152 results about "Terminal guidance" patented technology

A semi-physical simulation system for developing a laser terminal guidance ammunition round comprises a rocket sled test platform, a target simulation device, an attitude simulation device and a simulation platform. The rocket sled test platform launches a rocket sled loaded with a target and measures speed and position information of the rocket sled in real time. The target simulation device utilizes laser to irradiate the target to enable a seeker of the laser terminal guidance ammunition round to be capable of receiving laser from the target to perform target tracking. The attitude simulation device comprises a three-axle rotary table capable of loading the seeker and is used for simulating attitude movement of a missile body. The simulation platform analyzes and processes the speed and position information of the rocket sled and a seeker generating signal according to a preset simulation mathematical model to obtain information related with attitude change of the missile body and relative movement of a missile and correspondingly controls the rotation of the three-axle rotary table. The rocket sled test platform is located outside the outdoor, and the rocket sled test platform and the simulation platform perform data transmission through a radio station and achieve time synchronization.

The invention discloses a several-for-one collaborative guidance method based on target probability distribution information, and belongs to the field of aircraft guidance. The problems that due to the fact that an existing guidance method is adopted in a near space hypersonic target, the target detection precision is low, the trajectory prediction accuracy is low, terminal guidance time is too short, so that the requirement for the overload response speed of an interceptor is too high are solved. The method is characterized by comprising the steps that relative motion measuring information between the interceptor and the target is acquired; Kalman filtering is used for processing the relative motion measuring information by considering the state value of the interceptor, the estimated value of the target state at the current moment is obtained, and a target terminal moment location probability density function is obtained by considering the kinetic equation of target motion; the integrals of the target terminal moment location probability density function in a plurality of interceptor reachable sets are constructed as a collaborative guidance cost function; several-for-one collaborative guidance optimizes problem solving is carried out. The method is applied to a plurality of aircrafts to intercept the hypersonic target in a collaborative mode, and has wide application prospects.

The invention discloses a multi-process constraint and multi-terminal constraint terminal guidance law based on a generalized target control miss distance concept. The multi-process constraint and multi-terminal constraint terminal guidance law comprises the following steps that 1, stored standard control and an aircraft state initial value are uploaded; 2, a ballistic integral is predicated; 3, process constraints are corrected; 4, a target control trajectory one-order variational model is solved; 5, replacement criteria control is solved. Compared with a traditional terminal guidance law, the guidance law can meet the traditional terminal constraint requirements for the miss distance, the falling angle, the falling point attack angle and the like, a strong-nonlinearity state variable V can be restrained to a needed value, and meanwhile the process constraints can be processed in a loop-locked mode.

The invention relates to a suboptimal guidance method with terminal angle constraint, and belongs to the field of guidance technologies. The suboptimal guidance method has six advantages that influence, on the guidance process, of aerodynamic characteristics (aerodynamic resistance and the gravity) of an air vehicle is considered, so that the method is closer to a practical situation; people only need to know initial conditions and tail end requirements of a flight and real-time state information of the air vehicle to achieve terminal guidance, and therefore the amount of needed information is small; an obtainable trajectory inclination angle terminal value is wide in range; an obtained controlled quantity is smooth in variation and is easy for an attitude control system to track; a weight matrix Q is the function based on a residual height y-to-go (or a residual firing range x-to-go), and the quantities are easy to measure in a project; not only can an object be hit with an expected terminal angle, but also a constant value speed object can be hit with an expected attitude.

The invention discloses a linear pseudo-spectrum GNEM guidance and control method. According to the linear pseudo-spectrum GNEM guidance and control method, an original nonlinear optimal control problem is converted into a group of problems for solving a system of linear algebraic equations based on the concept of GNEM through linearization and Gauss pseudo-spectrum dispersion by combining nonlinear approximation model predictive control, linear quadratic optimal control and a Gauss pseudo-spectrum method. The linear pseudo-spectrum GNEM guidance and control method has the advantages that the calculation efficiency of solving the optimal control problem is quite high, and high calculation precision can be obtained just through several nodes; in addition, a final solution can be presented through a smooth function relevant to control over disperse nodes, and the method is quite suitable for online calculation. It is indicated through a simulation result that by applying the method to tail-section attack guidance with terminal angle constraint, compared with an MPSP method, the linear pseudo-spectrum GNEM guidance and control method has higher calculation efficiency and high calculation precision and can be completely applicable to a guidance framework for terminal guidance, and a smaller required overload command is generated compared with self-adaptive tail-section proportion guidance.

