Inertial sensor system for missiles and methods for flight-phase-dependent inertial sensor measurement

Abstract

Inertial sensor system (IMU) for use in an inertial navigation system of a missile (30), comprising: a first inertial sensor module (B1) comprising at least two first inertial sensors (C1, ..., Ci) of a first inertial sensor type and a first, multiplexer (MUX1) coupled with the outputs of the first inertial sensors (C1, ..., Ci); a second inertial sensor module (B2) comprising at least one second inertial sensor (D1, ..., Dj) of a second inertial sensor type different from the first inertial sensor type; and a control processor (P) which is coupled to the first multiplexer (MUX1) and which is designed to control the first multiplexer (MUX1) to output the sensor measurement values ​​of an inertial sensor selected by the control processor (P) from the first inertial sensors (C1, ..., Ci), wherein the control processor (P) is further designed to select an inertial sensor of the first inertial sensors (C1, ..., Ci) and an inertial sensor of the second inertial sensors (D1, ..., Dj) depending on an instantaneous flight phase of the missile (30).

Description

TECHNICAL AREA OF INVENTION

[0001] The invention relates to an inertial sensor system for missiles, a missile with such an inertial sensor system, and a method for flight-phase-dependent inertial sensor measurement during the flight of a missile. TECHNICAL BACKGROUND

[0002] Light missiles, such as guided missiles or guided grenades, are often equipped with an inertial navigation or control system (inertial navigation system). Due to the increased susceptibility to errors of inertial navigation systems, such light missiles frequently integrate additional sensors. The signals from these sensors provide further information about the missile's position and serve to detect and, if necessary, correct errors in the output values ​​of the inertial navigation system. To optimize the control of light missiles, all sensor signals can be validated using Kalman filters, which can incorporate dynamic measurement data such as altimeter data, data from a global navigation satellite system, or magnetometer data as additional parameters.

[0003] Often, conflicting requirements arise, since on the one hand, high-quality control functions with a sufficient safety requirement level according to IEC 61508 / IEC61511 with sufficient measuring range, good error tolerance and high bandwidth are desirable for light aircraft, while on the other hand, the inertial sensor systems to be used should also be cost-efficient.

[0004] Document US 7,171,303 B1 discloses a navigation system with an IMU for the synchronous provision of GPS measurements and inertial measurements.

[0005] Document US 8 239 162 B2 describes an inertial sensor system with a base having several physically separate sectors on which three groups of orthogonally aligned angular velocity sensors are positioned, with each group being positioned on a different sector of the base.

[0006] For example, in the publication DE 10 2011 115 971 A1, a method for determining inertial measured quantities using an inertial measuring unit comprising inertial sensors is proposed, in which at least two inertial sensors of different detection accuracy each detect the same measured quantity independently of each other in only partially disjoint measured quantity value ranges and combine the detected measured quantity values ​​to form an error-corrected measurement signal.

[0007] The publication EP 2 352 968 A1 discloses an inertial measuring unit with sensor clusters consisting of several cost-effective MEMS inertial sensors of low measuring accuracy.

[0008] The publication EP 2 594 891 B1 describes a method for flight-phase-dependent inertial sensor measurement during the flight of a body. SUMMARY OF THE INVENTION

[0009] One of the objectives of the invention is to find solutions for accurate and long-term stable inertial sensor systems for use in missiles that are easy and inexpensive to implement.

[0010] These and other problems are solved by an inertial sensor system with the features of claim 1, a missile with the features of claim 7, and a method for flight-phase-dependent inertial sensor measurement in a missile with the features of claim 8.

[0011] According to a first aspect of the invention, an inertial sensor system for use in an inertial navigation system of a missile comprises a first inertial sensor module, a second inertial sensor module, and a control processor. The first inertial sensor module has at least two first inertial sensors of a first inertial sensor type and a first multiplexer coupled to the outputs of the first inertial sensors. The second inertial sensor module has at least one second inertial sensor of a second inertial sensor type, different from the first inertial sensor type.The control processor is coupled to the first multiplexer and the second multiplexer, and is designed to control the first multiplexer to output the sensor readings of an inertial sensor selected by the control processor from the first inertial sensors, and to control the second multiplexer to output the sensor readings of an inertial sensor selected by the control processor from the second inertial sensors.

