Abstract:
System and methods of increasing reliability of determined location information by using two integration filters are provided. An exemplary embodiment integrates inertial navigation system information and global navigation satellite system (GNSS) information in a real time Kalman filter; determines a real time location of the aircraft with the real time Kalman filter based upon the INS information and the GNSS information; delays the GNSS information by an interval; integrates the INS information and the delayed GNSS information in a delay Kalman filter; determines a predictive location of the aircraft with the delay Kalman filter based upon the INS information, the delayed GNSS information, and the interval; and in response to an inaccuracy of the real time location determined from the real time Kalman filter, selects the predictive location determined from the delay Kalman filter as a new real time location of the aircraft.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    An aircraft inertial navigation system is typically integrated with a global navigation satellite system (GNSS) receiver to provide reliable and continuous location information to assist the aircraft in performing various maneuvers, particularly when the aircraft is landing. In such systems, the information from the inertial navigation system and the information from the GNSS receiver are integrated together using a Kalman filter. The Kalman filter generates high integrity and continuous location information. The use of Kalman filters, and other types of filters, are well known. 
         [0002]    For example, U.S. Pat. No. 5,184,304 to Huddle, entitled “Fault-Tolerant Inertial Navigation System,” U.S. Pat. No. 5,583,774 to Diesel, entitled “Assured-Integrity Monitored-Extrapolation Navigation Apparatus,” and U.S. Pat. No. 6,549,829 to Anderson et. al, entitled “Skipping Filter For Inertially Augmented Landing System,” all of which are incorporated by reference herein in their entirety, are examples of such integrated navigation systems that use filters to provide high integrity and continuous location information. 
         [0003]    The reliability of the location information provided by the GNSS receiver is a critical issue. If the location information provided by the GNSS receiver is inaccurate, or if the information is not available from the GNSS receiver, then sufficiently accurate location information may not be available to assist the aircraft in its performing various maneuvers. An exemplary fault mode may occur if one or more of the electronic components of the GNSS satellite, or the GNSS receiver fails, such that the location information becomes inaccurate. In such situations, location information from the GNSS receiver should not be used. Another exemplary fault mode may occur if the clock times used by the GNSS satellites drift and become sufficiently inaccurate so as to render the output location information inaccurate. Or, one or more of the GNSS satellites may not be in its expected orbital location, such as when a GNSS satellite moves in an unexpected or unknown manner from its designated orbital route, such that the determined location information based upon a signal received from the GNSS satellite is incorrect. Other fault modes may occur if the information from the GNSS system is not available. An exemplary fault mode in this case is one or more GNSS satellites signals are blocked from being received by the GNSS receiver, such that not enough GNSS signals are available for determining the user location information. 
       SUMMARY OF THE INVENTION 
       [0004]    Systems and methods of increasing reliability of determined location information by using two integration filters are provided. An exemplary embodiment has a real time filter that is operable to receive first location information from an inertial navigation system, that is operable to receive second location information from a global navigation satellite system (GNSS) receiver, and that is operable to determine real time location information of the aircraft based upon integration of the first location information and the second location information; a time interval unit operable to receive the second location information from the GNSS receiver, and operable to delay the received second location information by a time interval; a delay filter communicatively coupled to the time interval unit, operable to receive from the inertial navigation system the first location information, operable to receive from the time interval unit the second location information delayed by the interval, and operable to determine predictive location information of the aircraft based upon the first location information, the second location information delayed by the interval, and the interval; and an airborne GNSS device that is operable to identify a corruption of the second location information, and in response to determining that the second location information is corrupted, that is further operable to cause the predictive location information to be output to at least one of a flight management system and an auto pilot system. 
         [0005]    In accordance with further aspects, an exemplary embodiment integrates inertial navigation system information and GNSS information in a real time Kalman filter; determines a real time location of the aircraft with the real time Kalman filter based upon the inertial navigation system information and the GNSS information; delays the GNSS information by an interval; integrates the inertial navigation system information and the delayed GNSS information in a delay Kalman filter; determines a predictive location of the aircraft with the delay Kalman filter based upon the inertial navigation system information, the delayed GNSS information, and the interval; and in response to an inaccuracy of the real time location determined from the real time Kalman filter, selects the predictive location determined from the delay Kalman filter as a new real time location of the aircraft. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS  
         [0006]    Preferred and alternative embodiments are described in detail below with reference to the following drawings: 
           [0007]      FIG. 1  is a block diagram of an embodiment of a navigation integrity system; and 
           [0008]      FIG. 2  is a block diagram of an alternative embodiment of the navigation integrity system. 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
       [0009]      FIG. 1  is a block diagram of an embodiment of a navigation integrity system  100 . The exemplary embodiment of the navigation integrity system  100  receives location information from an inertial navigation system  102  and an airborne global navigation satellite system (GNSS) device  104 . The navigation integrity system  100  integrates the received GNSS location information with the inertial navigation system  102  such that a highly accurate and reliable location of the aircraft is communicated to the flight management and/or auto pilot systems  106 . It is appreciated that the location information determined by the navigation integrity system  100  may be provided to other aircraft systems. 