The invention discloses a trajectory optimization method of a midcourse guidance with constraints. The trajectory optimization method is used for reducing a shift velocity of a terminal guidance in anintercepting missile. The method comprises the following steps: establishing a kinematic equation of a whole trajectory in a longitudinal plane of the intercepting missile including a trajectory constraint and a terminal constraint, and constructing an intercepting missile trajectory optimization model according to an optimal trajectory of midcourse guidance with constraints set by the method; and optimizing a solution of the trajectory by adopting an hp self-adaptive pseudo-spectral method. According to the method, the shift velocity of the terminal guidance in the intercepting missile can be reduced, a good flight regime of the terminal guidance can be provided and an intercepting precision is improved. And the method is a promising trajectory manner.

The invention discloses a finite time integral sliding mode terminal guidance law based on a super-twisting algorithm. The guidance law comprises the following steps: step 1, establishing a three-dimensional target-missile relative motion equation; step 2, designing a guidance law. A stability analysis is conducted on the guidance law. The invention provide the finite time integral sliding mode terminal guidance law based on the super-twisting algorithm by combining an integral sliding mode, a super-twisting algorithm and a finite time disturbance observer. According to the finite time integral sliding mode terminal guidance law based on the super-twisting algorithm, the finite time can be converged, the buffeting phenomenon can be eliminated, so that the maneuvering acceleration of a target can be better estimated, and the guidance accuracy is high; and the guidance law can be used for intercepting high-speed large-maneuvering targets such as unmanned aerial vehicles, ballistic missiles and the like, and has a wide application prospect.

The invention discloses a single-platform multi-sensor information integration processor and an experimental system. The single-platform multi-sensor information integration processor comprises a communication module, a control module, an information integration processing module and a power module. The experimental system comprises the single-platform multi-sensor information integration processor and a computer which are connected through a USB (universal serial bus) data line to interchange data. The experimental system is capable of simulating the terminal guidance process of multi-mode compound guidance ammunition, and comprises a small seeker loop and a large relative movement control loop for a guided missile and a target, wherein the small seeker loop comprises a digital seeker servo system of the computer and the multi-sensor information integration processor, and the large relative movement control loop comprises a missile-borne integrated control computer of the computer and the multi-sensor information integration processor. Information integration processing algorithm can be simulated in real time, and processing results can be displayed on line by the aid of the computer.

A terminal guidance and / or steering process for a projectile towards a target, process in which the orientation of a velocity vector {right arrow over (Vp)} is determined then a guidance law is applied and finally a steering algorithm enabling the projectile to be reoriented towards its target, process wherein the three components of the terrestrial magnetic field {right arrow over (H)} are measured in a projectile-linked reference marker (OxmYmZm) and these measurements are used in the guidance law and / or steering algorithm as a fixed reference marker enabling the orientation at least partially of the projectile-linked reference marker with respect to the terrestrial reference marker (GXfYfZf).

The invention discloses a reference sight angle signal-based design method of a multi-constraint terminal guidance law. The design method comprises the following steps: a guidance model of a guided missile and a static target in a horizontal plane is built, and a sight angle second-order dynamic equation is obtained; a reference sight angle signal is designed based on a quartic function of time, and unknown parameters in the reference sight angle signal are solved through boundary conditions, attack time and angle constraints; a guided missile advance angle kinetic equation is built to obtainthe reference sight angle signal expected to be tracked under a view field constraint of a guidance head; and the multi-constraint terminal guidance law is obtained based on an error model. The tracking reference sight angle signal-based method is adopted; no model linearization hypothesis exists; the residual flight time does not need to be estimated; and the method is suitable for a terminal guidance task with attack time, attack angle and guidance head view field constraints, and has relatively high applicability and relatively high reliability.

Provided is a terminal guidance method for maneuvering aircraft multi-terminal constraint backstepping sliding mode. The method comprises three main steps: step 1, analyzing terminal constraint of a terminal guidance problem and number; step 2, determining a sliding mode variable structure control order, designing a sliding variable quantity; step 3, and designing limited time backstepping, to obtain the guidance law of the whole system. The design basis of the method is the backstepping design of reaching in limited time, and the design of the terminal guidance law is a problem of reaching in a limited time. The backstepping design in variable structure control is a regression design method combining selection of a Lyapunov function and the design of a controller. Beginning from a minimum order differential equation of the system, the method designs in a stabilization manner layer by layer, and finally global stabilization is realized, so as to provide the control law of the whole system.