[0012] According to a second aspect of the invention, a method for flight-phase-dependent inertial sensor measurement comprises the steps of determining a current flight phase of the aircraft, selecting a first active inertial sensor from at least two first inertial sensors of a first inertial sensor type depending on the determined flight phase, selecting a second active inertial sensor from at least two second inertial sensors of a second inertial sensor type different from the first inertial sensor type depending on the determined flight phase, and processing the sensor measurements of the two active inertial sensors in a control processor.

[0013] According to a third aspect of the invention, a missile comprises an inertial sensor system according to the first aspect of the invention.

[0014] A key aspect of the invention is to provide two or more redundant sensors for some or all sensor types of an inertial sensor system. These redundant sensors must have different measurement characteristics, such as different bandwidths, vibration resistances, acceleration resistances, measurement accuracies, temperature dependencies, scale factor errors, long-term stability, drift biases, or similar properties. By providing multiple sensors of each sensor type, a central control unit of the inertial sensor system can select a set of sensors as the temporarily optimal one for each flight phase and use the measured values ​​of the sensors in this selected set as the basis for the measurement.This makes it possible to respond to the respective requirement profiles of the different flight phases, where the respective sensor characteristics contribute to the reliability of the overall measurement to varying degrees.

[0015] A particular advantage of the solutions according to the invention is that the costs and required installation volume of such inertial sensor systems can be significantly reduced, since the sensors used only need to be optimized with respect to a smaller selection of sensor properties. Large, heavy, and expensive sensors that must simultaneously meet all conflicting requirements, such as a wide measuring range and high measuring accuracy, can therefore be advantageously avoided.

[0016] Advantageous designs and further developments result from the additional sub-claims as well as from the description with reference to the figures.

[0017] According to some embodiments of the inertial sensor system according to the invention, the inertial sensor system comprises a memory coupled to the control processor in which configuration data sets for error compensation of each of the first and second inertial sensors are stored. In some embodiments, the control processor can further be configured to select one inertial sensor of the first inertial sensors and one inertial sensor of the second inertial sensors depending on the configuration data sets stored in the memory.

[0018] According to the invention, in the inertial sensor system, the control processor is further configured to select one inertial sensor from the first set of inertial sensors and one inertial sensor from the second set of inertial sensors depending on the current flight phase of the missile. In some embodiments, it may be possible for the control processor to determine the current flight phase of the missile based on instantaneous sensor readings from at least one inertial sensor from the first and second sets of inertial sensors.

[0019] According to some further embodiments of the inertial sensor system according to the invention, the first type of inertial sensor can comprise accelerometers. According to some further embodiments of the inertial sensor system according to the invention, the second type of inertial sensor can comprise gyroscopes.

[0020] According to some further embodiments of the inertial sensor system according to the invention, the second inertial sensor module can have at least two second inertial sensors and a second multiplexer coupled to the outputs of the second inertial sensors. The control processor can furthermore be configured to control the second multiplexer to output the sensor measurement values ​​of an inertial sensor of the second inertial sensors selected by the control processor.

[0021] According to some embodiments of the method according to the invention, the first inertial sensor type comprises acceleration sensors. According to some embodiments of the method according to the invention, the second inertial sensor type comprises gyroscopes.

[0022] According to some further embodiments of the method according to the invention, the processing of the sensor measurements includes calculating the current flight attitude of the missile based on acceleration sensor measurements and yaw rate sensor measurements.

[0023] According to some further embodiments of the method according to the invention, the method further comprises the step of compensating the error parameters of the two active inertial sensors depending on configuration data sets for each of the first and second inertial sensors stored in a memory coupled with the control processor.

[0024] According to some further embodiments of the method according to the invention, the method further comprises the steps of determining, by the control processor, whether the missile enters a next flight phase, selecting a new first active inertial sensor from at least two first inertial sensors of a first inertial sensor type depending on the next flight phase, selecting a new second active inertial sensor from at least two second inertial sensors of a second inertial sensor type different from the first inertial sensor type depending on the next flight phase, and processing the sensor measurements of the two new active inertial sensors in the control processor.

