Abstract: Technologies for guidance of a spin-stabilized orbital rocket are described herein. The spin-stabilized rocket includes a guidance controller. The guidance controller computes parameters of a burn of a second-stage engine of the rocket to reach a desired nominal orbit subsequent to burnout of the first stage of the rocket. The guidance controller computes the burn parameters of the second-stage engine based upon one or more desired orbit parameters and a current position and velocity of the rocket. The computation of the burn parameters is based upon a simulated point-mass model of the motion of the rocket. The guidance controller then controls the rocket to initiate a second-stage burn having the computed burn parameters.

Abstract: An apparatus comprising an interface, a memory and a processor. The interface may be configured to receive (i) sensor data samples during operation of a vehicle and (ii) data from a telemetry system. The memory may be configured to store the sensor data samples over a number of points in time. The processor may be configured to (i) analyze the sensor data samples stored in the memory to determine (a) operational parameters of the vehicle and (b) information associated with a second vehicle and (ii) adjust the operational parameters in response to the information associated with the second vehicle. The data from the telemetry system and the information associated with the second vehicle may be used to determine an amount of space between the vehicle and the second vehicle. The operational parameters may be adjusted to keep the amount of space within a pre-determined range.

Abstract: A device, system, and method for shaping the trajectory of a projectile employing a Gravity bias, Gbias. The system includes a seeker, a guidance filter, a pitch rate filter, an actuator, pitch/yaw/roll coupled aerodynamics, and lateral rate sensors. It receives roll orientation input to a guidance and control autopilot; it applies Additional Gbias to that produced by the null rate command to the lateral control loops of the guidance and control autopilot device. The lateral rate command is equal to the desired Additional Gbias divided by an estimate of the projectile velocity. The Additional Gbias is translated to a rate command and incorporated into guidance loop commands to boost an Inherent Gbias to shape the trajectory of the projectile to the target.

Abstract: Embodiments include active protection systems and methods for an aerial platform. An onboard system includes radar modules, detects aerial vehicles within a threat range of the aerial platform, and determines if any of the aerial vehicles are an aerial threat. The onboard system also determines an intercept vector to the aerial threat, communicates the intercept vector to an eject vehicle, and causes the eject vehicle to be ejected from the aerial platform to intercept the aerial threat. The eject vehicle includes alignment thrusters to rotate a longitudinal axis of the eject vehicle to substantially align with the intercept vector, a rocket motor to accelerate the eject vehicle along an intercept vector, divert thrusters to divert the eject vehicle in a direction substantially perpendicular to the intercept vector, and attitude control thrusters to make adjustments to the attitude of the eject vehicle.

Abstract: A method for guiding a ballistic missile to a target has a first mode guidance process that drives a magnitude of a velocity-to-go (Vgo) vector toward zero. On regular intervals, a proportional velocity deficit value is calculated as equal to a time constant (Tau) multiplied by a specific force magnitude (sf). When the magnitude of the Vgo vector has been driven to less than or equal to the proportional velocity deficit value by way of the first mode guidance process, a second mode guidance process is initiated. The second mode guidance process constrains the magnitude of the Vgo vector to be equal to the proportional velocity deficit value throughout the remaining portion of powered flight.

Abstract: A rotating electrical machine for a motor vehicle comprising a rotor (2) supplied with an excitation current (ie), a stator (3) comprising a polyphase winding and coupled to the rotor (2), a thermal protection module (4) suitable for evaluating at least one temperature (T1, T2, T3) in the machine and for comparing the temperature with an associated thermal protection threshold (Th), a control device (5) supplying the excitation current (ie) depending on an operation mode command (RQ) and the temperature comparison, so as to operate the rotating electrical machine (1) according to a mode of operation chosen from a nominal generator mode in which the machine is configured to deliver a first maximum power or at least one extended generator mode in which the machine is configured to deliver a second maximum power greater than said first maximum power.

Abstract: The invention relates to a remote-controlled system comprising: —at least one ground interface (3), from which an operator can control a remote-controlled vehicle; —at least one mission unit (7, 8) in said vehicle; and —a data link between said interface (3) and said mission unit (7, 8). Said system is characterized in that it comprises, on the ground and in the vehicle, security monitoring systems (6, 10) suitable for approving and/or authenticating critical data and/or commands exchanged between the ground and the vehicle and also suitable for verifying the integrity of said data. It is thus possible to use, on the ground as on board the vehicle, interfaces and units with a low level of criticality at the same time as interfaces and units with the highest level of criticality.