A multi-function radio frequency (MFRF) module integrates command guidance, active and semi-active terminal guidance (and possibly passive) and fuzing sensors for gun-launched munitions into a single assembly. The MFRF module can be incorporated into a variety of different gun-launched munitions to execute missions currently performed by guided missiles. The MFRF module is programmable during munition activation to select the guidance mode, active or semi-active, and a primary fuze mode, proximity or height of burst.

The invention discloses a front view navigation positioning method of an aircraft, comprising two stages of a remote navigation stage and a distance navigation stage; in the remote navigation stage, the position of the landmark centroid in a real-time figure is firstly identified, the estimated position of target centroid in the real-time figure is calculated by combining the position deviation between the target centroid and the landmark centroid in the geodetic coordinate system and current aircraft imaging attitude parameters furthermore, the relevant position difference value of the aircraft is calculated, and then a target search area is selected by taking the estimated position of the target centroid in the real-time figure; in the distance navigation stage, the target is directly positioned in the target search area, and then the relevant position difference value of the aircraft is calculated according to the direct positioning result. By means of landmark identification of the peripheral areas of the target, the target can be indirectly positioned for navigation, ensuring the robustness and adaptability of aircraft navigation positioning and overcoming the defect that the aircraft can be not directly positioned and navigated in the process of flight because the aircraft deviates from a planned flight path owing to errors.

The invention discloses a strapdown terminal guidance method for an unmanned aerial vehicle. The method comprises the following steps of firstly, collecting image information of a target area by a full-strapdown seeker of the unmanned aerial vehicle; secondly, splicing the image information, and carrying out suspicious target detection on the spliced image; thirdly, determining an actual positionof a suspicious target on the image, and adjusting an air route according to the actual position in order to make the unmanned aerial vehicle fly around the suspicious target; fourthly, confirming thetarget and tracking the target; and fifthly, after effective tracking is carried out on the target, enabling a ground station to issue an attack instruction, enabling a guiding system of the unmannedaerial vehicle to receive the instruction, judging whether the guiding system updates the guidance information, if the guidance information is updated, enabling the guiding system to apply terminal guidance control and judge whether a terminal guidance process is completed and whether a terminal guidance task is cancelled in real time, and if the guidance information is not updated, the terminalguidance process is completed or the terminal guidance task is canceled, returning the fourth step. According to the method, the cost is low, and a problem of navigation guidance control based on thefull-strapdown seeker of the micro-miniature unmanned aerial vehicle is solved.

The invention relates to a vision sensing-based bionic staged composite terminal guidance method, which belongs to the technical field of guidance. A staged tracking strategy of turning intercepting to tail chasing when a dragonfly tracks a fruit fly is simulated, a sliding mode guidance law of controlling a target parallax angle is used in an early stage of guidance to quickly adjust the flying direction of an aerocraft, and the aerocraft quickly enters a tail chasing state; in a terminal stage of guidance, a motion camouflage guidance law with a focus located at infinity is used to guide theaerocraft to track the target, and in a transition stage between the early stage and the terminal stage, a two-step slip intersection law is used. According to the terminal guidance method disclosedin the invention, when a micro aerocraft which depends on a vision sensor to detect a target intercepts a quick moving target, the required overload is smaller in comparison with the traditional guidance method, the target is closer to the visual field center of the vision detector, and the target can be ensured to be hard to escape from the visual field of the vision detector during the guidanceprocess.

The invention relates to a multi-constrained online Gauss pseudo spectrum type terminal guidance method. The multi-constrained online Gauss pseudo spectrum type terminal guidance method comprises the following steps that (1) terminal trajectory optimization is conducted on a nonlinear state equation according to the Gauss pseudo spectrum method; (2) in the same guidance cycle, an optimal control command is introduced into the actual flying process according to the clock maintaining mechanism; (3) in the same guidance cycle, online reconstruction of a guidance command is conducted according to the whole trajectory information of the prior cycle and the current state quantity; (4) in the next guidance cycle, the step (2) and the step (3) are executed repeatedly till an aircraft approaches a target; (5) when the aircraft approaches the target, the target is hit according to the proportional guidance method. According to the multi-constrained online Gauss pseudo spectrum type terminal guidance method, the terminal values of the high-nonlinearity state quantities such as the speed, the attack angle and the attack direction can be controlled, deviation caused by various kinds of interference can be well eliminated during flight, and the optimal control command can be reconstructed rapidly.

