EP4232771A1

EP4232771A1 - Interceptor missile and method for steering same

Info

Publication number: EP4232771A1
Application number: EP21791293.0A
Authority: EP
Inventors: Thomas Kuhn
Original assignee: Diehl Defence GmbH and Co KG
Current assignee: Diehl Defence GmbH and Co KG
Priority date: 2020-10-21
Filing date: 2021-10-08
Publication date: 2023-08-30
Anticipated expiration: 2041-10-08
Also published as: KR20230100722A; AU2021366148B2; WO2022084068A1; DE102020006465A1; AU2021366148A1; DE102020006465B4; IL301348A; AU2021366148A9; US20230251070A1; IL301348B2; EP4232771B1; IL301348B1; EP4232771C0

Abstract

In a method for steering a steerable interceptor missile (2) for intercepting a moving target (8), during a midcourse phase (PM) of the interception, which interceptor missile is driven by an engine (10), the interceptor missile (2) is steered with the aid of real steering commands, which are produced at respective steering times on the basis of free control parameters (SP), which free control parameters are in the form of a current parameter vector (10), the free control parameters (SP) are continually and repeatedly optimized during the midcourse phase (PM) with the aid of an optimization method (14) for optimizing the control parameters (SP), the optimization method (14) is carried out in parallel with the actual steering, newly learned information about the movement of the target (8) and/or information about the flight of the interceptor missile (2) is used in the optimization method (14) as soon as said information is available, and optimized control parameters (SP) are accepted into the current parameter vector (10) after said optimized control parameters are provided by the optimization method (14). An interceptor missile (2) contains the current parameter vector (10) and a control and evaluation unit (50) for carrying out the method according to the invention.

Description

Interceptor missile and method of guiding it

The invention relates to the guidance of an engine-powered guided missile interceptor for intercepting a moving target, particularly a target missile, during a midcourse phase of interception, and to such an interceptor missile.

An interceptor missile is launched to defend against a moving target, in particular an approaching target missile. After a launch phase, in which the interceptor missile leaves its launch base and begins its flight roughly in the direction of the target, the midcourse phase of its flight follows. This serves to cover most of the distance to the target and to get close to it, in particular so close that the interceptor missile's onboard systems are sufficient to be able to hit the target with pinpoint accuracy in an endgame that follows the midcourse phase.

DE 102010 032 281 A1 discloses a method for controlling a guided missile driven by an engine, in which a processing means of the guided missile calculates a trajectory property of a trajectory to a target point during flight and controls the flight of the guided missile depending on the trajectory property. When calculating the trajectory property, non-guided flight processes that influence the flight speed of the guided missile and are controlled by the process means are taken into account. Incorporation of future or present flight processes controlled by the processing means into flight control based on proportional navigation is possible, however complex. Such incorporation is easier when the processing agent uses miss-point navigation, specifically a technique called Zero Effort Miss (ZEM) navigation, rather than proportional navigation.

The object of the invention is to propose improvements with regard to an interceptor missile or the guidance of an interceptor missile in the mid-course phase of its flight to a moving target.

The object is achieved by a method according to patent claim 1 for guiding an engine-driven steerable interceptor missile. Preferred or advantageous embodiments of the invention, as well as other categories of the invention, result from the further claims of the following description and the attached figures.

The interceptor missile or its flight serves to intercept a moving target, in particular a target missile. The steering method can be referred to as model predictive guiding. It is performed during a midcourse phase of interception. The midcourse phase is the phase of the interceptor missile's flight from its launch to entry into the endgame. The endgame begins with the activation of the on-board target search sensor system (search head). During the mid-course phase, the target data originate in particular from the sensors of the higher-level weapon system and are transmitted to the interceptor via a data link. The desired successful impact on the target represents the end of the mission. Alternatively, the mission is aborted (end of mission) if the target cannot be reached or is finally missed or entry into the endgame is not possible or the interception is aborted or ended for other reasons . Then the proposed tax procedure also ends.

The interceptor missile is actually steered as follows: At the respective steering times, the interceptor missile generates real steering commands for itself based on the free control parameters currently present in the interceptor missile at the steering time, which are present in the form of a parameter vector. "Real" means that the interceptor missile actually uses the guidance commands generated in this way is steered. Optionally, additional values can also flow into the steering commands, for example (free) parameters that are not part of the parameter vector.

The method assumes that when the interceptor missile enters the midcourse phase, i.e. when it is already in flight, or from this point in time and expediently until the end of the midcourse phase or even beyond, a current parameter vector of free control parameters for the interceptor missile, from which the real steering commands are then generated. In particular, this parameter vector forms a suitable initial value for the steering, possibly also for an optimization, as explained below.

The free control parameters are continuously and/or repetitively optimized in the course of the midcourse phase using an optimization method for optimizing the control parameters. This optimization or the optimization method takes place parallel to the actual steering. This can be understood in such a way that the control parameters in the current parameter vector (basis of the real steering) initially remain unchanged. In this case, the parameters currently used for the steering are not optimized directly, but rather—figuratively speaking—a copy or an image of these control parameters or the parameter vector outside of the actual steering process. In this respect, the optimization can also take place independently of the actual steering, which initially does not have to be influenced by the optimization taking place in parallel therewith.

