Patent Abstract:
A navigation system with resume logic and mode logic provides as an output an accurate navigation solution using multiple RF sensors. The resume logic determines which sensors are currently providing good data to the navigation system. The mode logic selects an operating mode of the navigation system and selects which data to use for calculating corrections to the navigation solution. The mode logic makes the selections based on the results of the resume logic. The resume logic continues to test data from the sensors. If a sensor that has previously provided erroneous data starts providing good data, the mode logic will automatically select that data for use in calculating the corrections to the navigation solution. The tracking of RF transmitters by the multiple RF sensors is controlled using a plurality of available inertial and non-inertial sensors.

Full Description:
FIELD  
       [0001]     The present invention relates generally to avionic systems, and more particularly, using multiple aiding sensors in a deeply integrated navigation system. The present invention, in addition to being applied to airborne applications, can be applied to land and underwater applications.  
       BACKGROUND  
       [0002]     A pilot receives information from many sources during take-off, flight, and landing of an aircraft. The aircraft includes avionic systems designed to collect data, perform calculations on the data, and present the data to the pilot. For example, the aircraft may include an inertial navigation system (INS), an Attitude Heading Reference System (AHRS), an air data computer, a roll-pitch-yaw computer, a mission computer, various displays, and other avionic systems. Some avionic systems may include one or more sensors that collect data, such as attitude, heading, altitude, and air speed. The same or other avionic systems may process this data. Avionic displays may present the data to the pilot in a usable format.  
         [0003]     If one of the sensors becomes inoperable, the pilot may have to rely on the information he can obtain from other sensors to continue the flight and land the aircraft safely. For example, the INS and the AHRS systems may provide similar information to the pilot. Both systems may provide attitude and heading information to the pilot. If there was a problem with the INS, the pilot may obtain some of the same information from the AHRS. Additionally, it is important that the data that the pilot is receiving is accurate. So if the INS is providing data, but the data is erroneous, the pilot should use the data obtained from the AHRS and ignore the INS data.  
         [0004]     Typically, when one of the sensors fails or provides erroneous data, the sensor and/or the associated avionic system is deactivated. For example, if the INS fails, the INS would be deactivated and the pilot would rely on data obtained from the AHRS. The pilot would not be able to use data from the INS again unless the pilot manually re-started the INS. Additionally, the pilot has no way of knowing if the INS has resumed providing reliable data until after re-starting the INS. Accordingly, a pilot will typically land the aircraft using the AHRS data and trouble shoot the INS once the aircraft is grounded.  
         [0005]     While the previous example is presented using the INS and the AHRS, other avionic systems and/or sensors may provide an overlap of information such that if one fails, the pilot still has access to some data. This redundancy of information provides for safer flights.  
         [0006]     Additionally, the redundancy of information may improve the accuracy of some avionic systems. For example, the aircraft may include both an INS and a global positioning satellite (GPS) receiver, or other radio frequency (RF) ranging system, such as Time Difference of Arrival (TDOA) and Galileo. Both the INS and the GPS receiver may provide estimates of the aircraft&#39;s position. In addition, the data from the GPS receiver may be used to calibrate the INS, while the GPS receiver may use the data from the INS to quickly re-establish tracking of a satellite in which the GPS receiver has temporarily lost contact. Thus, the integration of the INS and GPS receiver provides more accurate and robust data to the pilot.  
         [0007]     The integration of the INS and the GPS receiver may be described as loosely, closely, tightly, or deeply coupled. A loosely coupled system may be described as a stand-alone GPS receiver integrated with a stand-alone INS. The GPS receiver passes position, velocity, and time (PVT) information obtained from four satellites to the INS. The INS uses the PVT information to correct inertial errors that are commonly associated with INS operation. However, the GPS data passed to the INS becomes unusable when less than four satellites are available to the GPS receiver.  
         [0008]     A closely coupled system may be described as a stand-alone GPS receiver integrated with a stand-alone INS. However in a closely coupled system, in addition to the GPS receiver passing PVT information obtained from typically four satellites to the INS, the INS passes velocity, acceleration, and angular rate information to the GPS receiver. The GPS receiver can use this information when tracking satellites and to re-acquire a satellite signal that has been lost. However, as with the loosely coupled system, the GPS data passed to the INS becomes less usable when less than four satellites are available to the GPS receiver. For example, when only three satellites are available, GPS receivers typically continue to output horizontal position, but do not output valid altitude data.  
         [0009]     In a tightly coupled system, the GPS receiver provides pseudorange and/or deltarange data to the INS. The GPS receiver contains tracking loops for tracking data from multiple satellites. The tracking loops provide pseudorange and deltarange measurements to the INS. The pseudorange measurements are an output of a delay lock loop, which is used for tracking code phase, while the deltarange measurements are an output of a phase lock loop, which is used for tracking carrier phase. The pseudorange and deltarange measurements are used by a Kalman filter in the INS to calculate errors, which sends correction data to a navigation computation.  