The invention aims at solving the technical problems of how to realize cooperative jamming of anti-ship missile terminal guidance radar by using chaff projectiles among platforms in a multiplatform formation when the multiplatform formation is threaten by an anti-ship missile, as well as uncertainty of passive interference modes adopted to avoid an interference vacuum phase under the condition that starting time of the terminal guidance radar cannot be determined. When a certain platform is tracked by the terminal guidance radar, an interference area in which other platforms can be cooperated with the tracked platform is divided according to distribution conditions of the other platforms, and eventually, a specific cooperative jamming platform and an interference mode needing to be adopted are determined according to the area in which anti-ship missile is located. The method disclosed by the invention is particularly suitable for a ship-borne platform to use.

The invention relates to the technical field of weapon systems and utilization engineering, and discloses a random quick smooth second-order sliding mode terminal guidance method. The method comprises the steps that firstly, a random uncertain guidance model is built; secondly, a random quick smooth second-order sliding mode control law is made; and thirdly, the random quick smooth second-order sliding mode control law is applied to the random uncertain guidance model to obtain a finite-time convergence random quick smooth second-order sliding mode guidance law. The method has the beneficial effects that the sliding mode variable convergence speed of the provided random quick smooth second-order sliding mode guidance law is obviously superior to that of a non-finite time guidance law; random maneuvering of a target is considered, and compared with a certain quick smooth second-order sliding mode method, the miss distance is smaller; the provided guidance law can restrain great buffeting of the control quantity of the guidance middle segment, and therefore the control quantity is easier to achieve, and the line-of-sight rate is more stable; the guidance law has the high anti-jamming capacity and capacity of dealing with random maneuvering of the target; and wide application prospects are achieved.

The invention discloses a terminal guidance forward-looking radar echo simulation method based on a missile motion parameter. The method comprises the steps: changing a sampling period of a missile motion parameter into a pulse repetition period through interpolation according to a missile-borne radar parameter; setting a ground scattering point within a ground area irradiated by missile-borne radar, and obtaining the slant-range history of the scattering point according to the position information of the scattering point and the missile motion parameter; obtaining phase information of a scattering point echo distance range time domain filter and phase information of the scattering point according to the slant-range history of the scattering point and the pulse width of the radar, and constructing a scattering point echo azimuth time domain filter according to the azimuth information of the scattering point and a radar antenna direction map; obtaining echo information of the scattering point according to the scattering coefficient of the scattering point, the phase information of the scattering point, the scattering point echo distance range time domain filter and the scattering point echo azimuth time domain filter, and obtaining echo information of a whole radar irradiation scene by means of a time domain coherent superposition method.

The invention discloses a parameter perturbation self-adaptive vertical attack guidance method for a guidance rocket projectile. By means of the method, the requirement of the terminal guidance segment of the rocket projectile with the vertical attack requirement for the overload capacity can be reduced easily. After the rocket projectile flies over the apex of a trajectory, the manners of numerical integration and the like are adopted, the position of the projection of the projectile body mass center of the projectile is calculated after the projectile flies under the standard full effectiveattack angle in a pressing manner to meet the vertical trajectory condition, the controlled miss distance is obtained through calculation, the estimated controlled miss distance is utilized in the first stage to achieve feedback control, the command of a steering engine is adjusted in real time, gliding range extending is achieved when the projectile is father away from a target, the projectile israpidly pressed to form a vertical trajectory when close to the target, and meanwhile flight errors generated by disturbance are further eliminated. After the vertical trajectory is formed by the projectile, the projectile is converted into terminal guidance, and the projectile is guided according to the proportion to attack the target.