Newly detected information about the movement of the target and/or information about the flight of the interceptor missile is included in the optimization process as soon as it is available. The optimization method can therefore always be based on the most up-to-date and best available data about the circumstances of the current mission.

Optimized control parameters are then transferred to the current parameter vector after, in particular as soon as, they are available as a result of the optimization method. The optimized control parameters only influence the actual steering when the optimized parameters have been transferred to the current parameter vector on which the real steering is based. In the picture above, the optimized copy of the parameter vector is only then spoken integrated, transferred or taken over into the parameter vector actually used for steering.

The invention is based on the following core idea:

An interceptor missile is guided in the midcourse phase in order to hit a moving and, in particular, (potentially) maneuvering target. For this purpose, the real steering commands are calculated from a vector (parameter vector) of free (control) parameters. These free parameters can be target path angles, but they can also be target values for the lateral acceleration. In addition, in the case of a multi-stage engine, the ignition times of the respective stages can be treated as free parameters or, in the case of a controllable engine, its setpoint thrust or the time-discrete setpoint thrust profile. The invention is based on the idea that these free parameters are continuously and/or repetitively optimized over the course of the midcourse phase using a search method for parameter optimization. This optimization takes place independently and parallel to the actual steering. Whenever there is new information on the target movement or the flight history of the interceptor (interceptor missile), this is included in the optimization to determine better, ideally optimal parameters. The improved parameters are used by the actual steering as soon as they are available.

The following embodiments or variants of the invention are conceivable:

The optimization takes place in particular with regard to a target function (quality function/criterion/quality value), which in turn is based on a Zero Effort Miss (ZEM) prediction of the trajectories of the target and interceptor. As soon as the closest approach of target and interceptor (ZEM) is reached, the prediction (modified ZEM method) is aborted. The resulting trajectories are evaluated by the target function (quality value). For example, it is evaluated how close the interceptor comes to the target (ZEM), what speed the interceptor has at the end of its trajectory (maneuverability in the endgame), how long the fight lasts (Tgo), and at what angle the target and interceptor meet and any other sub-criteria. The target trajectory, ie the trajectory of the target, is predicted in particular on the basis of suitable hypotheses. For example, the target can be assumed to continue the current maneuver until a minimum approach speed is reached, in order to then continue flying in a straight line with maximum thrust (evasive maneuver). Or there is information on possible attack targets that suggest appropriate target maneuvers. Likewise, the target can be assumed to have a ballistic or pseudo-ballistic trajectory. The hypotheses regarding the target trajectory are based on prior knowledge and the observation of the target trajectory made up to the current point in time. There is extensive literature on this. Hypothesis formation and use is not a subject of this invention.

In the prediction of the interceptor trajectory, the free parameters to be optimized are used in particular, for example by converting the target trajectory angle or the target values for the lateral acceleration through a behavior model of the missile, in the case of a multi-stage engine the ignition times of the stages are selected accordingly and in the case of a controllable engine of the thrust or overrun is adjusted accordingly. In particular, the fuel consumption and loss of mass as well as the existing restrictions (minimum and maximum controllable thrust, no thrust after the fuel has been used up) are taken into account, as is the case with the well-known ZEM method, resistance and gravitation. In particular, an increment control ensures that events such as the ignition of an engine stage or reaching the ZEM are calculated precisely in terms of time.

Since the calculation of the target function each time includes the relatively complex step size-controlled simulation of the combat, it makes particular sense to use search methods that achieve the optimum or at least a significant improvement with relatively few iterations. Gradient-based methods are rather unsuitable for this due to the necessary approximation of the gradient using difference quotients. The simplex method according to Nelder Mead has proven to be very robust. This has been known in the prior art for about 60 years. According to the invention, improved steering of the interceptor missile towards the target is achieved due to the continuously updated control parameters and their use for real steering.

In a preferred embodiment of the method, the following optimization method is carried out, it being possible for the steps or the method to be ended or aborted at the end of the midcourse phase. The interceptor missile is then steered in the end game, which—like the start phase—is not part of the present patent application.

In a step a), a specifiable or specified parameter vector is selected as the current candidate of an MPC (Model Predictive Control) optimization method. The MPC method is used to potentially determine control parameters that are improved compared to the specified parameter vector. The current candidate thus forms a starting value for optimizing the control parameters using the MPC method.

In a step or method section b), a set of possible candidates (first and further subsequent candidates) for an improved parameter vector is determined as follows in or by carrying out the MPC optimization method; each of the candidates is assigned a quality value, which is also determined as part of the MPC process. The sentence can contain any number of candidates, whereby there can also be only one candidate which, for example, is always replaced when a better candidate is present. The number of candidates to be used is only a question of the chosen optimization method. Of course, powerful methods work with multiple candidates. For example, the widespread Nelder-Mead method operates with a simplex of n+1 parameter vectors, where n denotes the length of the parameter vector. But that plays no role for the idea of the MPG or the present invention. Even a "dumb" random search method of randomly varying the current parameter vector, evaluating the result, and continuing with the variation as the new current parameter vector in the event of an improvement would work. (Apart from the excessive computing time.) So the theorem contains in particular 1 to n candidates, where n depends on the optimization method, which is however not the subject of the invention.