         [0010]    Integrity of the location information provided by the airborne GNSS device  104  should be maintained if the navigation integrity system  100  is to provide accurate location information to the flight management and/or auto pilot systems  106 . The provided accurate location information aids in the various maneuvers performed by the aircraft, such as, but not limited to, landing. 
         [0011]    The airborne GNSS device  104  includes a GNSS receiver  108  that determines location information based upon the reception of signals transmitted by a plurality of GNSS satellites. As noted above, the GNSS receiver  108  may not provide accurate information if various fault modes occur. An exemplary fault mode may occur when a GNSS satellite is transmitting a signal with an incorrect time (caused by time clock drifting). Yet another exemplary fault mode may occur when the GNSS satellite is not in its designated orbital position (caused by movement of the GNSS satellite in an unexpected or unknown manner from its designated orbital route). Other exemplary fault modes may occur when a component of the on-board airborne GNSS device  104  and/or the on-board GNSS receiver  108  fails such that the location information is no longer output by the on-board airborne GNSS device  104 . Accordingly, embodiments of the navigation integrity system  100  are operable to determine that the location information output by the airborne GNSS device  104  is inaccurate and/or if the information is no longer available. 
         [0012]    Fault modes may be detected by a satellite based augmentation system (SBAS) which monitors the integrity of the information transmitted in a GNSS satellite signal. The SBAS (not shown), typically residing in geo-synchronous GNSS satellites, indicates presence of an error in the GNSS satellite signal or some other problem in the GNSS satellite. A fault indication signal, the SBAS integrity information  110 , is received by the airborne GNSS device  104  and indicates the inaccuracy of the real time location determined by the GNSS receiver  108 . The SBAS integrity information  110  may be separately transmitted or integrated into the GNSS signal. The SBAS integrity information  110  is then communicated from the GNSS satellite. The SBAS integrity information  110  may be received by the GNSS receiver  108  directly (such as when the fault indication signal generated by the SBAS integrity monitor  110  is integrated into the GNSS satellite signal). Under current available technologies, the SBAS integrity information  110  is received by the airborne GNSS device  104  with no more than thirty seconds delay. (It is appreciated that the time to receive the SBAS integrity information  110  may become less with improving SBAS technologies.) 
         [0013]    Alternatively, or additionally, fault modes occurring at the GNSS satellite may be detected by a ground based augmentation system (GBAS) which monitors the integrity of the information transmitted in the GNSS satellite signal. The GBAS (not shown), typically resides in a land-based installation that monitors GNSS satellite signals. The GBAS indicates presence of an error or fault in the information transmitted in the GNSS satellite signal or may indicate other problems with the GNSS satellite, such as when a satellite control center indicates a problem with the GNSS satellite. GBAS integrity information  112  is communicated from the GBAS to the airborne GNSS device  104  on the aircraft and indicates the inaccuracy of the real time location determined by the GNSS receiver  108 . Under current available technologies, the GBAS integrity information  112  is received by the airborne GNSS device  104  with no more than six seconds delay. (It is appreciated that the time to receive the GBAS integrity information  112  may become less with improving GBAS technologies.) 
         [0014]    The exemplary embodiment of the navigation integrity system  100  illustrated in  FIG. 1  has a processor system  114 , a real time Kalman filter  116 , a delay Kalman filter  118 , a memory  120 , and a time interval unit  122 . In the exemplary embodiment, integrity logic  124  resides in the memory  120 . The integrity logic  124  is retrieved and executed by the processor system  114 . The operational functions of the integrity logic  124  are described below. Alternatively, the functions may be implemented as firmware, or a combination of firmware and software, by alternative embodiments of the navigation integrity system  100 . 