[0025] The above embodiments and further developments can be combined with one another as appropriate. Further possible embodiments, further developments, and implementations of the invention also include combinations of features of the invention described previously or subsequently with regard to the exemplary embodiments, even if not explicitly mentioned. In particular, the person skilled in the art will also add individual aspects as improvements or additions to the respective basic form of the present invention. BRIEF SUMMARY OF THE CHARACTERS

[0026] The present invention will be explained in more detail below with reference to the exemplary embodiments shown in the schematic figures. These figures show: Fig. 1 an exemplary illustration of an inertial sensor system according to an embodiment of the invention; Fig. 2 an exemplary illustration of an inertial sensor system according to a further embodiment of the invention; Fig. 3 a schematic block diagram of an inertial sensor system according to a further embodiment of the invention; Fig. 4 a schematic block diagram of an inertial sensor system according to a further embodiment of the invention; Fig. 5 an exemplary illustration of a missile with an inertial sensor system according to a further embodiment of the invention; Fig. 6 an exemplary illustration of a missile with an inertial sensor system according to a further embodiment of the invention; Fig. 7 a flowchart of a method for flight-phase-dependent inertial sensor measurement on a flying object according to a further embodiment of the invention; and Fig. 8 a flowchart which shows exemplary details of the procedure of Fig. 7 represents.

[0027] The accompanying figures are intended to provide a further understanding of the embodiments of the invention. They illustrate embodiments and, in conjunction with the description, serve to explain the principles and concepts of the invention. Other embodiments and many of the aforementioned advantages become apparent with reference to the drawings. The elements of the drawings are not necessarily shown to scale. Directional terminology such as "above," "below," "left," "right," "over," "below," "horizontal," "vertical," "front," "back," and similar terms are used for explanatory purposes only and are not intended to limit the general public to specific embodiments as shown in the figures.

[0028] In the figures of the drawing, identical, functionally equivalent and similarly acting elements, features and components - unless otherwise stated - are each provided with the same reference symbols. DESCRIPTION OF EXAMPLES OF EXECUTION

[0029] Missiles within the meaning of the present invention comprise all ballistic or unguided missiles and guided missiles that can move along definable flight paths in and outside of airspace. Missiles within the meaning of the present invention particularly include light drones, guided missiles, guided grenades, all types of cruise missiles, surface-to-air missiles, air-to-surface missiles, anti-tank guided missiles, anti-ship missiles, air-to-air missiles such as air-to-air missiles or surface-to-air missiles, anti-rocket missiles, and anti-satellite missiles.

[0030] An inertial measurement unit (IMU) within the meaning of the present invention is a spatial combination of several inertial sensors, such as accelerometers, gyroscopes, and similar sensors. IMUs are sensory measuring devices of inertial navigation systems (INS), which are used, among other things, in aircraft for flight navigation and for the control-based stabilization of the aircraft in space. To detect six possible kinematic degrees of freedom, an IMU typically has at least three mutually orthogonal accelerometers (translation sensors) for detecting translational motion in the x, y, and z directions, and at least three mutually orthogonal gyroscopes (rotational rate sensors) for detecting rotational or gyroscopic motions about the x, y, and z axes.IMUs can therefore provide at least three linear acceleration values ​​for translational motion and at least three angular velocity values ​​for rotation rates. In an INS, the linear velocity along the trajectory and the position in space relative to a reference point can be determined from the linear acceleration values, possibly after compensation for gravitational acceleration. Integrating the three angular velocity values, relative to a reference point, yields the orientation in space.

[0031] Fig. Figure 1 shows an exemplary illustration of an inertial sensor system 10. The in Fig. The components of the inertial sensor system 10 shown in Figure 1 are only exemplary, and further components can be implemented in connection with the inertial sensor system 10. For example, it may be possible for the inertial sensor system 10 to have a similar or analogous function to that shown in Figure 1. Fig. The inertial sensor system IMU, shown in block diagram 3 and functionally explained below, is implemented. The inertial sensor system 10 can, for example, be part of an inertial navigation system (INS), which is designed, for example, for use in a missile. Such a missile 30, which can have an inertial sensor system 10 (or inertial sensor system IMU), is exemplified in Fig. Figure 5 illustrates this. The inertial sensor system 10 can be configured, in particular, as a strapdown inertial sensor system, i.e., an IMU in which the inertial sensors of the inertial sensor system 10 are rigidly connected to the outer frame of the missile 30. In particular, the accelerometers of such strapdown inertial sensor systems also exhibit a sensor measurement component attributable to the rotational motion of the missile 30, which can be compensated for using sensor signal values ​​from the gyroscopes to determine the translational acceleration values.