Process section b) comprises steps c1) to c5):

In a step or method section c1), a modified ZEM method (Zero Effort Miss) is carried out as follows on the basis of the current candidate. The modified ZEM method includes steps d1) to d4):

In a step or method section d1), iterative predictions are made as follows at the respective step times; the method section d1) comprises the steps d2) to d4):

In a step d2), a possible intercept trajectory of the interceptor missile is predicted based on the current candidate, taking into account the guidance of the interceptor missile. The steering is based on virtual steering commands that are only generated as part of the optimization process, but are not used to actually steer the interceptor missile. Instead, the steering commands are used to virtually determine the predicted trajectory. However, the generation of the virtual steering commands can be identical to the generation of the real steering commands. This creates a realistic simulation of the trajectory.

In a step d3), a possible target trajectory of the target is predicted on the basis of hypothetical maneuvers of the target.

In a step or a loop d4), the steps d2) to d3) are repeated iteratively until a ZEM approximation of the intercept trajectory and the target trajectory is achieved. After the end of the loop, both trajectories (and possibly a corresponding remaining flight time, see below) are available until the ZEM is reached (i.e. the minimum distance between the trajectories, ideally zero if the interceptor missile can actually reach the target according to the prediction). Now proceed as follows with step c2):

In a step c2), based on the results of the ZEM method (the results are in particular: trajectories, ZEM, predicted duration Tgo des Flight along the trajectories until reaching the ZEM, etc.) a current quality value is determined using a quality criterion and assigned to the current candidate.

In a step c3), the current candidate is filed together as the first or further candidate in the set of candidates. The current quality value is assigned to the candidate as a quality value and is also stored in the sentence. When step c3) is reached for the first time, a first value pair consisting of a candidate and a quality value is stored, when it is reached again (see below) a second, then a third, etc., until the MPC method is completed and the set of candidates is thus available. For example, after ten runs of steps c1) to c4), there are ten candidates with their quality values.

In a step c4), the existence of an end criterion for the optimization or the MPC optimization method is checked. If this has not yet been achieved, i.e. the MPC optimization method has not yet completed its optimization of the current candidate or candidate in the sentence, the two steps e1) and e2) are carried out:

In step e1) the current candidate is varied towards a varied candidate using an M PC search method. A second candidate is created from the first candidate, a third from the second, etc. The search method is used for the numerical optimization of the free parameters.

In step e2), the varied candidate that has just been determined is then adopted as the current candidate and the process continues with step c1) or returns to it.

In a step c5), which represents the alternative to step c4), namely if the end criterion has been reached, the process continues with step f):

In a step f), a return is made to step a) and the method is continued there.

During the entire midcourse phase or the execution of the above-mentioned procedural steps (in a certain sense parallel to this) it becomes predeterminable Correction times selected according to a correction criterion one of the currently available candidates and replaced the current parameter vector with the selected candidate. As a result, optimized control parameters are adopted in the current parameter vector. From this point in time, the real steering commands can then be generated on a modified basis, namely on the basis of a changed parameter vector or one that has been improved in relation to the target, or optimized control parameters.

Optionally, in step c3) one or more of the candidates determined in the MPC method are discarded according to a rejection criterion and removed from the set together with their quality values. The sentence is kept correspondingly small and candidates that are not required are removed.

The "modification" of the ZEM method is that in a conventional or known or usual ZEM method, both a virtual guidance of the interceptor missile using a parameter vector in the form of the current candidate is taken into account, as well as hypothetical maneuvers for the target trajectory of the target .

The procedure begins after launch, i.e. when the interceptor missile is already in flight. At the moment the procedure begins, it can therefore be assumed that a current parameter vector is already available, which was used to steer the interceptor missile in the launch phase. The current parameter vector at the end of the start phase can therefore be selected in particular as a predeterminable parameter vector of the method.

The process can be aborted when the midcourse phase is over and the endgame control begins. Both the MPC method and the ZEM method are known in a wide variety of forms in the prior art, so that they are not explained in more detail here. Any forms of the respective known individual methods can be used and combined in embodiments of the invention. In particular, the prediction step width in the ZEM method is controlled according to known procedures, for example in such a way that the time steps become smaller as the target is approached. When "Hypothetical maneuvers" of the target are particularly suitable: unaccelerated movement (zero effort), ballistic trajectory, known or suspected evasive maneuvers or any other a priori knowledge of the target. When determining the trajectory of the interceptor missile and/or the target, passive effects such as non-controllable thrust, decreasing weight depending on fuel consumption, air resistance, etc. are taken into account in addition to the free parameters.