         [0010]     In the tightly coupled system, the GPS receiver sends pseudorange and deltarange to the INS for all satellites that are being tracked. The INS may continue to use the data obtained from the GPS receiver even when fewer than four satellite signals are being tracked. The INS in a tightly coupled GPS/INS system can continue to use the GPS data with less than four available satellites because each pseudorange and deltarange measurement is an independent measurement.  
         [0011]     A deeply coupled system includes a GPS function and an Inertial Measurement Unit (IMU). The GPS function may be defined as the processing associated with computing the GPS data, while the IMU is generally described as the inertial sensing component of the INS, providing data directly to a computer. In a deeply coupled system, a stand-alone GPS receiver may not exist. For example, the functions of the GPS receiver may be resident in a single processor, along with the INS function.  
         [0012]     The computer performs the INS computations. However, in contrast with the tightly coupled system which uses pseudorange and deltarange data, measurements from all available satellites are processed by the Kalman filter using in-phase (I) and quadrature (Q) signals, which are calculated in the GPS function. The Kalman filter calculates the errors and sends correction data to the navigation computation and to the GPS function. The information sent to the GPS function includes commands to replica code generators to enable the GPS function to track the GPS satellites. This capability eliminates the need for standalone tracking loops in the GPS function. By combining information from multiple satellites and the inertial sensors, the deeply coupled system is able to track the satellites under higher interference or jamming levels.  
         [0013]     Accordingly, the more integrated the GPS/INS system becomes, the more robust the navigation system becomes. Additional benefits may be obtained by integrating data from other sensors. For example, data from the deeply integrated GPS/INS system may be used to calibrate the air data computer and a magnetometer. The air data computer and magnetometer may then be used as aids should GPS data become unavailable and the performance of the INS has degraded to a level where air data or magnetometer aiding will improve the accuracy of the navigation solution.  
         [0014]     It would be beneficial to use a deeply integrated GPS/INS system in a navigation system that is operable to automatically resume using data from a sensor that resumes providing reliable data. Accordingly, the pilot may operate the aircraft using the best data available from the avionic sensors.  
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0015]     Presently preferred embodiments are described below in conjunction with the appended drawing figures, wherein like reference numerals refer to like elements in the various figures, and wherein:  
         [0016]      FIG. 1  is a block diagram of a deeply integrated navigation system, according to an exemplary embodiment;  
         [0017]      FIG. 2  is a table summarizing available outputs in the navigation solution depending on the aiding sources available, according to an exemplary embodiment,  
         [0018]      FIG. 3  is a block diagram of a system processor, according to an exemplary embodiment;  
         [0019]      FIG. 4A  is a block diagram of a method of calculating GPS sensor data in an aiding mode, according to an exemplary embodiment;  
         [0020]      FIG. 4B  is a block diagram of a method of calculating GPS sensor data in a deep integration mode, according to an exemplary embodiment;  
         [0021]      FIG. 5  is a block diagram of mode logic, according to an exemplary embodiment; and  
         [0022]      FIG. 6  is a flow chart of resume logic, according to an exemplary embodiment.  
     
    
     DETAILED DESCRIPTION  
       [0023]      FIG. 1  is a block diagram of a deeply integrated navigation system  100 , according to an exemplary embodiment. The navigation system  100  includes an Inertial Measurement Unit (IMU)  102 , one or more GPS sensors  104 , a magnetometer  106 , an air data computer  108 , and a system processor  110 . The navigation system  100  may also include additional entities not shown in  FIG. 1 . In a preferred embodiment, the navigation system  100  may be provided in a single package.  
         [0024]     The IMU  102  may provide acceleration and angular rate data. Typically the IMU  102  relies on three orthogonally mounted acceleration sensors and three nominally orthogonally mounted inertial angular rate sensors, which can provide three-axis acceleration and angular rate measurement signals. Accordingly, the IMU  102  may include three accelerometers  112  and three gyroscopes  114 . The three accelerometers  112  may be any type of accelerometer, such as a force re-balance, resonating beam, or MEMS accelerometer. The three gyroscopes  114  may be any type of gyroscope, such as a ring laser or MEMS gyroscope. The three accelerometers  112  and three gyroscopes  114  may be packaged together with a processor and associated navigation software. Alternatively, self-contained IMU packages may also be used.  
         [0025]     Additionally, the IMU  102  may include inertial electronics  116 . The inertial electronics  116  may be used to convert the acceleration and angular rate data obtained by the IMU  102  into a digital representation of the data. The inertial electronics  116  may also provide compensation of the acceleration and angular rate data. This compensation may include compensation in which compensation coefficients or other characteristics of the compensation are updated by the system processor  110 . The combination of the IMU  102  and the system processor  110  may be referred to as an INS.  