[0032] The inertial sensor system 10 comprises a carrier platform 18a on which a circuit board 18b is mounted. The carrier platform 18a serves to mechanically stabilize the components on the circuit board 18b and can have, on the one hand, laterally arranged mounting tabs 17 for attaching the inertial sensor system 10 to a frame of a missile, and on the other hand, mounting pins 19a on the underside of the carrier platform 18a. The mounting pins 19a can be used to stack several inertial sensor systems 10 on top of each other by inserting the mounting pins 19a into identically arranged mounting openings 19b in the circuit board 18b of another inertial sensor system 10. The mounting pins 19a serve both as a mechanically stable connection and as spacers between the two stacked inertial sensor systems 10. The number of mounting pins 19a and / or mounting openings 19b can vary, even if in Fig. 1 Four mounting openings 19b are shown as examples. Likewise, the positioning of the mounting openings 19b and the mounting pins 19a in relation to the plan view of the essentially planar inertial sensor system 10 is not limited to the figures shown in Figure 1. Fig. The positioning shown is limited to the outer corners. Furthermore, it may be possible to use floor plan shapes other than the circular one. Fig. 1. To select an exemplary shape for the mounting openings 19b and the mounting pins 19a, for example polygonal plan or cross-sectional shapes.

[0033] The inertial sensor system 10 has a central control processor 12a, which is supplied with electrical power via a power supply module 11. The central control processor 12a can be, for example, an ASIC, an FPGA, a PLD, a digital signal processor, a microcontroller, or another suitable portable computing device capable of processing incoming sensor signals and converting processed sensor signals into output values ​​that can be output at input / output interfaces 14a of the inertial sensor system 10. In addition to the input / output interfaces 14a, the inertial sensor system 10 can also have system-connecting interfaces 13, via which several stacked inertial sensor systems 10 or their control processors 12a can communicate with each other and exchange data. The central control processor 12a can also have internal memory.It may also be possible that a separate (in . Fig. 1 memory module (not explicitly shown) is provided on board 18b, which is coupled to the central control processor 12a, and from which the central control processor 12a can retrieve stored data for processing.

[0034] The inertial sensor system 10 comprises a number of inertial sensors of various types. An inertial sensor type is a group of sensors designed to measure a specific inertial sensor parameter, such as linear accelerometers, rotational accelerometers, gyroscopes, and similar sensor types. The physical design of an inertial sensor is irrelevant for its classification as a sensor type, provided the sensor can output the measured values ​​specified by that sensor type for the given parameter. Inertial sensors within a group assigned to a particular sensor type can differ in their measurement characteristics. For example, inertial sensors of the same sensor type may have different bandwidths, vibration resistances, and other characteristics.

[0035] Inertial sensors may exhibit differences in acceleration stability, measurement accuracies, temperature dependencies, drift (bias), long-term stability, or scale factor error. These measurement characteristics can be determined through calibration procedures, such as using rotary tables, centrifuges, or temperature chambers, depending on the stimulation parameters, and stored as error parameters. Compensation parameters for controlling the inertial sensors can be calculated for each determined measurement characteristic. These parameters allow the output values ​​of the respective inertial sensors to be compensated before further processing in a control processor. The compensation parameters for each inertial sensor used can be stored in a configuration data set in a memory location of the inertial sensor system 10.

[0036] The inertial sensor system 10 can, for example, comprise three spatially oriented gyroscopic sensors or rotation rate sensors 15a, 15b, 15c, each fixedly arranged orthogonally to one another on the circuit board 18b. The inertial sensor system 10 can, for example, additionally comprise a three-dimensional gyroscopic sensor 15d. The inertial sensor system 10 can, for example, comprise three spatially oriented linear accelerometers 16a, 16b, 16c, each fixedly arranged orthogonally to one another on the circuit board 18b. In some cases, redundant accelerometers with the same orientation can be used for one or more of the linear accelerometers 16a, 16b, 16c, as in Fig. Figure 1 shows an example of the linear accelerometer 16a. Furthermore, the inertial sensor system 10 can, for example, additionally include a three-dimensional linear accelerometer 16d. It should be clear that the number, orientation, relative positioning on the circuit board 18b, and dimensions of the inertial sensors 15a, 15b, 15c, 15d, 16a, 16b, 16c, and 16d shown in Figure 1 are described in Figure 1. Fig. 1 are only exemplary and that a multitude of other implementation options can also be realized.