The invention discloses a control system for a strapdown guidance unmanned aerial vehicle. The control system comprises an image seeker, an image processing module and a navigation guidance control module; the image seeker is strapped on the unmanned aerial vehicle and is used for capturing the image information of a target area; the image processing module sends the image information to a groundstation; the ground station selects an attack target from the image information, triggers a strap-down terminal guidance attack instruction, and transmits the position information of the selected target in the current frame image and the strap-down terminal guidance attack instruction to the image processing module; the image processing module obtains guidance information after receiving the position information and the instruction, and transmits the guidance information to the navigation guidance control module through a fast real-time subscription and distribution protocol; the navigation guidance control module receives the guidance information transmitted by the image processing module in real time and thereafter obtains the servo control quantity of the unmanned aerial vehicle throughan outer-ring flight path control loop and an inner-ring attitude control loop, so as to perform attitude control on the unmanned aerial vehicle.

The invention discloses a minimum overload terminal guiding method with terminal angular constraint. By using the method, the minimum theoretical value of a required overload under an ideal system can be obtained, and large terminal angle high-precision terminal guidance is obtained on the premise of being close to the minimum ideal required overload as far as possible. According to the content, the method comprises the steps that firstly, a circular-linear guidance law uCLGL(t) is calculated for the ideal system according to a target position and the terminal angular constraint, so that the lower limit of the required overload under the large terminal angle is determined, and the design is used for guiding a guided weapon to achieve the maneuvering capability required by the large terminal angle; secondly, according to a non-ideal system with lagged links of an automatic pilot, the tail of the circular-linear guidance law is switched into a polynomial guidance law with first-order derivative compensation, so that the combined guidance law is formed, the switching condition is given, the required overload behind the switching point is not larger than the required overload in front of the switching point, and meanwhile terminal position and angular errors can be effectively reduced.

The invention discloses a guidance rocket vertical attack guidance method based on falling point estimation. By using the method, a terminal guidance stage can be entered at a large falling angle, andthe requirement of the missile terminal guidance stage on overload capacity is lowered. In the early stage of the falling arc section, a trajectory is adjusted in a manner of firstly pulling up and then pressing down, so that a guidance rocket can enter the terminal guidance stage at a nearly vertical angle, the guidance control difficulty of the terminal guidance is reduced, vertical strike canbe effectively realized, and the attack efficiency of a warhead is improved. In addition, by introducing a cross point miss distance, the optimal control opportunity of the pull-up operation and the press-down operation in the third stage can be effectively obtained, and the mode is simple and reliable. Meanwhile, the ideal trajectory is tracked in the first stage, the trajectory is adjusted in the second stage according to the zero-effect prediction miss distance, the maneuverability of a rocket projectile on the whole trajectory is fully utilized, and the method has the advantages of being simple in algorithm, easy to realize, wide in attack range and the like.

The invention relates to a terminal guidance section simulation system, which comprises a guidance head and a movement simulation platform, wherein the guidance head adopts a simulation guidance head and consists of a holder and a camera; the movement simulation platform is of a net-type supporting frame which is provided with a guide rail installation beam and a transmission belt, wherein a precise guide rail is installed on the bottom part of the guide rail installation beam, a large installation platform is installed between two guide rail installation beams, a small installation platform is installed on the side surface of the guide rail installation beam, a roll-over drag chain is installed at the left side of a stepper motor, and the simulation guidance head is installed below the large installation platform. By adopting the simulation guidance head, the terminal guidance section simulation system has the advantages of long service life, high cost performance and simple structure; and the movement simulation platform adopts a steel tube net rack as a supporting frame, adopts the precise guide rail as the movement restriction, adopts the large-power stepper motor, the high-strength transmission belt and the roll-over drag chain as the towing, and adopts the computer to carry out the movement control, so the simulation system has the advantages that the simulated range of the earth surface for missile to pass is large, the platform motion is stable and is free from jumping, and the computer can conveniently perform the precise control.

The invention relates to the technical field of radar signal processing and discloses an angle-measuring and identifying method for target and bait interferences in a main lobe of monopulse radar. The method comprises the following steps: with an improved radar receiver, extracting an echo signal of an opposite angle difference channel and judging if bait interference exists; with single target angle measurement and single pulse angle measurement for target and bait, identifying true and false target angles; and controlling an angle tracking gate of the monopulse radar according to the angle measurement output of a true target. According to the angle-measuring and identifying method provided by the invention, the receiver is used for extracting a double-difference signal and processing the signal, so as to realize the angle position estimation for two undistinguishable targets, complete bait interference detection and true target identification, increase the interference detection and identification accuracy and calculation efficiency and enhance the instantaneity of interference detection. The angle-measuring and identifying method provided by the invention has important practical significance in increasing the angle distinguishing capability of monopulse radar and improving existing target indication radar and tracking guidance radar and terminal guidance radar.