According to one embodiment of the invention, the Zero Effort-Miss steering (ZEM) is combined with the model predictive control approach (Model Predictive Control, MPC). The control parameters for designing the trajectory of the interceptor missile are permanently (time of correction) and online (i.e. during the course of the midcourse phase, by the interceptor missile itself) using prediction models for target (step d3), predicted course of the trajectory based on suspected maneuvers, etc. ) and interceptor missiles (interceptor, step d2, predicted trajectory based on guidance model, etc.).

The method can thus also be referred to as "Model Predictive Guidance (MPG)".

According to the method, the possibility arises of using an estimate of the target acceleration (hypothetical maneuvers) for the guidance of the interceptor missile. The proposed method forms a starting point for the guidance of a missile approaching over long distances with a controllable thrust profile (according to a free parameter), such as an interceptor missile based on a ramjet drive (ramjet interceptor).

According to one embodiment of the invention, a modified MPC is applied in the field of missile guidance. According to one embodiment of the invention, the combination of MPC and ZEM prediction modified and interacting in this way results in the guidance of an interceptor missile.

According to one embodiment of the invention, during the midcourse phase, in particular, two operations or processes run side by side or in parallel and to a certain extent independently of one another: The first process is the generation of Steering commands based on a currently (at the moment of generation of the steering command) present parameter vector. The second process is the optimization of the parameter vector. Using the MPC method, a set of possible alternative parameter vectors is generated and each of these parameter vectors is evaluated with a quality value. A modified ZEM prediction is used within the MPC method. Based on the second process, if necessary, namely if one is found, an optimized parameter vector (e.g. better quality value than the first parameter vector, which corresponds to the current one from the steering command generation) is selected and the current parameter vector in the first process is replaced by the optimized parameter vector. In the first process, the steering commands are then generated on the basis of the improved, replaced parameter vector.

In particular, both processes run independently of one another in that a certain number of steering commands are generated from one and the same parameter vector before the parameter vector from the second process is replaced at a later point in time. The reason for this is, for example, that the MPC method takes a certain amount of time before an improved parameter vector is found, but in the meantime steering commands continue to be generated at shorter time intervals.

In a preferred embodiment--as already explained above--the predeterminable parameter vector is specified in step a) by selecting the last current parameter vector of a start phase preceding the midcourse phase as the predeterminable parameter vector. In an alternative embodiment, a parameter vector is chosen that corresponds to the prediction of a direct approach to the target. In particular, two MPC assessments are carried out based on these two different first candidates, and that parameter vector which leads to the better quality value (quality, measure of quality) is selected as the first candidate. This ensures good starting conditions for the MPC process in the midcourse phase.

In a preferred embodiment, reaching the end criterion in step c5) is selected as the correction time and—as the correction criterion—that candidate selected from the set associated with the best quality value. Thus, the end of the optimization is awaited and only then is the parameter vector actually used for steering replaced. This new parameter vector is the best (best quality value) for route guidance that could be determined using the optimization.

In a preferred embodiment, step c3) is selected as the correction time and in step c3) the current candidate (which has just been or is currently being saved as a candidate in the sentence together with its quality value) is taken over as the current parameter vector, if its assigned quality value is the best of all quality values previously present in the set. Thus, the current parameter vector, ie the parameter vector used for the real generation of steering commands, is updated not only after the optimization method has been completed (step c5)), but already during its processing. Optimizations thus go earlier in the steering behavior and thus the trajectory of the interceptor missile.

This strategy of replacing the current parameter vector according to the alternatives mentioned can also be varied for different process runs of the MPC process.

In a preferred embodiment, in step e1) the variation towards a further or varied candidate is carried out at least partially on the basis of the previous candidates and their quality values. One or more or all of the candidates/quality values previously determined in the MPC method or set are therefore used in the search method in order to enable an improved determination of a next potential candidate.

In a preferred embodiment, the quality criterion contains at least as a sub-criterion: a minimum deviation from the target (ZEM, closest approach to/distance of the interceptor missile from the target) and/or a maximum terminal speed when it hits the target and/or a minimum remaining flight time to the target and/or or a desired angle of incidence on the target. The corresponding sub-criteria or their result values are assigned evaluation factors in particular in order to finally generate a quality value. All of these sub-criteria are to those that are ultimately decisive for a successful or even most effective approach to / combat of the goal.

In a preferred embodiment, the method is designed for an interceptor missile whose engine is a solid booster or a two-pulse engine or a controllable engine. In the case of a solids booster, in particular its remaining burning time for the midcourse phase is taken into account in step d2) in the method. In the case of a two-pulse engine, in particular its ignition time for the ignition of the second engine stage is taken into account as a free parameter in the parameter vector and, in particular, also optimized within the framework of the M PC method. In the case of a controllable (or regulatable) engine, e.g. a ramjet, in particular its thrust control value or the course of the thrust control value over time is taken into account as a free control parameter in the parameter vector and in particular optimized. For all three variants, the weight of the interceptor missile, which decreases with fuel consumption, is taken into account in particular in step d2).