[0037] Fig. Figure 2 shows a stack of two inertial sensor systems 10 and a cross-system control system 21, which is located as the top stack level in Fig. 2 is implemented. Together, the two inertial sensor systems 10 and the control system 21 form an inertial sensor system 20, which uses the two inertial sensor systems 10 as redundant sensor systems for fault detection and, if necessary, fault correction. It should be clear that more than two inertial sensor systems 10 can be used in the inertial sensor system 20 and that the placement of the cross-system control system 21 in the stack does not necessarily have to be at the top level (as in Fig. 2 (shown as an example) is limited.

[0038] The cross-system control system 21 comprises a power supply module 11 and a central system processor 12b, which is supplied with electrical energy via the power supply module 11. The central control processor 12b can be, for example, an ASIC, an FPGA, a PLD, a digital signal processor, a microcontroller, or another suitable portable computing device capable of processing incoming sensor signals and converting these signals into output values ​​that can be displayed at input / output interfaces 14b. The control processors 12a and 12b of the inertial sensor system 20 are connected via the system-connecting interfaces 13, through which the stacked inertial sensor systems 10, or rather their control processors 12a, can communicate with each other and with the central control processor 12b of the cross-system control system 21 and exchange data.The central control processor 12b may also have internal memory. It may also be possible that a separate (in . Fig. 2 (not explicitly shown) memory module on board 18b of the cross-system control system 21 is provided, which is coupled to the central control processor 12b, and from which the central control processor 12b can retrieve stored data for processing.

[0039] The central control processor 12b can receive the output data from the central control processors 12a of all inertial sensor systems 10 and combine them using appropriate plausibility checks. For example, the individual inertial sensor systems 10 can provide redundant data, which reduces the overall reliability of the inertial sensor system 20. It may also be possible to equip the individual inertial sensor systems 10 with inertial sensors with different measurement characteristics in order to minimize the probability of identical and therefore difficult-to-detect errors occurring. A "lock step" functionality can be implemented in the central control processor 12b, meaning that the central control processor 12b only outputs a complete sensor output value after the central control processors 12a of the individual inertial sensor systems 10 have output error-free, verified data for each measurement period.This enables dual modular redundancy, allowing the entire system 20 to exhibit higher fault tolerance. To implement a "lock step" functionality, the central control processor 12b can contain two processing cores that can execute the same instructions in parallel, and whose results are then mutually or cross-correlated checked.

[0040] The central control processor 12b can also be designed to initiate a built-in self-test routine (BIST) to verify measurement results from inertial sensors by means of redundantly available additional inertial sensors and, if necessary, to detect and correct any malfunction.

[0041] The modular design of the inertial sensor system 20 enables scalable configurations with multiple base modules from inertial sensor systems 10, which are interconnected via the system-wide control system 21. Depending on the number of base modules, various performance parameters are possible, and the configuration of the overall system can be flexibly adapted to the required safety standards. For example, the occurrence of Byzantine errors can be prevented with just three base modules: A single error on one of the modules, including the control system 21, can be reliably detected by majority vote. Accordingly, the probability of reliably detecting an error increases when more than three base modules are used.

[0042] Fig. Figure 4 shows a schematic block diagram of an inertial sensor system 20, which comprises two base modules 1 and 2 and a control system 21 coupled to both base modules. The power supply 11 of the base modules provides electrical energy to all inertial sensors 15a, 15b, 15c, 15d, 16a, 16b, 16c, and 16d, as well as to the central control processor 12a. The sensor readings from the individual inertial sensors are combined on the control system 21 via system interfaces 13 by means of the central control processor 12a of the individual base modules. The central control processor 12b of the control system 21 then combines and evaluates the incoming sensor readings according to predefined algorithms.

[0043] The in the Fig. 2 and Fig. The modular inertial sensor system 20 shown in section 4 can, for example, be used in a missile 40, which is exemplified in Fig. Figure 6 illustrates this. The inertial sensor system 20 can, for example, be part of an inertial navigation system (INS), which is designed, for instance, for use in a missile. Such a missile 40, which may have an inertial sensor system 20 (or inertial sensor system IMU), is shown as an example in Figure 6. Fig. Figure 6 illustrates this. The inertial sensor system 20 can be configured, in particular, as a strapdown inertial sensor system, i.e., an IMU in which the inertial sensors of the inertial sensor system 20 are rigidly connected to the outer frame of the missile 40. In particular, the accelerometers of such strapdown inertial sensor systems also exhibit a sensor measurement component attributable to the rotational motion of the missile 40, which can be compensated for using sensor signal values ​​from the gyroscopes to determine the translational acceleration values.