Abstract: A method and apparatus for generating navigation solutions. A global positioning system based navigation solution unit; an inertial navigation solution unit; a correction unit, a limiter, and an adding unit. The global positioning system based navigation solution unit is capable of generating a first navigation solution. The inertial navigation solution unit is capable of generating a second navigation solution. The correction unit is capable of generating a raw correction. The limiter is capable of selectively modifying the raw correction to fall within a selected range of corrections to form a correction. The adding unit is capable of adding the correction to the second navigation solution to form a navigation solution.

Abstract: A new method for determining an attitude of a sensor with respect to a moving body is proposed. A movement vector is measured by a sensor mounted in the moving body. In addition, the attitude of the sensor with respect to the moving body is determined using the movement vector measured by the sensor when the moving body starts to move.

Abstract: A 3-D environment mapping systems and methods of dynamically mapping a 3-D environment includes identifying at least five point correspondences associated with image frames of a scene recorded by a camera from a plurality of viewpoints. A down direction is determined from IMU motion data. The down direction is an absolute global downward direction. Components of the scene that are vertical are identified using the down direction, an iteration of 3-D position of each of the five point correspondences, and a current iteration of refined relative locations and poses of the camera. A 2-D map is generated of the scene by mapping the identified components of the scene in an eagle view of the scene.

Abstract: A rocket vehicle includes a controller that integrates operation of a variable-vector main thruster and attitude control thrusters. When the main thruster is firing and roll is commanded, the controller can provide roll moment by firing only a single attitude control thruster, while changing the thrust vector of the main thruster to offset any pitch/yaw moments induced by the firing of the single attitude control thruster. The single attitude control thruster may be a thruster on the leeward side of the rocket vehicle. Since there is a lower wall pressure on the leeward side of the rocket vehicle, the thruster efficiency is improved by accomplishing roll by use of a single thruster (which may be one of a pair of thrusters used to achieve roll in one direction). A significant reduction in fuel use may be accomplished.

Abstract: A method for visualizing a position of a vehicle configured for navigating a pipe within the pipe on a display device wherein the visualization on the display device is at least partially performed based on a virtual representation and the pipe is graphically represented at least in sections and preferably in a form of lines defining its enveloping surfaces on the display device, wherein the vehicle and preferably its movement is also represented on the display device thus within the pipe.

Abstract: Inertial measurement unit apparatus for use with guidance systems are disclosed herein. An example guidance system includes an inertial measurement unit removably coupled in a cavity of a guidance wafer via an access port of the guidance wafer defining a port axis that is non-parallel relative to a longitudinal axis of the guidance wafer.

Abstract: The present invention relates to a method for measuring a time delay between two output signals measured by two adjacent optical sensors, the time delay resulting from a movement of an object moving in a given direction. The method comprises the steps: carrying out a spatial filtering of the signals Ph1 and Ph2 delivered by a first and a second optical sensor; calculating a first order derivative and carrying out a temporal low-pass filtering of the signals; calculating the second order derivative of the signals; measuring the delay between the signals; and is characterized in that the step of measuring the delay is a feedback loop based on an estimated delay between the temporally filtered signals. Device and set of devices for measuring the angular velocity of a luminance transition zone and steering aid system for fixation and tracking a target comprising at least one such luminance transition zone.

Abstract: A method and system for piloting a craft with a rear propulsion unit are disclosed. The method can include a servo loop where the attitude (?M) of the craft (1) is measured in the vicinity of the rear end (1R) of the craft, then the orientation (?) of the propulsion means (2), which can be oriented relative to the rear end (1R), is adjusted as a function of the attitude measurement (?M) in such a way that the craft (1) is stabilized on its flight path.

Abstract: Embodiments of a flight-control system and methods of reducing the probability of a roll-control reversal in a canard-controlled flight vehicle are generally described herein. In some embodiments, the flight-control system may monitor angular velocities of the flight vehicle to detect the onset of instability and adaptively control an acceleration limit of the flight vehicle based on the detected instability to reduce the probability of a roll-control reversal. The onset of instability may be detected by persistently high angular velocities. The acceleration limit may be further adaptively controlled based on an approach of a vehicle trim limit.