In a preferred embodiment, a currently predicted remaining flight time of the interceptor missile until the end of its mission is additionally determined in step a). The end of the mission is in particular the meeting with the target or the reaching of a minimum distance to the target (ZEM). This remaining flight time (also "Tgo") can optionally be used as a free control parameter and/or as a sub-criterion for the quality criterion (e.g. the lowest possible remaining flight time) and/or for determining increments in the ZEM method. As part of the ZEM prediction, a current remaining flight time can then also be determined in each case, namely as the point in time at which the ZEM was reached.

In a preferred variant of this embodiment, the current parameter vector and thus in particular also the specifiable parameter vector and/or the current candidate etc. is one in which at least one of the free parameters is a value based on the remaining flight time or a sequence based on the remaining flight time of fractional values. A corresponding value is, for example, the above-mentioned ignition point for a two-pulse engine. For example, a sequence of partial values becomes one for the thrust control value (as a free parameter). controllable engine is selected as follows: the remaining flight time to the destination in an optimization process is divided into n, eg n=5, particularly equally long periods of time and each period of time is assigned a specific thrust control value as a constant partial value. In the optimization process, a thrust curve running in n or five steps (corresponding to the time segments) over time is used and optimized in step d2) for the prediction of the trajectory of the interceptor missile. In particular, with a variable thrust profile, a free control parameter is available so that the interceptor missile can react particularly well to highly agile evasive maneuvers by the target. For this variant of the method, the most up-to-date remaining flight time of the interceptor missile to the target is therefore always predicted.

In a preferred variant of this embodiment (in the alternative or variant of a sequence of partial values), the predicted remaining flight time is taken into account in step d2) in step d2) in such a way that it is divided into a specifiable number of time segments using the ZEM method is taken into account, and a different partial value is taken into account for each time period. As explained above, the parameter vector thus contains a free parameter, which in turn is formed from a value sequence of the partial values and represents, for example, a thrust curve in 5 stages/time segments.

In a preferred variant of this embodiment - as already explained above - the value or the partial values are dependent on the remaining flight time or determined by reaching a specific point in time or several ignition times of a respective first or further combustion stage of one or more engines of the interceptor missile. This embodiment is suitable for interceptor missiles which contain one or more single-stage or multi-stage engines, it being possible for such a stage of the engine to be assigned its own ignition point to be optimized.

In a preferred variant of this embodiment--as already explained above--at least one of the values or partial values is a thrust control value dependent on the remaining flight time for an engine of the interceptor missile whose thrust can be controlled. Here the dependency exists, for example, in an intermittent or continuous variation in thrust during the remaining flight time.

In a preferred variant of this embodiment, at least one of the partial values is a control value, dependent on the remaining flight time, for a transverse acceleration element of the interceptor missile. The activation of a corresponding lateral acceleration element leads to a lateral acceleration, i.e. a change in direction of the interceptor missile. The explanations given above for a controllable overrun curve apply analogously to a control of a corresponding transverse acceleration according to a time curve.

In a preferred embodiment, the current parameter vector (in particular also a specifiable candidate, see above) is selected that contains at least two path angles for the trajectory of the interceptor missile as two free parameters. This results in a particularly simple optimization problem for the M PC method. This also enables a particularly rapid reaction to highly agile targets, so that the interceptor missile can follow them particularly well.

The object of the invention is also achieved by an interceptor missile according to patent claim 15. During its flight, the interceptor missile is at least temporarily driven by its engine (at least one) and can be steered using real steering commands. The process can continue to be used even after all engines have burned out. For this purpose, it has a steering device, eg controllable rudders or lateral acceleration means, eg control nozzles, which are actuated on the basis of steering commands and are used to steer the flying interceptor missile. The interceptor missile is still used to intercept a target. The interceptor missile contains a respective current parameter vector of free control parameters for the interceptor missile, on the basis of which, as explained above, real steering commands for the steering are generated. The interceptor missile also contains a control and evaluation unit. The control and evaluation unit is set up to carry out the method according to the invention. The interceptor missile and at least some of its embodiments and the respective advantages have already been explained in connection with the method according to the invention.

The control and evaluation unit is set up or adapted or configured to carry out the method according to the invention. "Set up" / "adapted" / "configured" is to be understood in such a way that the control and evaluation unit is not only suitable for carrying out the relevant steps/functions, but rather was specifically designed for this. The control and evaluation unit is “set up” accordingly, in particular by programming a computing device contained therein or by hard wiring.

The invention is based on the following findings, observations and considerations and also has the following embodiments. The embodiments are sometimes also referred to as “the invention” for the sake of simplicity. The embodiments can also contain parts or combinations of the above-mentioned embodiments or correspond to them and/or optionally also include embodiments that have not been mentioned before.

New types of hypersonic weapons such as e.g. an HGV (hyperonic glide vehicle) or an HCM (hyperonic cruise missile) form a new threat as targets against which conventional interceptor missiles can hardly be successfully used.