[0044] The functional process of an inertial sensor measurement using the inertial sensor system 20 is described in the Fig. 7 and Fig. 8 explained in more detail. It states Fig. 7 describes the process flow of a method for combining inertial sensor measurements in the control system 21 after all raw data from the sensors of the base modules have been acquired. The processing of the sensor data in the individual base modules, which is subject to the processing according to the method of Fig. Section 7, which precedes this, is discussed in greater detail in connection with the flowchart of the Fig. 8 explained.

[0045] Following an initialization phase (not explicitly shown) of all base modules and the system-wide control system, the sensor measurements within a base module 1 or 2 are preprocessed by the respective central control processor 12a in steps M(x) with x = 1, ..., n. In step M1, all raw sensor data are acquired, and in step M2, they are compensated with calibration data. In step M3, if no measurement range exceedances have already been detected, measurements from similar inertial sensors are compared and classified as plausible if a predefined error threshold for the deviation is not exceeded. In step M4, the plausible inertial sensor measurements from sensors of the same sensor type are optionally fused, and in step M5, the central control processor 12a outputs the base module data via the system interfaces 13.This is done separately for each of the basic modules 1 and 2.

[0046] In control system 21, after preprocessing M(x), all basic module outputs, including any error messages from the basic modules, are received and processed in step S1. This step S1 can be performed in lockstep mode to keep all basic modules synchronized. If a sufficient number of valid basic module outputs have been received, deviations between the individual basic module outputs can be detected in step S2. Parallel comparisons of sensor data belonging to the same inertial sensor type can be performed, so that if predefined maximum deviation tolerances are not exceeded, all basic module outputs can be fused in step S3, depending on the sensor type, and made available externally in step S4 as the overall output of the inertial sensor system 20. If necessary, corrective measures can be initiated before fusion if correctable errors occur.In the event of an uncorrectable error, a corresponding error message will be generated externally.

[0047] Fig. Figure 3 shows an inertial measurement unit (IMU) system for use in an inertial navigation system of a missile, such as the one in Fig. 5 depicted missiles 30. The inertial sensor system IMU is fundamentally similar to the inertial sensor system 10 of the Fig. The inertial sensor system 10 is constructed and essentially implements the functionality of forming virtual sensor group sets. For this purpose, the inertial sensor system 10 has at least two inertial sensor modules B1, B2, ..., Bm, each containing inertial sensors of the same type. For example, a first inertial sensor module B1 comprises at least two first inertial sensors C1, ..., Ci of a first inertial sensor type, and a second inertial sensor module B2 comprises at least two second inertial sensors D1, ..., Dj of a second inertial sensor type different from the first. For example, the first inertial sensors C1, ..., Ci can be accelerometers and the second inertial sensors D1, ..., Dj can be gyroscopes. An exemplary third inertial sensor module Bm can comprise at least two third inertial sensors E1, ..., Ek of a third inertial sensor type different from the first two.It should be clear that the variables i, j, k and m can take on different values ​​of two or more than two, and that the example shown is the . Fig. Figure 3, with three sensor modules and three sensors each, is merely an example. For the inertial sensor system (IMU), it is essential that one of the inertial sensor modules B1, B2, ..., Bm has more than one inertial sensor to allow a selection of inertial sensors depending on the operating conditions, at least in one of the inertial sensor modules B1, B2, ..., Bm. The other inertial sensor modules B1, B2, ..., Bm may, under certain circumstances, have only a single inertial sensor. For example, inertial sensor module B1 may have two inertial sensors C1, C2, from which a selection of the active inertial sensor can be made, while the other inertial sensor modules B2, ..., Bm may each have only a single inertial sensor D1 or E1. Under these circumstances, the multiplexer MUX2, ... can be used in the inertial sensor modules with only a single inertial sensor.MUXm can be dispensed with and the output of the single inertial sensor can be directly coupled to the control processor P.