Abstract: The invention relates to a flying device (1), in particular an air missile, which is of the inertial navigation type, which is rotatable and includes a pitch rate gyroscope (2), wherein said device also comprises automatic control means (8) for rotating same about the longitudinal axis (3) thereof, regularly alternating the direction of rotation, so as to cancel the effect of the scale factor of the pitch rate gyroscope (2), in addition to correcting ordinary drifts.

Abstract: A scalable, inert munition data recorder cartridge. This inert cartridge allows users to record data from within a weapon system feed chute, breach, and extractor port. This cartridge is adaptable to any weapon system, packaging unit, environmental chamber, or other ammunition holding and storage device. Although the cartridge is designed to interface with a weapon and ammunition packaging based on its shape, it still functions as designed regardless of its location and application. This would allow the cartridge to be placed anywhere an actual live cartridge of ammunition could be placed.

Abstract: The invention relates to a guidance system comprising estimation means able to estimate, in the course of flight, the attitude and the aerodynamic speed of a projectile, as well as the variations in the speed of the wind, on the basis of guidance orders formulated by guidance means of the guidance system, of a reference trajectory and of measurements obtained by measurement means of the system, using a model of the dynamic behavior of the projectile and a model of the dynamics of the wind.

Abstract: A SAR image recorded by a reconnaissance system is transferred as a reference edge image together with the data of the trajectory as a reference. The signal of the infrared seeker head of the missile is converted into a virtual SAR edge image and compared to the SAR reference image to calculate the precise position of the missile.

Abstract: An apparatus and method improves the fault tolerance of a rocket or missile guidance system which includes a resonant sensor. When improper initialization is detected, the resonant sensor is reinitialized, repeatedly if necessary, until normal operation is achieved. Improper initialization is detected by comparing data from the guidance system with pre-specified physical limits to roll, pitch, yaw, and/or other features of the flight scenario. Embodiments can also detect a fault condition due to an error signal from a “Built-in-Test” (BIT) module. The initialization sequence initiated by the invention can be identical to the power-on sequence, or it can be a separate, reinitiating sequence. Subsequent resets are initiated as needed, for example until the burn of the rocket fuel and the associated vibrations have ceased and the resonant sensor has been successfully initialized.

Abstract: A projectile's payload is oriented (independently or by orientation of the projectile itself) toward a target just prior to firing (e.g., detonation of the payload), e.g., for munitions providing an increased kill and casualty area and a fire “in defilade” (left, right, backwards or at any angle) capability.

Abstract: Systems and methods of controlling solid propellant gas pressure and vehicle thrust are provided. Propellant gas pressure and a vehicle inertial characteristic are sensed. Propellant gas pressure commands and vehicle thrust commands are generated. A propellant gas pressure error is determined based on the propellant gas pressure commands and the sensed propellant gas pressure, and vehicle thrust error is determined based on the vehicle thrust commands and the sensed vehicle inertial characteristic. Reaction control valves are moved between closed and full-open positions based on the determined propellant gas pressure error and on the determined vehicle thrust error. The system and method allow the reaction control valves to operate at variable frequencies or at fixed frequencies. The system and method also allows propellant pressure to be commanded to follow a predetermined pressure profile or commanded to vary “on-the-fly.

Abstract: A projectile, such as a missile, rolls during at least a portion of its flight, while retaining its roll reference to enable navigation during the rolling period of flight. The roll reference may be retained by using a sensor, such as magnetometer, to periodically check and correct the roll reference. Alternatively or in addition the missile may alternate roll directions, for example varying roll rate in a substantially sinusoidal function. By rolling the missile inaccuracies in an inertial measurement unit (IMU) of the missile may be ameliorated by being to a large extent canceled out by the changes in orientation of the missile as the missile rolls. This enables use of IMUs with lower accuracy than would otherwise be required to obtain accurate flight. Thus accurate flight may be accomplished with less costly IMUs, without sacrificing the ability to navigate.

Abstract: Higher Order Sliding Mode (HOSM) control techniques are applied to the Guidance Control (G&C) of interceptor missile in which velocity may be steered by combination of main thrust, aerodynamic lift and lateral on-off divert thrusters, and attitude may be steered by continuous or on-off actuators. Methods include the pointing of the seeker, its associated estimation processes, a guidance law that uses concurrent divert mechanisms, and an attitude autopilot. The insensitivity of the controller to matched disturbances allows the concurrent usage of the divert mechanisms without adverse effect on the accuracy. The controller also allows the de-coupling of the control of roll, pitch and yaw channels, and usage quaternions to represent body attitude and it provides control perfect robustness. While it conceivable to design separately the components of the G&C method, it is widely accepted that designing them in an integrated fashion usually produces a better result.