The invention is based on the idea of using an interceptor missile, for example a so-called "Ramjet Interceptor" (RJI), namely a multi-stage missile based on a ramjet drive, against such hypersonic targets, ie against hypersonic weapons. The invention is also based on the idea of creating a guidance concept for the midcourse phase of such an interceptor missile. While there are various concepts for the endgame, which are mostly based on recognizing the target maneuver and switching directly to the guidance of the interceptor missile, the midcourse phase of existing concepts is limited to the best possible determination of the point of encounter (predicted intercept point = PIP) and the railway to this. For the new target class of potentially highly manoeuvrable Hypersonic Glide Vehicles (HGV) or Hypersonic Cruise Missiles (HCM), this is sufficient PIP approach is not sufficient, since the approach of the interceptor missile takes a comparatively long time and the target can deviate greatly from an originally suitable PIP during this time.

So far, steering in the midcourse phase has mostly been based on the PIP approach. This is determined with the help of all available a priori knowledge of the respective weapon system and the interceptor missile only has the task of approaching this PIP and ensuring a handover from the instructing sensor of the weapon system (radar) to the on-board sensor (seeker).

In a more flexible approach, the interceptor missile sets its own PIP. In any case, significant target maneuvers in the midcourse phase result in a shift of the PIP and thus a deviation from the original optimal trajectory. For this reason, new methods are based on classifying target maneuvers that are as insignificant as possible and avoiding an unnecessary relocation of the PIP and the associated loss of energy. A concept for the explicit treatment of maneuvering targets in the midcourse phase is the aim of the present invention. As mentioned above, there are a variety of well-known solutions regarding endgame-guidance (terminal guidance).

The MPC approach in embodiments of the present invention consists in not only predicting the target and interceptor trajectories (trajectories of target and interceptor missile) with a ZEM predictor, but also suitable control parameters using numerical, real-time capable search methods, e.g. in every nth steering cycle to optimize. The prediction of target and interceptor movement, especially based on the zero-effort approach, is a well-known concept. In predicting the target trajectory, the estimated target accelerations can be used using game theoretic hypotheses. In a conceivable realization, for example, the goal can be assumed to be to minimize the approach speed of the interceptor with the current maneuver (evasive maneuver). The method is based on the idea of free control parameters to be optimized. In a first approach, this can be the orbit angle of the interceptor at the current point in time. If the optimization has determined the optimal orbit angle, then you can this is interpreted as a target path angle (changed current parameter vector) and commanded in the form of a path steering (generation of the real steering commands).

The optimization can use a similar cost function (quality) as in an offline method (determination of the path by a guidance system outside of the interceptor missile). Criteria such as minimum departure, maximum terminal speed, minimum remaining flight time (time to go = Tgo) and geometric requirements such as specific impact angles in the cost function can be used. In further implementations, e.g. the ignition time of a second engine pulse can be used as a parameter to be optimized. Finally, it is possible to discretize the controllable thrust curve of a jet, ramjet, or gel engine in terms of time (sequence of partial values) and to optimally determine iteratively in the sense of the MPC.

Further features, effects and advantages of the invention result from the following description of a preferred exemplary embodiment of the invention and the attached figures. Show, each in a schematic principle sketch:

FIG. 1 shows a basic diagram of the method according to the invention.

The method is used to steer an interceptor missile 2, which is driven by an engine 4--here the thrust of which can be controlled--and can be steered, here by the control of tail units, not shown in detail. The steering takes place through the implementation of real steering commands 6 in the interceptor missile 2 to the engine 4 and the tail units, which is not explained in any more detail. The interceptor missile 2 is used to intercept a target 8. The method is carried out exclusively in a midcourse phase PM of the flight of the interceptor missile 2, i.e. intercepting the target 8.

The steering is based on a current parameter vector 10 . The parameter vector 10 contains a series, here three, of free control parameters SP1-3 for the interceptor missile 2. The control parameters SP1 and SP2 are path angles, the control parameter SP3 is a thrust control value for the engine 4, which includes a total of five sub-values SP3a-e. A respective remaining flight time Tgo of the interceptor missile 2 until hitting the target 8 is divided into five equally long periods of time. In each of these time segments, the engine 4 is controlled in sequence by a corresponding thrust control value SP3a-e.

At the beginning of the process, the launch phase of the interceptor missile 2 has just ended and the midcourse phase PM begins. A current parameter vector 10 is present when entering the midcourse phase PM. At the respective steering times, here every 10 ms, a respective real steering command 6 is generated from the parameter vector 10 and the interceptor missile 2 is steered using these steering commands 6 .

The method starts with a step a), in which a specifiable parameter vector 15 is selected as the current candidate 12 of an MPC optimization method 14 . In the present case, the specification is made in such a way that the current parameter vector 10 present from the end of the starting phase is used as the parameter vector 15 that can be specified. The MPC optimization method 14 serves to determine an improved parameter vector which is intended to replace the current parameter vector 10 .

The MPC optimization method 14 now begins. Within this method (step or loop b)), a set 16 of possible candidates 18a-c, three in this example, with quality values 20a-c assigned in each case, is determined. Each of the candidates 18a-c is a possible parameter vector that could replace the parameter vector 10 if it promised better mission success than the parameter vector 10 that is actually present at the moment.