         [0015]    As noted above, the inertial navigation system  102  generates real time location information. The inertial navigation system  102  is a well-known system that determines the real time location information based on various sensors  126  in the aircraft, such as acceleration sensors (accelerometers), rotation sensors (gyros), heading sensors (magnetometer), altitude sensors (barometric altitude), and the like (not shown). Based on a previous known location (which is presumed accurate), the inertial navigation system  102  computes the real time location based on the information provided by the above-described sensors. 
         [0016]    The real time Kalman filter  116  receives location information from the inertial navigation system  102  and the airborne GNSS device  104 . The received location information is integrated together by the real time Kalman filter  116  to generate accurate real time location information (assuming that location information received from the inertial navigation system  102  and the airborne GNSS device  104  are both accurate within design accuracy thresholds). The real time location information is communicated from the real time Kalman filter  116  to the processor system  114 . The processor system  114  communicates real time location information to the flight management and/or auto pilot systems  106  to aid in the various maneuvers performed by the aircraft, such as, but not limited to, landing. (It is appreciated that the real time location information determined by the real time Kalman filter  116  may be provided to other aircraft systems.) 
         [0017]    The time interval unit  122  also receives location information from the airborne GNSS device  104 . A predefined time interval is added to the information received from the airborne GNSS device  104 , thereby resulting in a time delay corresponding to the interval when the location information is communicated from the time interval unit  122 . 
         [0018]    In an exemplary embodiment that is monitoring the GBAS integrity information  112 , the time interval is approximately equal to six seconds. Six seconds corresponds to a time period that is sufficient for verification of the integrity of the location provided by the airborne GNSS device  104  based upon the GBAS integrity information  112 . In another exemplary embodiment that is monitoring the SBAS integrity information  110 , the time interval is approximately equal to thirty seconds. Thirty seconds corresponds to a time period that is sufficient for verification of the integrity of the location provided by the airborne GNSS device  104  based upon the SBAS integrity information  110 . Any suitable time interval may be used by the various embodiments of the navigation integrity system  100  corresponding to the times associated with receiving and/or processing the SBAS integrity information  110  and/or GBAS integrity information  112 . 
         [0019]    The delay Kalman filter  118  receives the time delayed location information from the time interval unit  122  and location information from the inertial navigation system  102 . This location information is integrated together by the delay Kalman filter  118  to generate accurate location information that is delayed by the time period delay (assuming that received location information from the inertial navigation system  102  and the airborne GNSS device  104  are both accurate within the design accuracy thresholds). 
         [0020]    Then, the delay Kalman filter  118  computes a predictive location from the time delayed location information received from the time interval unit  122  and location information received from the inertial navigation system  102 . That is, the delay Kalman filter  118  receives the time delayed location information from the time interval unit  122  and location information from the inertial navigation system  102 , integrates this location information, and then generates a predictive location for the aircraft. The predictive location information is communicated from the delay Kalman filter  118  to the processor system  114 . 
         [0021]    Embodiments of the navigation integrity system  100  are operable to assess the integrity of the real time location information provided by the real time Kalman filter  116 . In the event that the navigation integrity system  100  determines that the real time location information provided by the real time Kalman filter  116  has become corrupted (no longer accurate within acceptable tolerances), the navigation integrity system  100  provides the predictive location information from the delay Kalman filter  118  to the flight management and/or auto pilot systems  106 . (It is appreciated that the predictive location information determined by the delay Kalman filter  118  may be provided to other aircraft systems.) 
         [0022]    Because the time interval introduced by the time interval unit  122  is preferably greater than the time required for determining that the real time location information provided by the real time Kalman filter  116  has become corrupted, embodiments of the navigation integrity system  100  are able to change over from the corrupted real time location information to the predictive location information such that the flight management and/or auto pilot systems  106  continuously receive sufficiently accurate location information for purposes of navigation. That is, since the location information determined by the GNSS receiver  108  was accurate when originally collected (at the earlier time which corresponds to the delay time), the predictive location information determined by the delay Kalman filter  118  (based on the altitude, speed and heading in effect at the time that the delayed location information was determined) will be sufficiently accurate for navigation of the aircraft. Accordingly, the delay Kalman filter  118  uses the altitude, speed and heading of the aircraft, the time interval, the time delayed location information and the information from the various sensors  126  in the aircraft to accurately predict the location of the aircraft. 
         [0023]    For example, an exemplary embodiment of the navigation integrity system  100  uses a six second delay time when the GBAS integrity information  112  is used to assess integrity of the location information determined by the GNSS receiver  108 . Assume that the real time location information provided by the real time Kalman filter  116  is valid at a current time, t 0 . The predictive location information determined by the delay Kalman filter  118 , based upon information collected six seconds earlier (t 0-6 ), and forward predicted by the six second interval, is also accurate (assuming that the information collected six seconds ago was accurate). 