[0048] While the inertial sensors within a given inertial sensor module all share the same sensor type, they can have different sensor designs and / or measurement characteristics. For example, primary sensors with high measurement accuracy but large size and / or expensive design can be used alongside secondary sensors with lower measurement accuracy but small size and / or cost-effective design. The rationale behind using multiple different sensors of the same type is that a missile's flight can typically be divided into different phases with significantly different requirements. For instance, a launch phase is characterized by high acceleration and low rotation rate.During such a phase, a high bandwidth is required for accelerometers, while the yaw rate sensors, although highly vibration-resistant and acceleration-independent, should have only minor distortion effects due to the short duration of the launch phase. In contrast, the free-flight phase following the launch phase is significantly longer, during which the accelerometers require a lower bandwidth but should exhibit very low distortion effects.

[0049] Depending on the flight phase, a central control processor P can select one or a subgroup of the inertial sensors available in each of the inertial sensor modules B1, ..., Bm as the active inertial sensors. This selection has the advantage that the inertial sensors with the most suitable performance parameters for the respective flight phase always provide the actual measured values, thus optimally minimizing the flight-phase-dependent dominant error terms. This type of adaptation can be applied to various performance parameters: bandwidth, vibration resistance, acceleration resistance, measurement accuracy, temperature dependence, drift value ("bias"), long-term stability value, or scale factor error value or scale factor nonlinearity. This results in "virtual" sensor systems for each flight phase, consisting of subgroups of selected active inertial sensors, which can be controlled with their own set of compensation parameters.

[0050] The central control processor P can determine the flight phase, for example, according to a predefined temporal sequence or use the measurement data from the inertial sensors themselves as a basis for decision-making. It may also be possible to specify the current flight phase to the central control processor P via an external source through an I / O data interface coupled to the control processor P. The inertial sensor system (IMU) can also include a power supply component (SV) that provides electrical power to the individual inertial sensor modules or their sensors. The power supply component (SV) can also provide electrical power to the central control processor P for its operation.

[0051] The inertial measurement unit (IMU) can have a memory (MEM) in which configuration data sets for error compensation of each of the first and second inertial sensors are stored. The control processor (P) can then select the currently active inertial sensors based on the configuration data sets stored in the memory (MEM) that describe this combination.

[0052] The selection of the currently active inertial sensors by the control processor P is carried out via multiplexers MUX1, MUX2, ..., MUXm, each of which is coupled to the outputs of the inertial sensors. The multiplexers MUX1, MUX2, ..., MUXm are controlled by the control processor P via selection signals Sel1, Sel2, ..., Selm to cause the respective multiplexers MUX1, MUX2, ..., MUXm to output the sensor readings O1, O2, ..., Om of the selected inertial sensors. The sensor readings O1, O2, ..., Om can then be passed to the control processor P for further processing.

[0053] A method for flight-phase-dependent inertial sensor measurement during the flight of a missile, which uses an inertial sensor system (IMU) of the Fig.The process, which can be carried out in step 3, first involves determining the current flight phase of the missile. Depending on the specific flight phase, a first active inertial sensor can be selected from at least two first inertial sensors C1, ..., Ci of a first inertial sensor type, depending on the specific flight phase; for example, an accelerometer. Likewise, a second active inertial sensor can optionally be selected from at least two second inertial sensors D1, ..., Dj of a second inertial sensor type different from the first; for example, a rate-of-turn sensor. The sensor readings of one active inertial sensor (or...)(Optionally, the two active inertial sensors) can then be processed in a control processor P for each flight phase, for example by the control processor P passing on the acquired acceleration sensor measurements and rotation rate sensor measurements to a flight control system to calculate the current flight attitude of the aircraft.

[0054] The error parameters of the two active inertial sensors can be compensated depending on configuration data sets describing this combination, which are stored in a memory MEM coupled to the control processor P.

[0055] During the missile's flight, the control processor P can monitor whether the missile is entering a new flight phase. If so, new first and second active inertial sensors can be selected depending on the flight phase the missile has currently entered. The control processor P then uses the sensor readings from the two new active inertial sensors for further processing.

[0056] In the preceding detailed description, various features have been summarized in one or more examples to improve the clarity of the presentation. However, it should be clear that the above description is merely illustrative and in no way limiting. It serves to cover all alternatives, modifications, and equivalents of the various features and embodiments. Many other examples will be immediately and directly clear to the person skilled in the art based on their technical knowledge, given the above description.

[0057] The exemplary embodiments were selected and described to best illustrate the principles underlying the invention and its practical applications. This enables those skilled in the art to optimally modify and utilize the invention and its various exemplary embodiments with regard to the intended purpose. In the claims and the description, the terms "including" and "comprising" are used as neutral language terms for the corresponding terms "comprehensive." Furthermore, the use of the terms "a," "a," and "an" is not intended to fundamentally exclude multiple features and components described in this way.