Based on the current candidate 12, a modified ZEM method 22 is now carried out in a step c1): In a step or a loop d1), the following steps are carried out iteratively at the respective step times t1, 2, 3,...

In a step d2), a possible intercept trajectory 24 of the interceptor missile 2 is predicted. For this purpose, virtual steering commands 7 (corresponding to the real steering commands 6) are determined based on the current candidate 12 at the respective step times t1, 2, 3, . The trajectory 24 results from the temporal or spatial sequence of the locations. In other words, iteratively simulates how the Interceptor missile 2 would move if the current candidate 12 were used as the parameter vector 10 for its guidance.

Furthermore, in a step d3) corresponding to the step times t1, 2, 3, For example, it is assumed that the target 8 flies a specific avoidance curve to be assumed in order to escape the interceptor missile 2 .

According to a step or a loop d4), steps d2) and d3) are repeated iteratively for as many points in time t1, 2, 3, . This ends the ZEM method 22 .

In the example, the results 32 of the ZEM method 22 are the achievable ZEM approach 30, an updated remaining flight time Tgo, the impact speed and the impact angle of the interceptor missile 2 on the target 8, etc.

In a step c2), a current quality value 33 for the respective candidate 12 is determined on the basis of these results 32 and assigned to it. The assignment is based on a quality criterion 36.

In a step c3), the current candidate 12 is stored together with its ascertained good value 33 in the set 16 as a candidate 18a-c with a good value 20a-c. In the first run, the quality value 20a is assigned to the candidate 18a, in later runs the quality value 20b is assigned to the candidate 18b and stored in record 16, etc.

In a step c4), an end criterion 38 for the optimization method 14 is now checked. If this is not achieved, in a step e1) the current candidate 12 is varied using an MPC search method 40 to form a varied candidate 42 . In a step e2), this is accepted as the current candidate 12 and the MPC Optimization method 14 started again with the now optimized or modified candidate 12 .

In the example, the optimization method 14 is run through three times, resulting in three candidates 18a-c with associated quality values 20a-c. Then the end criterion 38 has been reached, namely in this case the fixed number of three method runs.

Since the end criterion 38 has been reached, a return is made to step a) in order to calculate a new set 16.

The process ends or is aborted when the midcourse phase PM has ended.

During the course of the method, one of the candidates 18a-c is selected according to a correction criterion 44 at a given correction point in time TK and is used from then on as the current parameter vector 10 for real guidance of the interceptor missile 2. In the example, the correction time TK is when the end criterion 38 is reached. The correction criterion 44 is the selection of that candidate 18a-c from the set 16 to which the best quality value 20a-c in the current set 16 is assigned.

An alternative possibility is to choose step c3) as the correction point in time TK and (from the second check/determination of the quality value) to make the best of the previously checked candidates 18a-c the parameter vector 10. The best one is when its quality value 20b-c is better than the quality values 20a-c of the candidates 18a-c previously present in set 16.

In step a), a currently predicted remaining flight time Tgo of the interceptor missile 2 to the target 8 is also determined in order to have a time base for the evaluation of the control parameters SP3a-e in step d2). An updated remaining flight time Tgo is also available as part of the result 32 at the end of a respective run of the ZEM method 22 and can be used from then on. The current parameter vector 10 is present in the interceptor missile 2 in each case. The interceptor missile 2 also contains a control and evaluation unit 50, here a central computer, which is set up to carry out the method according to the invention. The "establishment" takes place here by means of correspondingly powerful hardware and programming in order to implement the method.

List of reference symbols interceptor missile engine steering command (real)

7 steering command (virtual)

8 goal

10 parameter vector (current)

12 candidate (current)

14 MPC Optimization Methods

15 parameter vector (specifiable)

16 set

18a-c candidate

20a-c quality rating

22 ZEM method

24 Intercept Trajectory

26 flight maneuvers (hypothetical)

28 target trajectory

30 ZEM approach

32 results

33 quality value (current)

36 quality criterion

38 end criterion

40 MPC search method

42 candidate (varies)

44 correction criterion

50 control and evaluation unit

SP control parameters

Tgo remaining flight duration

PM Midcourse Phase t1 , 2, 3, ... Step time

TK correction time

Claims

- 24 - Claims

1. A method for guiding an engine (4) powered guided interceptor missile (2) to intercept a moving target (8) during a midcourse phase (PM) of interception, the interceptor missile (2) using real guidance commands (6 ) is steered, which are generated at the respective steering times using free control parameters (SP), which are present in the form of a current parameter vector (10), the free control parameters (SP) in the course of the midcourse phase (PM) using an optimization method (14) to optimize the control parameters (SP) are optimized continuously and/or repetitively, with the optimization method (14) taking place in parallel with the actual steering, with newly recognized information on the movement of the target (8) and/or information on the flight of the interceptor missile ( 2) be included in the optimization process (14) as soon as they are available, with optimized control parameters (SP) being incorporated into the current parameter vector (10). en after they are available from the optimization process (14).