         [0026]     The one or more GPS sensors  104  may be a GPS receiver, TDOA, Galileo, or any other RF ranging system. It is understood that the term GPS sensor as used in this specification includes any RF ranging system. The GPS sensor  104  may provide three-dimensional PVT information. Additionally, the GPS sensor  104  may provide pseudorange and deltarange (PR/DR) information and/or in-phase and quadrature (I &amp; Q) information. If multiple GPS sensors  104  are used with separate and suitably located antennas, then an estimate of heading may be computed using the information from the GPS sensor  104 . Typically, the GPS sensor  104  is used in conjunction with the INS to provide a more robust navigation solution. The INS, alone or in conjunction with other aids, may provide data to the aircraft when a satellite signal is temporarily lost due to interference. The GPS sensor  104  may use the INS data to quickly regain a lost satellite signal. Additionally, the INS may use GPS data for initialization, calibration, and/or aiding.  
         [0027]     The magnetometer  106  may detect the Earth&#39;s magnetic field. Data from the magnetometer  106  may be used to determine the heading of the aircraft. This information may be used to initialize the system or as an aid to the INS. The system processor  110  may use the heading information from the magnetometer  106  in combination with GPS PVT information, GPS PR/DR information, GPS I &amp; Q information, GPS-derived heading information, or inertial-derived heading information to provide an improved heading reading to the pilot. The INS, in conjunction with other aiding sensors, may also be used to calibrate the magnetometer  106 .  
         [0028]     The air data computer  108  may be used to calculate altitude, vertical speed, air speed, and a mach number of the aircraft. Other calculations may also be possible, such as air temperature. Pressure transducers within the air data computer  108  may be used to collect data. The system processor  110  may be used to convert the data collected by the pressure transducers and provide altitude, vertical speed, air speed, and mach number outputs. The INS may use outputs from the air data computer  108  for calculating altitude, as an aiding sensor to improve the overall navigation solution, for use in reversionary modes, and during initialization. The INS, in conjunction with other aiding sensors, may also be used to calibrate the air data computer  108 .  
         [0029]     If sufficient aiding information is not available and the inertial sensors do not accurately compute position and velocity sufficient to maintain an application&#39;s attitude requirements, then the navigation system  100  may limit its computations to data typically computed by an AHRS function. The AHRS function is typically defined as an inertial system that outputs only attitude, which includes pitch, roll, and heading. An AHRS function typically does not output position and velocity data. Additionally, errors in pitch, roll, and heading for the AHRS function are typically bounded by one or more aiding sources or the accelerometers  112  through the use of a slaving of the attitude to a gravity vector. The aiding sources or accelerometers  112  may be used in a slaving loop to improve the estimate of attitude. In this approach, the navigation system  100  may use a Kalman filter to correct the attitude based on the gravity vector. Thus, the AHRS function is not depicted in  FIG. 1 , because it does not exist as a separate function.  
         [0030]     The system processor  110  may receive data from the sensors, provide error correction, and provide as an output the navigation solution. The system processor  110  may include any combination of hardware, firmware, and/or software operable to receive the data, process the data, and calculate a navigation solution. The navigation solution may be a three-dimensional position, three-dimensional velocity, and three-dimensional attitude solution. Other avionics systems may use the navigation solution. For example, the aircraft&#39;s position may be displayed for the pilot on a head-up display.  
         [0031]     When in the navigation mode (e.g., when navigation is engaged), the type and quality of the navigation outputs may depend on the type and quality of data received by the system processor  110 . The type and quality of the navigation outputs may also depend on the performance of the inertial sensors. For example, the quality of the data may be better when the system is being aided by deep integration than when the system has only the heading aid enabled. For example, the navigation solution may include position, velocity, and attitude data if the aircraft is being aided by GPS PVT. As another example, the navigation solution may include just the attitude data if the aircraft is being aided by the magnetometer with leveling engaged. Additionally, the navigation system  100  may be in standby mode, in which case no navigation solution may be provided. Other modes may be possible, such as a test mode.  
         [0032]     The engage leveling capability is typically accomplished through one of two methods, the traditional method and the Kalman filter method. Traditionally, the leveling capability has been accomplished through conventional filtering techniques that simply slave an attitude matrix to level based on the assumption that the gravity vector is vertical. However, this can cause problems during coordinated turns in which the acceleration vector is not vertical. During these periods, the leveling loop may be temporarily disengaged. The fact that the aircraft may be in a coordinated turn may be based on the attitude of the aircraft. For example, the leveling loop may be disengaged when the roll angle is greater than five degrees.  
         [0033]     Another method of implementing the leveling capability is to perform the leveling through the Kalman filter located within the system processor  110 . The Kalman filter method may overcome or reduce the problem associated with the assumption that the gravity vector is vertical. The Kalman filter may contain additional states to model the leveling capability. The Kalman filter may use a comparison of a predicted and an actual acceleration component to calculate an attitude error. The attitude error may be used to perform the leveling.  
