8.8) 2RDRZ This line deviation dc aimed Ay can be used instead of the variation of distance AZ for trajectory correction in the case of a moving earth objective 3. Flight time or trajectory correction can take place

  constantly throughout the duration of the target approach phase.

  It will be understood that the invention is not limited to the examples of

  embodiment described but that it can be applied to many other examples.

  It is thus for example that it is possible to use as emission signals of the missile, instead of the microwave signals described, optical signals from the UV to IR domains, for example the signals from a laser. In this case, of course, the spectral sensitivity of the receiving antennas must be adapted to the emission spectrum.

  of the transmitting antenna s. This can be achieved by using, for example, photo-

  transistors as receiving antennas.

  It should be noted that the terrestrial objective or objective on the ground can be an objective located on water (for example a ship) so that, in this case, the term ground

  in the sense of the invention denotes water.

Claims (20)

  1. Self-guiding method of a guided missile provided with a military head, in which it is a projectile or a submissile provided with an active search head which emits signals, the projectile or the submissile being used for the fight against stationary or mobile terrestrial objectives, characterized in that - during its phase of approach of the objective the missile (1) follows a trajectory (2) curved, - on this trajectory (2) curved the S / C ratio of the signals S reflected by the terrestrial objective (3) to the echo C of the ground (4) is kept constant or at least substantially constant and - the trajectory vector (21) of the missile (1) is directed towards the ground (4) during the objective approach phase.   2. Method according to claim 1, characterized in that, on the side of the missile (1), the distance RZ to the objective point Z, the distance RD to the meeting point D of the natural speed vector VG of the missile (1) with the ground (4) and the angle of sight lines y between the two distances RZ and RD are determined by means   the FLSAR (Eorward-Looking-ynthetic-Aperture-Badar) process.   3. Method according to any one of claims 1 and 2, characterized   in that the line of sight angle there is measured directly by the monopulse method.   4. Method according to any one of claims 1 to 3, characterized   in that the missile (1) describes a trajectory (2) which is substantially circular   in the case of a stationary terrestrial objective (3).   5. Method according to claim 4, characterized in that the S / C ratio is given approximately by the following relation: s, _s S const. - S C RZ C   ZB v sin '2fs VG in which S = signal reflected by the terrestrial objective (3), RZ = distance between the missile (1) and the terrestrial objective (3) ("distance from the objective"), y = angle between the distance RZ between the missile (1) and the terrestrial objective (3) and the distance RD between the missile (1) and the meeting point D of the eigenvector vector VG of the missile (1) ("angle of lines "), c = speed of light, fs = frequency of emission of the missile radar (1), B = operating bandwidth of the missile radar (1), VG = proper speed of the missile (1).   6. Proc6d6 according to any one of claims 1 to 3, characterized   in that a trajectory correction is accomplished in the case of a target terrestrial (3) mobile.   7. Method according to claim 6, characterized in that the trajectory correction is accomplished in such a way that the missile (1) is guided on an estimated point of impact which is distant from the momentary position of the objective of a distance AZ and in that the estimation of this distance AZ is accomplished according to the following formula: AZ = Vz. rT * in which rT * = estimated impact instant, VZ = speed of the terrestrial objective (3) which moves on the ground (4),   AZ = distance traveled by the terrestrial objective (3) in the plane (,).   8. Method according to any one of claims 6 and 7, characterized   in that the speed of the terrestrial objective (3) is determined from the ratio of the radial speed of the terrestrial objective (3) to the natural speed VG of the missile (1) multiplied by the cosine of the angle of lines of sight y and is taken into account for the trajectory correction.   9. Method according to any one of claims 6 to 8, characterized   in that the trajectory correction is calculated online or is read from a table memorized.   10. Method according to any one of the preceding claims,   characterized in that the transverse acceleration bq which must be applied to the missile (1) is given by the following relation: 2V2 sin0 b = G 0 q R Zou in which VG = natural speed of the missile (1), Y = angle of lines of sight at the time of target acquisition,   RZo = distance from the objective at the time of acquisition of the objective.   11. Method according to any one of the preceding claims,   characterized in that the trajectory correction is based on the modification of   the angle of sight lines Ay in the case where the objective on the ground (3) is mobile.   12. Device for implementing the method according to one   any one of the preceding claims, characterized in that a group   antenna (77) formed by transmitting antennas and receiving antennas constitutes a fixed group connected to the missile (1) which is connected to a transmitting and receiving unit (79) or consists of a transmitting and receiving unit constituted by one or more projectors and by detectors, which operates optically and which constitutes a fixed group connected to the missile, and in that the detectors are connected to an operating unit.   13. Device according to claim 12, characterized in that the group of antennas (77) consists of a transmitting antenna (s) and at least three, preferably four, receiving antennas (a-d) which are preferably of size substantially identical.   14. Device according to any one of claims 12 and 13,   characterized in that the receiving antennas (a-d) are grouped around   the transmitting antenna (s) in or on the missile (1).   15. device according to any one of claims 12 to 14,   characterized in that the group of antennas (77) has radiation lobes wide antenna.   16. Device according to any one of claims 12 to 15,   characterized in that the transmitting and receiving unit (79) contains an objective identification logic (7) and in that a geometry calculator (8) and a regulator (9) for guiding the missile (1 ) are connected downstream of this logic objective identification (7).   17. device according to any one of claims 12 to 16,   characterized in that the transmitting and receiving unit (79) contains a control unit (64) and in that the transmitting antenna (s) is controlled by this control unit (64) via a modulator (63), a control stage (62) connected downstream from the previous one and a power stage (61). 18. Device according to claim 17, characterized in that the control unit (64), in which monopulse signals are preferably generated   also, controls the objective identification logic (7).   19. Device according to any one of claims 12 to 16,   characterized in that the receiving antennas (ad) are each connected to the objective identification logic (7) via an amplifier (51), a filter (52), an adjustable amplifier (53), an A / D converter (54), a   R-FFT filter (55) and a V-FFT filter (56).   20. Device according to claim 12, characterized in that it is   provided one or more lasers acting as projectors.