2. The method according to claim 1, characterized in that the following optimization method (14) is carried out during the midcourse phase (PM): a) a specifiable parameter vector (15) is selected as the current candidate (12) of a model predicted control (MPC) optimization method (14) for determining improved control parameters (SP), b) in the M PC optimization method (14) a set ( 16) possible candidates (18a-c) for an improved parameter vector (10) with associated quality values (20a-c) determined as follows: c1) based on the current candidate (12), a modified Zero Effort Miss (ZEM) method (22 ) carried out as follows: d1) at the respective step times (t1 -3) are iteratively predicted as follows: d2) a possible intercept trajectory (24) of the interceptor missile (2) under

Consideration of virtual guidance commands (7) of the interceptor missile (2) based on the current candidate (12), d3) a possible target trajectory (28) of the target (8) based on hypothetical maneuvers (26) of the target (8), d4) the steps d2) to d3) are repeated iteratively until a ZEM approximation (30) of interception trajectory (24) and target trajectory (28) is reached, c2) based on the results (32) of the ZEM method (22) becomes a current quality value (33) determined using a quality criterion (36) and assigned to the current candidate (12), c3) the current candidate (12) is in the sentence (16) successively as the first (18a) or further candidate (18b-c) together with the current quality value (33) is stored as an assigned quality value (20a-c), c4) if an end criterion (38) of the optimization has not yet been reached: e1) the current candidate (12) becomes a varied candidate (42), e2) the varied candidate (42) becomes the current candidate from now on (12) is accepted and the process continues with step c1), c5) if the end criterion (38) is reached, the process continues as follows: f) a return is made to step a), with during the midcourse phase (PM) at predeterminable correction times (TK) one of the candidates (18a-c) is selected according to a correction criterion (44) and the current parameter vector (10) is replaced by the selected candidate (18a-c). is, in order to thereby take over optimized control parameters (SP) in the current parameter vector (10).

3. The method according to claim 2, characterized in that reaching the end criterion (38) in step c5) is selected as the correction time (TK) and that candidate (18a-c) from the set (16) is selected as the correction criterion (44). , to which the best quality value (20a-c) is assigned.

4. The method according to claim 2 or 3, characterized in that step c3) is selected as the correction time (TK) and in step c3) the current candidate (12) just stored as the candidate (18a-c) is additionally used as the correction criterion (44). is accepted as the current parameter vector (10) if its associated quality value (20a-c) is the best of all quality values (20a-c) previously present in the set (16).

5. The method as claimed in one of claims 2 to 4, characterized in that in step e1) the variation towards a varied candidate (42) is carried out at least partially on the basis of the previous candidates (18a-c) and their quality values (20a-c). .

6. The method according to any one of claims 2 to 5, characterized in that the quality criterion (36) contains at least as a sub-criterion: a minimum deviation from the target (8), a maximum final speed when hitting the target (8), a minimum remaining flight time to Target (8), a desired angle of impact on the target (8).

7. The method according to any one of claims 2 to 6, characterized in that in step a) a currently predicted remaining flight time (Tgo) of the interceptor missile (2) is determined until its mission end. - 27 -

8. The method according to claim 7, characterized in that the current parameter vector (10) used is one in which at least one of the free parameters (SP) is a value based on the remaining flight time (Tgo) or a sequence of partial values based on the remaining flight time (SP3a-e) is.

9. The method according to claim 8, characterized in that for the variant of a sequence of partial values (SP3a-e): in step d2) the predicted remaining flight time (Tgo) is taken into account in such a way that this in the ZEM method (22) in a predetermined number of time segments is divided, and for each time segment a different partial value (SP3a-e) is taken into account

10. The method according to any one of claims 7 to 9, characterized in that at least one of the values or partial values (SP3a-e) is an ignition time dependent on the remaining flight time (Tgo) of a respective first or further combustion stage of one or more engines (4) of the interceptor missile (2) is.

11. The method according to any one of claims 7 to 10, characterized in that at least one of the values or partial values (SP3a-e) is a thrust control value dependent on the remaining flight time (Tgo) for an engine (4) of the interceptor missile (2) that can be controlled with regard to its thrust power. is.

12. Method according to one of claims 7 to 11, characterized in that at least one of the partial values (SP3a-e) is a control value dependent on the remaining flight time (Tgo) for a lateral acceleration element of the interceptor missile (2).

13. The method according to any one of claims 2 to 12, characterized in that - 28 - one is selected as the current parameter vector (10) which contains at least two path angles for the trajectory of the interceptor missile (2) as two free parameters (SP1.2).

14. The method according to any one of the preceding claims, characterized in that the method for an interceptor missile (2) is designed, the engine (4) is a solid booster or a two-pulse engine or a controllable engine.

15. Interceptor missile (2), which is driven by an engine (4) and can be steered using real steering commands (6) and is used to intercept a target (8) that has a current parameter vector (10) of free control parameters (SP) for contains the interceptor missile (2) and which contains a control and evaluation unit (50) which is set up to carry out the method according to one of Claims 1 to 14.