         [0034]      FIG. 2  is a table that summarizes the available outputs in the navigation solution depending on the aiding sources available. Attitude data is generally always available. However, position and velocity data availability may be dependant upon the type of aiding sources that are available and the performance of the inertial sensors. Leveling may be engaged when performance of the inertial sensors is not sufficient to meet the application&#39;s attitude requirements and an aiding source that does not compute position and velocity is enabled. The navigation system  100  may output position, velocity, and attitude for a period of time, and then may switch to attitude-only outputs based on the time or covariance values on the Kalman filter.  
         [0035]      FIG. 3  is a block diagram of a system processor  300 , according to an exemplary embodiment. The system processor  300  may be substantially the same as the system processor  110  depicted in  FIG. 1 . The system processor  300  may include an IMU compensation element  320 , a navigation computation element  322 , mode logic  324 , a Kalman filter  326 , and two numerical controlled oscillator (NCO) command generators  328 ,  330 .  
         [0036]     The number of NCO command generators in the system processor  300  may be determined by the number of GPS, Galileo, or other RF ranging sensors providing data to the Kalman filter  326 . Accordingly, more or less than two NCO command generators may be located in the system processor  300 . Alternatively, the NCO command generators  328 ,  330  may be located outside the system processor  300 . The system processor  300  may include additional elements not depicted in  FIG. 3  as well. In an alternative embodiment, the IMU compensation element  320  may be partially or completely located in the IMU  102 .  
         [0037]     The system processor  300  may receive a variety of data from a variety of sources. The system processor  300  may receive the following types of data: compensated or uncompensated gyroscope data, compensated or uncompensated acceleration data, magnetic data, air data, external velocity data  310 , external attitude data  312 , and GPS data, otherwise referred to as sensor data.  
         [0038]     The external velocity data  310  and the external attitude data  312  may be provided by other inertial navigation systems on the aircraft, such as an aircraft INS. The external velocity data  310  may indicate the velocity of the aircraft, including when the velocity of the aircraft is zero (i.e., stationary). Further, the external velocity data  310  may be calculated as a conventional velocity measurement or as a conventional change in position over a specified time period (e.g., the Kalman filter interval).  
         [0039]     The system processor  300  may receive sensor data from the following sensors: gyroscope sensors  302 , acceleration sensors  304 , magnetic sensors  306 , air sensors  308 , and two GPS sensors  316 ,  318 . While two GPS sensors are depicted in  FIG. 3  more or less than two GPS sensors may provide GPS data to the system processor  300 . The system processor  300  may receive additional data as well, such as data from an odometer. Additionally, the system processor may receive an indication of whether or not to engage leveling  314 . The indication of whether or not to engage level  314  may be based on time since an aid became unavailable or on the Kalman filter  326  covariances.  
         [0040]     The sensors  302 - 308 ,  316 , and  318  may sense the state of the vehicle. Gyroscope electronics  332  may convert data from the gyroscope sensors  302  into a digital representation of the gyroscope data prior to sending the gyroscope data to the system processor  300 . Likewise, accelerometer electronics  334  may convert data from the acceleration sensors  304  into a digital representation of the acceleration data prior to sending the acceleration data to the system processor  300 .  
         [0041]     The format and accuracy of data from the two GPS sensors  316 ,  318  may depend on whether the navigation system  100  is in a deeply integrated mode or a PVT/PR-DR aiding mode. Further, if the navigation system  100  is in the PVT/PR-DR aiding mode, the format and accuracy of the GPS sensor data may depend on whether the two GPS sensors  316 ,  318  are in a PVT mode or a PR/DR mode. The GPS sensor data is further described with reference to  FIG. 4 . Galileo or other RF ranging sensors may be used instead of or in conjunction with the GPS sensors  316 ,  318 .  
         [0042]     The system processor  300  may provide as an output a navigation solution. The navigation solution may be a three-dimensional position, three-dimensional velocity, and three-dimensional attitude solution. However, the exact navigation solution may depend on the operational mode of the aircraft. Additionally, the system processor  300  may provide GPS outputs. The GPS outputs may be the data obtained from the two GPS sensors  316 ,  318 , with or without additional processing from the system processor  300 . Additional outputs are also possible. Other avionics systems may use the navigation solution and the GPS outputs. For example, the aircraft&#39;s position may be displayed for the pilot on a head-up display.  
         [0043]     The IMU element  320  may receive the data from the gyroscope electronics  332  and the acceleration electronics  334 . The combination of the gyroscope electronics  332  and the acceleration electronics  334  may be substantially the same as the inertial electronics  116  as depicted in  FIG. 1 . The IMU element  320  may provide compensation that uses information from multiple sensors (e.g., three accelerometers and three gyroscopes) to compensate one or more of the inertial sensors. For example, the IMU element  320  may compensate for coning, sculling, and/or gravitational effects.  