Abstract: A projectile's payload is oriented (independently or by orientation of the projectile itself) toward a target just prior to firing (e.g., detonation of the payload), e.g., for munitions providing an increased kill and casualty area and a fire “in defilade” (left, right, backwards or at any angle) capability.

Abstract: A missile has a pair of systems to provide acceleration information during flight. The primary system is a microelectromechanical systems (MEMS) inertial measurement unit (IMU) that provides accurate rate sensor output, such as providing pitch and yaw rates, at low cost, over a wide range of conditions. However MEMS IMUs are susceptible to temporary incorrect responses when subjected to shocks, such as acoustic-range shocks, for instance in the range of 10-20 kHz. The missile includes a secondary system to temporarily provide acceleration data during the periods following shocks, when the MEMS IMU does not provide valid (reliable or usable) rate sensor output, for use in estimating pseudo pitch and yaw rates. The secondary system may be an accelerometer that does not provide navigation-quality acceleration data, but does provide a sufficiently accurate response in order to maintain stable flight during the post-shock period.

Abstract: A system for estimating at least one of position, attitude, and heading of a vehicle is disclosed. The system includes at least three gyroscopes configured to output a signal indicative of inertial angular rates around three mutually orthogonal axes of the vehicle and at least three accelerometers configured to output a signal indicative of accelerations along three mutually orthogonal axes of the vehicle. The system further includes a triaxial magnetometer configured to output a signal indicative of a projection of ambient magnetic field on three mutually orthogonal axes of the vehicle. The system also includes a sensor configured to output a signal indicative of vehicle altitude and a differential pressure sensor configured to output a signal indicative of airspeed of the vehicle. The system further includes a device configured to receive the signals and estimate at least of one of position, attitude, and heading of the vehicle.

Abstract: An exemplary navigation system uses a master navigation component at a first location in a vehicle and a slave navigation component at a second location that is a variable displacement to the first location due to physical deformation of the vehicle. Static and dynamic location components provide static and dynamic information of the displacement between the first and second locations. An error estimator estimates errors in the navigational measurement data generated by the slave navigation component based on the navigational measurement data generated by the master navigation component and the displacement information provided by the static and dynamic location components. The master navigation component corrects the navigation measurement data of the slave navigation component based on the determined error and translates the corrected navigation measurement data of the slave navigation component into navigational measurement data in its coordinate system.

Abstract: A guided projectile may include a projectile body. An inertial measurement unit may be disposed within the projectile body, the inertial measurement unit including sensors to measure motion parameters relative to first, second, and third mutually orthogonal axes. Each of the first, second, and third mutually orthogonal axes may form an oblique angle with a longitudinal axis of the projectile body. A controller may be configured to control a trajectory of the guided projectile in response, at least in part, to measurement data received from the inertial measurement unit.

Abstract: The method for generating an integrated guidance law for aerodynamic missiles uses a strength Pareto evolutionary algorithm (SPEA)-based approach for generating an integrated fuzzy guidance law, which includes three separate fuzzy controllers. Each of these fuzzy controllers is activated in a unique region of missile interception. The distribution of membership functions and the associated rules are obtained by solving a nonlinear constrained multi-objective optimization problem in which final time, energy consumption, and miss distance are treated as competing objectives. A Tabu search is utilized to build a library of initial feasible solutions for the multi-objective optimization algorithm. Additionally, a hierarchical clustering technique is utilized to provide the decision maker with a representative and manageable Pareto-optimal set without destroying the characteristics of the trade-off front. A fuzzy-based system is employed to extract the best compromise solution over the trade-off curve.

Abstract: The invention concerns a method for improving determination of inertial navigation parameters (1) of a carrier (1) moving along modelizable kinematic movement components, the method including the following steps: (a) selecting, taking into account a kinematic model (13) of the carrier (1), at least one movement component whereby integration (11) of the inertial measurements (20) is assumed to give a predetermined value, (b) integrating (11) the inertial measurements (20) in accordance with the selected component(s), (c) determining (14) based on the component(s) a variation between the integration (11) obtained at step (b) and the predetermined value of step (a), (d) estimating on the basis of the variation(s) thus obtained a global inertial error resulting from internal errors associated with said modelizable components, and values to be updated of variable parameters of the kinematic model (13), (e) correcting said inertial navigation based on the thus determined global inertial error.