         [0024]    Then, assume that at some time i the location information determined by the GNSS receiver  108  (at t i  seconds) has become corrupted (no longer accurate within acceptable tolerances). The flight management and/or auto pilot systems  106  are then provided the predictive location information determined by the delay Kalman filter  118 , which was determined based on information collected six seconds earlier (at t i-6  seconds, which is six seconds after the current time t i  seconds). Accordingly, the flight management and/or auto pilot systems  106  have accurate location information so that the aircraft may perform its various maneuvers, such as, but not limited to, landing. 
         [0025]      FIG. 2  is a block diagram of an alternative embodiment of the navigation integrity system  100 . This embodiment includes a switch unit  202  that is operable to communicate the real time location information determined by the real time Kalman filter  116  or the predictive location information determined by the delay Kalman filter  118 . If the the SBAS integrity information  108  and/or GBAS integrity information  110  indicates that the real time location information determined by the GNSS receiver  108  is accurate, then the switch unit  202  communicates the real time location information from the real time Kalman filter  116  to the flight management and/or auto pilot systems  106 . On the other hand, if the SBAS integrity information  108  and/or GBAS integrity information  110  indicates that real time location information determined by the GNSS receiver  108  is corrupt, then the switch unit  202  communicates the predictive location information determined by the delay Kalman filter  118  to the flight management and/or auto pilot systems  106 . 
         [0026]    In a preferred embodiment, the switch unit  202  is controlled by the processor system  114 . Alternatively, the switch unit  202  may be controlled by the by another system device, such as, but not limited to, the airborne GNSS device  104 . The switching functions performed by the switch unit  202  may be implemented with any suitable type of electronic, solid state, or firmware type switching device or means commonly employed in the art. For example, a processor-based switch unit  202  would be implemented using a combination of software and firmware using components and methods commonly employed in the art of switching electrical devices. 
         [0027]    In an alternative embodiment, both the SBAS integrity information  108  and the GBAS integrity information  110  are monitored. In the event that the SBAS integrity information  108  indicates that the GNSS location information has been corrupted, a time interval corresponding to the SBAS integrity information  108  (here, the exemplary thirty seconds) is used. Thus, if the SBAS integrity information  108  indicates a corruption, the predictive location information determined by the delay Kalman filter  118  is based upon a time interval associated with the SBAS integrity information  108 . Alternatively, in the event that the GBAS integrity information  110  indicates that the GNSS location information has been corrupted, a time interval corresponding to the GBAS integrity information  110  (here, the exemplary six seconds) is used. Thus, if the GBAS integrity information  110  indicates a corruption, the predictive location information determined by the delay Kalman filter  118  is based upon a time interval associated with the GBAS integrity information  110 . 
         [0028]    In such an embodiment, the time interval unit  122  may be have two buffers, or other suitable means, for storing location information provided by the airborne GNSS device  104 . That is, location information corresponding to a first time interval (corresponding to the time interval associated with the SBAS integrity information  108 ) and a second time interval (corresponding to the time interval associated with the GBAS integrity information  110 ) are separately stored. 
         [0029]    In an alternative embodiment, the time interval unit  122  and the delay Kalman filter  118  are integrated into a single device. The delay may be implemented as software, firmware, or a combination of both. 
         [0030]    In an alternative embodiment, the filtering functions are implemented in another suitable integration filter type. For example, a skipping filter may be used. In other embodiments, the functionality of the Kalman filter is implemented as software, firmware, or a combination of software and firmware. Any suitable device or process that is operable to integrate location information generated by an inertial navigation system  102  and an airborne GNSS device  104  (or a GNSS receiver  108 ) may be used. 
         [0031]    In another embodiment, the predictive location and the real time location are compared by the processor system  114 . An inaccuracy of the real time location is determined when a difference between the real time location and the predictive location is different by a threshold distance. 
         [0032]    The above-described embodiments determined a corruption of the GNSS location information based upon the SBAS integrity information  108  and/or the GBAS integrity information  110 . In alternative embodiments, other information may be received and monitored to determine if the GNSS location information has been corrupted. 
         [0033]    While the preferred embodiment of the invention has been illustrated and described, as noted above, many changes can be made without departing from the spirit and scope of the invention. Accordingly, the scope of the invention is not limited by the disclosure of the preferred embodiment. Instead, the invention should be determined entirely by reference to the claims that follow.