         [0044]     Additionally, the Kalman filter  326  may provide estimates of gyroscope and accelerometer errors to the IMU element  320  and/or to the navigation solution. Inertial navigation systems experience drifts over time, which causes errors in the position, velocity, and attitude solutions. The errors may be caused by gyroscope drift, accelerometer bias, scale factor errors, and other error sources. The navigation corrections provided by the Kalman filter  326  may correct the errors in the navigation solution caused by these errors. The IMU element  320  may provide compensated IMU data to the navigation computation element  322 .  
         [0045]     The navigation computation element  322  may be software capable of blending the IMU data received from the IMU element  320  and the navigation corrections provided by the Kalman filter  326  to produce a navigation solution. The navigation computation element  322  may calculate the navigation solution by numerically solving Newton&#39;s equations of motion using the data received from the IMU element  320  and the Kalman filter  326 . The navigation solution may be referenced to a navigation coordinate frame. Possible navigation coordinate frames include earth centered inertial (ECI), earth centered earth fixed (ECEF), local level with axes in the directions of north, east, down (NED), and local level with a wander azimuth.  
         [0046]     Additionally, the navigation computation element  322  may provide the navigation solution to the Kalman filter  326 . The Kalman filter  326  may use the navigation solution to calculate estimates of future calculated navigation solutions. In this manner, the Kalman filter  326  may provide a recursive method of calculating the errors in the sensors used to compute the navigation solution.  
         [0047]     The navigation computation element  322  may also receive an input from the mode logic  324 . The mode logic  324  may indicate whether or not the navigation computation element  322  should provide a navigation solution. When the mode logic  324  indicates that the navigation computation element  322  should not provide a navigation solution, the navigation system  100  may be in standby mode. Alternatively, when the mode logic  324  indicates that the navigation computation element  322  should provide a navigation solution, the navigation system  100  may be in navigation mode.  
         [0048]     The mode logic  324  may be any combination of hardware, firmware, and/or software that is operable to determine whether a navigation solution should be provided, and if so, what type of navigational solution should be provided. For example, the aircraft may be in stand-by mode and not require a navigation solution. However, if the aircraft is in navigation mode, the mode logic  324  may determine whether the navigation solution should be based on the deeply integrated mode or the aiding mode. Further, if the aircraft is in the aiding mode, the mode logic may determine whether to use the PVT mode or the PR/DR mode. Additional information regarding the mode logic  324  is described with reference to  FIG. 5 .  
         [0049]     The Kalman filter  326  may be any combination of hardware, firmware, and/or software operable to provide an estimate. Kalman filters are well known in the art for use in providing correction data to a navigation computation element to provide a more accurate navigation solution. The Kalman filter  326  may receive data from the sensors and estimate navigation corrections of the aircraft&#39;s position, velocity, and/or attitude. The Kalman filter  326  may estimate navigation corrections using a model of the INS error dynamics. The Kalman filter  326  may provide the estimate to the IMU element  320 , the navigation computation element  322 , and the NCO command generators  328 ,  330 .  
         [0050]     The NCO command generators  328 ,  330  may receive parameter estimates from the Kalman filter  326  and the navigation solution from the navigation computation element  322  when the navigation system  100  is in the deep integration mode. The NCO command generators  328 ,  330  might not be used when the navigation system  100  is in the aiding mode. The NCO command generators  328 ,  330  may generate code and carrier command signals. The code and carrier command signals may be used in calculating the GPS sensor data in the deep integration mode. The function of the NCO command generators  328 ,  330  is further described with reference to  FIG. 4B .  
         [0051]      FIG. 4  is a block diagram of a method of calculating the GPS sensor data.  FIG. 4A  is a block diagram of a method of calculating the GPS sensor data in the aiding mode, while  FIG. 4B  is a block diagram of a method of calculating the GPS sensor data in the deep integration mode. In both the aiding mode and the deep integration mode, the GPS sensor  104  detects and receives data from orbiting satellites. The orbiting satellites broadcast a continuous series of radio signals, referred to as GPS signals, which are detected by the GPS sensor  104 . The radio signals contain information regarding the known position of the satellites. Based on the reception of the radio signals, the GPS sensor  104  is able to estimate the distance to each satellite, and the relative velocity of the satellites with respect to the GPS sensor  104 .  
         [0052]     The satellites currently broadcast on two frequencies. The two frequencies are referred to as L1 (1575.42 MHz) and L2 (1227.6 MHz). During the propagation of the GPS signals through the atmosphere, a loss in signal strength occurs. Accordingly, the GPS signals may be processed into usable signals. In the future, additional frequencies may be added to GPS satellites, or satellites using a different signal structure, such as Galileo, may become available. The techniques described herein may be used with Galileo, RF ranging systems, and these new frequencies and signal structures.  