Abstract: The effects of IMU gyro and accelerometer bias errors are significantly reduced in accordance with the present teachings by a system or method for commanding an IMU or vehicle through a series of preprogrammed maneuvers. The maneuvers can be designed to minimize the effects of other gyro errors including scale factor errors, nonlinearities, cross coupling/misalignment, and scale factor asymmetries. A sample maneuver is provided which demonstrates performance based on a sequence of roll and yaw maneuvers resulting in zero build up of error at the end of a maneuver cycle period as a result of these errors. Modification of the system involves the addition of control logic to determine the maneuver period, maneuver rate, and vehicle orientation. No additional hardware beyond possible fuel required to perform the maneuver is required.

Abstract: A missile has a pair of systems to provide acceleration information during flight. The primary system is a microelectromechanical systems (MEMS) inertial measurement unit (IMU) that provides accurate rate sensor output, such as providing pitch and yaw rates, at low cost, over a wide range of conditions. However MEMS IMUs are susceptible to temporary incorrect responses when subjected to shocks, such as acoustic-range shocks, for instance in the range of 10-20 kHz. The missile includes a secondary system to temporarily provide acceleration data during the periods following shocks, when the MEMS IMU does not provide valid (reliable or usable) rate sensor output, for use in estimating pseudo pitch and yaw rates. The secondary system may be an accelerometer that does not provide navigation-quality acceleration data, but does provide a sufficiently accurate response in order to maintain stable flight during the post-shock period.

Abstract: A guided projectile may include a projectile body, an inertial measurement unit disposed within the projectile body, one or more control surfaces extendible from the projectile body, and a controller which controls the one or more control surfaces in response, at least in part, to measurement data received from the inertial measurement unit.

Abstract: A guided projectile may include a projectile body. An inertial measurement unit may be disposed within the projectile body, the inertial measurement unit including sensors to measure motion parameters relative to first, second, and third mutually orthogonal axes. Each of the first, second, and third mutually orthogonal axes may form an oblique angle with a longitudinal axis of the projectile body. A controller may be configured to control a trajectory of the guided projectile in response, at least in part, to measurement data received from the inertial measurement unit.

Abstract: Flight management systems can behave erratically when the distance measurements on which they are based are subject to value jumps because they liken these value jumps to movements of the aircraft performed at speeds exceeding the performance levels of the aircraft for which they were designed. To avoid this, the proposed flight management system uses a filter to spread the distance value jumps in time, over periods of the order of those needed for the aircraft to come through the distance differences that they represent. This filter replaces a value jump with a ramp making up the difference and corresponding to a movement that remains within the performance scope of the aircraft.

Abstract: One example embodiment relates to a method of navigating an object. The method includes detecting when the object accelerates through the speed of sound and maneuvering the object based on when the object accelerates through the speed of sound. Another example embodiment relates to a system for navigating an object. The system includes a detector within the object. The detector determines when the object accelerates through mach one. The system further includes a guidance system within the object. The guidance system adjusts the flight of the object based on data received from the detector.

Abstract: A system and method of detecting, processing, and selectively responding to radio frequency transmissions detected by at least one projectile deployed above a geographic area.

Abstract: A Reconfigurable Nose Control System (RNCS) is designed to adjust the flight path of spin-stabilized artillery projectiles. The RNCS uses the surface of a projectile nose cone as a trim tab. The nose cone may be despun by the action of aerodynamic surfaces, to zero spin relative to earth fixed coordinates using local air flow, and deflected by a simple rotary motion of a Divert Motor about the longitudinal axis of the projectile. A forward section of the nose cone having an ogive is mounted at an angle to the longitudinal axis of the projectile, forming an axial offset of an axis of the forward section with respect to the longitudinal axis of the projectile. Another section of the nose cone includes another motor, the Roll Generator Motor, that is rotationally decoupled from the forward section and rotates the deflected forward section so that its axis may be pointed in any direction within its range of motion.