Claims

[1] Inertial sensor system (IMU) for use in an inertial navigation system of a missile (30), comprising: a first inertial sensor module (B1) comprising at least two first inertial sensors (C1, ..., Ci) of a first inertial sensor type and a first, multiplexer (MUX1) coupled with the outputs of the first inertial sensors (C1, ..., Ci); a second inertial sensor module (B2) comprising at least one second inertial sensor (D1, ..., Dj) of a second inertial sensor type different from the first inertial sensor type; and a control processor (P) which is coupled to the first multiplexer (MUX1) and which is designed to control the first multiplexer (MUX1) to output the sensor measurement values ​​of an inertial sensor selected by the control processor (P) from the first inertial sensors (C1, ..., Ci), wherein the control processor (P) is further designed to select an inertial sensor of the first inertial sensors (C1, ..., Ci) and an inertial sensor of the second inertial sensors (D1, ..., Dj) depending on an instantaneous flight phase of the missile (30). [2] Inertial sensor system (IMU) according to claim 1, further comprising: a memory (MEM) coupled to the control processor (P), in which configuration data sets for error compensation of each of the first and second inertial sensors are stored. [3] Inertial sensor system (IMU) according to claim 2, wherein the control processor (P) is further configured to select an inertial sensor of the first inertial sensors (C1, ..., Ci) and an inertial sensor of the second inertial sensors (D1, ..., Dj) depending on the configuration data sets stored in the memory (MEM). [4] Inertial sensor system (IMU) according to any one of claims 1 to 3, wherein the control processor (P) determines the instantaneous flight phase of the missile (30) on the basis of instantaneous sensor measurements of at least one inertial sensor of the first and second inertial sensors (C1, ..., Ci; D1, ..., Dj). [5] Inertial sensor system (IMU) according to any one of claims 1 to 4, wherein the first inertial sensor type comprises accelerometers and the second inertial sensor type comprises gyroscopes. [6] Inertial sensor system (IMU) according to any one of claims 1 to 5, wherein the second inertial sensor module (B2) comprises at least two second inertial sensors (D1, ..., Dj) and a second multiplexer (MUX2) coupled to the outputs of the second inertial sensors (D1, ..., Dj), and wherein the control processor (P) is further configured to control the second multiplexer (MUX2) to output the sensor measurement values ​​of an inertial sensor of the second inertial sensors (D1, ..., Dj) selected by the control processor (P). [7] Missile (30) with an inertial sensor system (IMU) according to any one of claims 1 to 6. [8] Method for flight-phase-dependent inertial sensor measurement during the flight of a missile, comprising the steps: Determining an instantaneous flight phase of the missile (30); selecting a first active inertial sensor from at least two first inertial sensors (C1, ..., Ci) of a first inertial sensor type depending on the determined flight phase; and Processing the sensor readings of the first active inertial sensor (C1, ..., Ci) in a control processor (P); Selecting a second active inertial sensor from at least two second inertial sensors (D1, ..., Dj) of a second inertial sensor type different from the first inertial sensor type, depending on the specific flight phase; and Processing the sensor readings of the second active inertial sensor (D1, ... Dj) in the control processor (P). [9] Method according to claim 8, wherein the first inertial sensor type comprises accelerometers and the second inertial sensor type comprises gyroscopes. [10] Method according to claim 9, wherein the processing of the sensor measurements comprises calculating the instantaneous flight attitude of the missile (30) based on acceleration sensor measurements and yaw rate sensor measurements. [11] Method according to any one of claims 8 to 10, further comprising the step: Compensating the error parameters of the active inertial sensors depending on configuration data sets describing this combination for each of the first and second inertial sensors (C1, ..., Ci; D1, ..., Dj) stored in a memory (MEM) coupled with the control processor (P). [12] Method according to any one of claims 8 to 11, further comprising the steps: Determine, by the control processor (P), whether the missile (30) enters the next flight phase; Selecting a new first active inertial sensor from at least two first inertial sensors (C1, ..., Ci) of a first inertial sensor type depending on the next flight phase; Selecting a new second active inertial sensor from at least two second inertial sensors (D1, ..., Dj) of a second inertial sensor type different from the first inertial sensor type, depending on the next flight phase; and Processing the sensor readings from the two new active inertial sensors in the control processor (P).

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