         [0053]     The GPS signals received by the GPS sensor  104  may be processed by the RF/IF and sampling element  402 . The RF/IF and sampling element  402  may be located in the GPS sensor  104 . The GPS signal may pass through a high-pass filter that rejects all parts of the signal that are not within the L1 or L2 bandwidths, producing a radio frequency (RF) signal. The RF signal may be down converted to an intermediate frequency (IF). The IF signal may be sampled to convert the IF signal into a digital form. The digital form of the IF signal may be provided to a bank of correlators  404  or for use in a software defined GPS radio. A software defined GPS radio may process data at frequencies below IF, at IF, or higher than IF, including at the original RF signal frequency.  
         [0054]     The bank of correlators  404  may be used to determine if a bank of replica generators  408  is generating a replica that is identical in structure, time, and frequency to the received GPS satellite signal. The result of this correlation may be a set of signals that includes an in-phase signal and a signal that is 90 degrees out of phase from a reference signal located within the GPS sensor  104 . The bank of correlators  404  and the bank of replica generators  408  may be located within the GPS sensor  104 .  
         [0055]     The number of satellites being tracked may determine the number of correlators in the bank of correlators  404 . A bank of correlators  404  may include three correlators for each satellite tracked to detect early, prompt, and late GPS signal transmissions. The bank of correlators  404  may provide outputs I E , I P , I L  (e.g., early, prompt, and late in-phase signals) and Q E , Q P , Q L  (e.g., early, prompt, and late quadrature signals) for each satellite tracked. The I &amp; Q data is often summed prior to use by other functions. Typically the data is summed for 20 msec, thus generating data samples at a 50 Hz rate. Of course, other data rates may be used.  
         [0056]     As depicted in  FIG. 4A , the bank of correlators  404  may provide the I and Q components of the GPS signal to a GPS tracking loop  406  in the aiding mode. The GPS tracking loop  406  may be located within the GPS sensor  104  or the system processor  110 . The GPS tracking loop  406  may include both a carrier tracking loop and a code tracking loop. The tracking loops provide pseudorange and deltarange measurement outputs. The pseudorange measurements are an output of a delay lock loop, which is used for tracking code phase, while the deltarange measurements are an output of a phase lock loop, which is used for tracking carrier phase. The pseudorange and deltarange measurements are used by the Kalman filter  326  to estimate navigation corrections when the navigation system  100  is in the aiding mode.  
         [0057]     The pseudorange and deltarange measurements may also be provided to the bank of replica generators  408  in the aiding mode. The measurements are scaled and time-phased to adjust the replica generators  408  to enable the replica generator  408  to continue to generate a replica that is identical in data, time, and frequency to the GPS signal received from the satellite.  
         [0058]     As depicted in  FIG. 4B , the bank of correlators  404  may provide the I and Q components of the GPS signal to a measurement and pre-processing element  410 , instead of the GPS tracking loop  406 , in the deep integration mode. The measurement and pre-processing element  410  may be located within the system processor  110 . The measurement and pre-processing element  410  may calculate a code error estimate (Ê τ ) and a carrier error estimate (Ê φ ) based on the early, prompt, and late I and Q components of the GPS signal, as provided by the bank of correlators  404 . The code and carrier error estimates may be calculated as follows.  
             E   =         I   E   2     +     Q   E   2                 (     Equation   ⁢           ⁢   1     )               L   =         I   L   2     +     Q   L   2                 (     Equation   ⁢           ⁢   2     )                   E   ^     τ     =       E   -   L       2   ⁢     (     E   +   L     )                 (     Equation   ⁢           ⁢   3     )                   E   ^     ϕ     =       tan     -   1       ⁡     (       Q   P       I   P       )               (     Equation   ⁢           ⁢   4     )             
 
 The code and carrier error estimates may be converted from the 50 Hz data rate to a 10 Hz data rate prior to sending the error estimates to the Kalman filter  326 . Other data rates may be used. 
 
         [0059]     The Kalman filter  326  may receive the code and carrier error estimates from the measurement and pre-processing element  410  and estimate navigation corrections of the aircraft&#39;s position, velocity, and/or attitude, as well as the GPS clock, clock drift, and other parameters associated with the GPS clock. The Kalman filter  326  may include 43 state vector elements. The state vector elements may include navigation errors (e.g., position, velocity, and attitude), GPS oscillator errors, range bias states, and inertial sensor errors. More or less than 43 vector elements may be used. For example, not all range bias state vectors may be used or additional range bias states may be added to enable the simultaneous tracking of more or less GPS satellites.  