The invention discloses a determining method for terminal guidance shift-exchange conditions of a reenter vehicle. The determining method includes (1), generating preset shift-exchange point information; (2), correcting the preset shift-exchange point information generated in step (1), and generating a preset shift-exchange point error ball; (3), designing guidance gain which is changeable with time; (4), determining a command smoothing time factor sequence sorted according to time; (5), determining shift-exchange logic, starting terminal guidance of a seeker when the shift-exchange logic is completely met; (6), in the process of midcourse guidance flight of the reenter vehicle, performing midcourse guidance by the aid of the guidance gain, which is changeable with the time, determined in step (3), judging whether the shift-exchange logic determined in step (5) is met or not in real time, starting the terminal guidance of the seeker when the shift-exchange logic is met, picking a factor from the command smoothing time factor sequence determined in step (4) according to order, performing smoothing on the midcourse guidance and the terminal guidance by the factor, performing guidance according to a command subjected to smoothing, and shifting to the simple terminal guidance of the seeker after preset time is reached.

A multi-function radio frequency (MFRF) module integrates command guidance, active and semi-active terminal guidance (and possibly passive) and fuzing sensors for gun-launched munitions into a single assembly. The MFRF module can be incorporated into a variety of different gun-launched munitions to execute missions currently performed by guided missiles. The MFRF module is programmable during munition activation to select the guidance mode, active or semi-active, and a primary fuze mode, proximity or height of burst.

The invention provides an analysis method of handover conditions of a hypersonic-velocity target interception missile based on interception geometry and belongs to the field of aircraft manufacturing technology. The method is technically characterized in that by considering all possible velocity direction of a target and according to the relative motion relation between the interception missile and the target, positions of all possible impact points are obtained through derivation, and the interception geometry is built; according to a research result of the interception geometry and the positional relation among the interception missile at the midcourse and terminal guidance handover moment, the target and the interception geometry, interception conditions required to be met for intercepting the hypersonic-velocity target are given; on the basis that the interception conditions are met, a calculation method of the positional condition and the solving process of the angle condition of the interception missile at the midcourse and terminal guidance handover moment are given, and the influence of maneuvering of the target on solving the angle condition of the interception missile is analyzed. Through the method, the problem that the hypersonic-velocity target cannot be intercepted through an existing interception method is solved.

The invention relates to an analytical guidance method for a standard control miss distance of a terminal guidance section. The analytical guidance method comprises the following steps: firstly, modeling the terminal guidance section; secondly, dividing a standard control trajectory at the terminal guidance section into a zero-order standard control trajectory and a first-order standard control trajectory according to a canonical perturbation theory; thirdly, performing analytic solution on a zero order item of the standard control trajectory at the terminal guidance section; fourthly, performing analytic solution on a first order item of the standard control trajectory at the terminal guidance section; and fifthly, acquiring a complete analytical solution of the standard control trajectory, and calculating a guidance command. The analytical guidance method disclosed by the invention has the advantages that firstly, an idea of taking a zero attack angle as standard control over the terminal guidance section of a homing missile is put forward, and a motion equation of the standard control trajectory is simplified; in addition, an attack angle required by guided correction is enabled to be located near zero, and speed loss brought by the guided correction is reduced; secondly, analytic solution is performed on a longitudinal trajectory at the terminal guidance section by adopting a canonical perturbation method, and high-precision analytical solutions of firing range, height, speed and trajectory inclination angle are obtained; and thirdly, a guidance method based on a generalized standard control miss distance is provided, and normal acceleration of the missile is greatly reduced; and meanwhile, the terminal attack angle is enabled to be converged to be zero.

The present invention extends to methods, systems, and computer program products for formulating lane level routing plans. In general, aspects of the invention are used in motorized vehicles to guide a driver to a terminal vehicle configuration according to a lane level routing plan that balances travel time with routing plan robustness. A lane level routing plan can be based on terminal guidance conditions (e.g., exiting a highway in the correct off ramp lane), statistical patterns of lanes themselves, current vehicle state, and state of the local environment near the vehicle. Lane level routing plans can be communicated to the driver with audio, visual, and / or haptic cues. Lane level routing plans can be revised online and in (essentially) real-time in response to changing conditions in the local environment (e.g., a trailing vehicle in a neighboring lane has decided to increase speed).