Abstract: A position sensing assembly includes a bearing element and a helically shaped rotational member used to drive a portion of a sensor assembly, such as a Digital Rotary Magnetic Encoder. Interaction between the bearing element and a helically shaped rotational member minimizes the presence of backlash in the position sensing assembly. Accordingly, as an actuator assembly drives both a control element, such as a flight control surface, and the position sensing assembly, the sensor assembly generates an output signal that accurately reflects the position of the control element.

Abstract: A thermally controlled gas bearing supported inertial measurement unit (IMU) system is provided. The system comprises a sensor assembly enclosing one or more sensors and a plurality of heating elements, wherein each of the plurality of heating elements is proximal to the sensor assembly. The system also comprises a plurality of temperature sensors configured to determine a temperature of a region of the sensor assembly and a control unit configured to adjust a temperature of at least one of the plurality of heating elements based on feedback from the at least one temperature sensor.

Abstract: A method of controlling flight of a missile includes using gyroscopes, such as pitch rate gyroscopes, to sense when a factor based on the angular rate of change of the missile exceeds a threshold value. One the threshold value is exceeded, a decision may be made to use one or more compensation thrusters to reduce the angular rate of change. The use of the compensation thrusters may correct residual angular velocities from a pitch over maneuver used to put the missile on an intended course. In addition, the compensation thrusters may be used to compensate for errors in missile heading induced after the pitch over maneuver, such as induced by misalignment of thrust provided by a main rocket motor of the missile. Multiple compensation thrusters may be used to compensate for angular changes in the pitch and yaw directions.

Abstract: Rolling airframe projectile guidance and stability systems are disclosed. Flight control surfaces, such as canards and/or tail fins are attached to a projectile airframe that is designed to roll during flight. Stepper motors are attached to the flight control surfaces and move the flight control surfaces in discrete increments. A control system generates signals that control the flight control surfaces. The control system may include a neural network that is trained to generate control signals in response to received inputs.

Abstract: A projectile, such as a missile, rolls during at least a portion of its flight, while retaining its roll reference to enable navigation during the rolling period of flight. The roll reference may be retained by using a sensor, such as magnetometer, to periodically check and correct the roll reference. Alternatively or in addition the missile may alternate roll directions, for example varying roll rate in a substantially sinusoidal function. By rolling the missile inaccuracies in an inertial measurement unit (IMU) of the missile may be ameliorated by being to a large extent canceled out by the changes in orientation of the missile as the missile rolls. This enables use of IMUs with lower accuracy than would otherwise be required to obtain accurate flight. Thus accurate flight may be accomplished with less costly IMUs, without sacrificing the ability to navigate.

Abstract: An inertial navigation system is provided. The system includes a sensor block, an outer shell that substantially surrounds the sensor block and a plurality of gas pads connected to the outer shell that float the sensor block in gas creating a near frictionless environment to allow the sensor block to rotate in all directions. Each of the plurality of gas pads is adapted to receive pressurized gas. The outer shell and the sensor block are separated by a gap created by the pressurized gas at each pad.

Abstract: A missile guidance system and method for guiding a missile, mainly horizontally flying, to pass a target at a desired passage height.

Abstract: A guidance method for a powered ballistic missile involves using an onboard computer to numerically simulate the flight path of the missile in real time, using a model with at least 3 degrees of freedom. The results of this simulation are used to update in real time an aim point and/or a predicted intercept point. An iterative process may be used in adjusting the aim point and/or the predicted intercept point. The process may be carried out until a specified number of steps have been completed, and/or until a specified heading error threshold of the aim point and a specified time of flight threshold have been achieved. The use of real time updating of an aim point of the missile advantageously takes into account variations in missile velocity and position due to individual variations in the rocket motor of the missile.

Abstract: A method and system for degimbalization of sensor outputs is provided. Data output from an embedded GPS inertial navigation system (EGI), which is mounted within a gimbaled system on a vehicle, is processed to provide degimballed navigation data. Generally, motion of the EGI is due to the vehicle motion and the gimbal motion. To provide navigation information relating solely to the vehicle, effects of the gimbal motion within the EGI outputs can be removed.

Abstract: A method for producing inertial measurement data is provided. The method comprises receiving raw inertial measurement data from one or more inertial sensors; receiving raw position data based on signals from a global navigation satellite system; processing the raw inertial measurement data and the raw position data with a filter to generate state variable estimates; and calculating enhanced inertial measurement data based on the raw inertial measurement data and the state variable estimates from the filter.