         [0060]     The Kalman filter estimate of navigation corrections may be transmitted to the NCO command generators  328 ,  330  at a 10 Hz data rate. Other data rates may be used. The NCO command generators  328 ,  330  may also receive the navigation solution as provided by the navigation computation element  322 . In a preferred embodiment, a 100 Hz data rate is used; however, other data rates may be used. The data received from the Kalman filter  326  and the navigation computation element  322  may be used to compute and estimate the satellite range over an interval. The interval is the period of time between when the NCO command generators  328 ,  330  are updated. In a preferred embodiment the update rate is 50 Hz, but other rates may be used.  
         [0061]     For each satellite tracked, the satellite&#39;s pseudorange at a start time ({circumflex over (ρ)} start ) is computed. At the end of the interval, the satellite&#39;s pseudorange at a stop time ({circumflex over (ρ)} stop ) is computed. The data received from the Kalman filter  326  and the navigation computation element  322  may be used to compute the pseudorange at the start time and the stop time. The pseudorange calculation at the stop time may then become the pseudorange calculation at the start time for the next interval. For each satellite tracked, the NCO command generators  328 ,  330  may calculate a code command (Code_cmd) and a carrier command (Carrier_cmd) that may be used to update the NCO command generators  328 ,  330 . The code and carrier commands may be calculated as follows. 
 
Δ{circumflex over (ρ)}={circumflex over (ρ)} stop −{circumflex over (ρ)} start   (Equation 5) 
 
Δ{circumflex over (τ)}=Δ{circumflex over (ρ)}/λ C ; λ C ˜29.3 m  (Equation 6) 
 
Δ{circumflex over (φ)}=Δ{circumflex over (ρ)}/λ L ; λ L ˜0.19 m for L 1   (Equation 7) 
 
Code_cmd=Δ{circumflex over (τ)}/0.02 sec  (Equation 8) 
 
Carrier_cmd=Δ{circumflex over (φ)}/0.02 sec  (Equation 9) 
 
         [0062]     The code and carrier NCO commands may be transmitted to the bank of replica generators  408  at a 50 Hz data rate. Other data rates may be used. The code and carrier commands may be used by the NCO command generators  328 ,  330  to adjust the bank of replica generators  408 . The update may enable the GPS sensor  104  to track the GPS satellites under higher interference or jamming levels.  
         [0063]      FIG. 5  is a block diagram of mode logic  500 , according to an exemplary embodiment. The mode logic  500  may be substantially the same as the mode logic  324  depicted in  FIG. 3 . The mode logic  500  may include a plurality of switches  502 - 522 . The switches  502 - 522  may be either hardware or software switches, but are preferably software switches. Eleven switches are depicted in  FIG. 5 ; however, more or less than eleven switches may be used in the mode logic  500 . The number of switches in the mode logic  500  may be related to the amount of sensor data collected by the navigation system  100  and by the number of operational modes used by the navigation system  100 .  
         [0064]     The navigate switch  502  may determine the mode of the navigation system (e.g., standby or navigate). The switches  504 - 518  may determine which data the Kalman filter  326  will use to calculate corrections to the navigation solution. The activate change in attitude switch  520  may be used to determine whether leveling should be engaged as described with reference to  FIG. 1 .  
         [0065]     The navigate switch  502  may be in the “open” or disabled position indicating that the navigation system  100  is in the standby mode. In the standby mode, a navigation solution may not be provided. Both the navigation computation element  322  and the Kalman filter  326  may be disabled. Accordingly, the position of the other switches  504 - 522  may be irrelevant to the operation of the navigation system  100 . If the navigate switch  502  is in the “closed” or enabled position, the navigation computation  322  and the Kalman filter  326  may be enabled.  
         [0066]     When the navigate switch  502  is enabled, the position of the other switches  504 - 522  may be relevant. Each of the switches  504 - 522  may operate independently from each other. Accordingly, the mode logic  500  may determine which flight data to use (e.g., air data, velocity, attitude) and whether to operate in the deep integration mode or the aiding mode. The mode logic  500  may determine which aiding data to use based on which sensors are providing accurate flight data. More specifically, the mode logic  500  may determine which flight data to use based on the results of the resume logic  600  described with reference to  FIG. 6 .  
         [0067]     When the activate deep integration switch  522  is enabled, the navigation system  100  may operate in the deep integration mode. When the activate deep integration switch  522  is disabled, the navigation system may operate in the aiding mode, which includes two sub-modes: PVT and PR/DR aiding. The activate PVT GPS switch  504  may be enabled for operating in the PVT aiding mode, while the activate PR/DR GPS switch  506  may be enabled for operating in the PR/DR aiding mode.  
         [0068]     When the activate GPS heading switch  508  is enabled, GPS heading data may be used by the Kalman filter  326  to calculate corrections to the navigation solution. The GPS heading data may be calculated by conventional means. For example, the GPS heading data may be calculated as described in commonly assigned U.S. Pat. Nos. 5,917,445; 6,088,653; and 6,114,988; which are fully incorporated herein by reference.  
         [0069]     When the activate magnetometer switch  510  is enabled, data from the magnetometer  106  may be used by the Kalman filter  326  to calculate corrections to the navigation solution. When the activate air data switch  512  is enabled, data from the air data computer  108  may be used by the Kalman filter  326  to calculate corrections to the navigation solution. When the activate velocity switch  514  and/or the activate attitude switch  516  is enabled, data from other inertial navigation systems on the aircraft may be used by the Kalman filter  326  to calculate corrections to the navigation solution. When the activate odometer aiding switch  518  is enabled, data from an odometer may be used by the Kalman filter  326  to calculate corrections to the navigation solution. The odometer reading may be used by the Kalman filter  326  to calculate position change over time (i.e., velocity).  
         [0070]      FIG. 6  is a flow chart of resume logic  600 , according to an exemplary embodiment. The resume logic  600  may be a software program located within the system processor  110 . The resume logic  600  may apply to all aiding sources depicted in  FIG. 1  (e.g., three accelerometers, three gyroscopes, the GPS sensor, the magnetometer, and the air data computer). The resume logic  600  may be implemented when no errors have been detected in previous data measurements. Alternatively, the resume logic  600  may be implemented after data from one or more sensors has been previously found to be erroneous.  
         [0071]     At block  602 , a measurement is taken. The measurement may be the PVT or PR/DR data measured by the GPS sensor  104 , the magnetic field data measured by the magnetometer  106 , the air data measured by the air data computer  108 , or any other sensor data measurement.  
         [0072]     At block  604 , the measurement is checked for validity. The sensor may provide one or more validity bits to the system processor  110 . The validity bit may be used to determine if the data from the sensor measurement is new. For example, system processor  110  may determine that the data had been previously transferred to the system processor  110  (e.g., old data), which would be an indication that the data is not valid or has been previously processed. The previously transferred data may or may not have been previously determined to have errors. The validity bits may provide other information indicating whether or not the data is valid as well.  
         [0073]     At block  606 , if the measurement is not valid based on the validity bit check, then the measurement is not used by the Kalman filter to calculate corrections to the navigation solution.  
         [0074]     At block  608 , the measurement is checked to determine if the measurement is self-consistent. The system processor  110  may have data regarding the operational capacity of the aircraft. As such, the system processor  110  may know the valid range of data from each of the sensors. For example, the air data computer  108  may provide the system processor  110  with the air speed of the aircraft. If the air speed data is greater than the maximum speed that the aircraft can travel, the system processor  110  may determine that the air speed data is erroneous.  
         [0075]     At block  606 , if the measurement is not self-consistent, then the measurement is not used to calculate corrections to the navigation solution. Using the example provided above, if the air speed data is erroneous, the air speed data from the air data computer  108  may not be used by the Kalman filter to calculate corrections to the navigation solution.  
         [0076]     At block  610 , the measurement is compared to a Kalman filter prediction  612 . The Kalman filter  326  may be used to predict the position, velocity, and/or attitude of the aircraft. The Kalman filter  626  may be located in the system processor  110  and receive recursive data measurements as shown at block  614 . If the data is not within a certain percentage of the Kalman filter prediction  612 , the data may contain an error. For example, a deviation of 3σ from the Kalman filter prediction  612  may indicate that the data is erroneous.  
         [0077]     At block  606 , if the measurement does not agree with the Kalman filter prediction  612 , then the measurement is not used by the Kalman filter to calculate corrections to the navigation solution.  
         [0078]     At block  614 , the measurement may be used by the Kalman filter  326  to provide corrections to the navigation solution. The measurement may be used whether or not the sensor previously provided erroneous data. For example, if the air data computer previously provided erroneous air speed data and the current air speed data has passed a validity check, a self-consistent check, and a Kalman filter prediction check, the current air speed data may be used to calculate the corrections by way of the Kalman filter  326 , which may be provided to the navigation computation element  322  to calculate the navigation solution. This data may be sent to the Kalman filter  326  as depicted at block  612 .  
         [0079]     The resume logic  600  may allow the navigation system  100  to automatically determine that a sensor that had previously provided erroneous data is currently providing valid data. The results of the resume logic  600  may be used by the mode logic  500  to determine which sensor data should be provided to the Kalman filter  326 . Thus, the navigation system  100  may provide the most accurate navigational solution to the pilot at all times. The resume logic  600  in combination with the mode logic  500  may eliminate the previous requirement for the pilot to manually re-start a sensor that has failed or provided erroneous data.  
         [0080]     It should be understood that the illustrated embodiments are exemplary only and should not be taken as limiting the scope of the present invention. While the invention has been described with reference to an aircraft, the invention may be applicable to other vehicles or devices, such as space vehicles, missiles, and pipeline inspection gear. The claims should not be read as limited to the described order or elements unless stated to that effect. Therefore, all embodiments that come within the scope and spirit of the following claims and equivalents thereto are claimed as the invention.

Technology Classification (CPC): 6