Abstract:
Potentially hazardous conditions encountered by an aircraft are signaled to the flight crew of the aircraft by a system of visible and audible annunciators. A first visible annunciator indicates cautions and warnings and a second visible annunciator indicates system classifications for detected conditions. The second visible annunciator indicates, for example, a threatened stall, a low fuel condition or excessive airspeed. An audible voice message played over the aircraft communication system is tailored to the potential hazard condition. A user operable button provides condition-specific cancellation authority.

Description:
RELARED APPLICATION  
       [0001]    This application claims priority to U.S. Provisional Application serial number 60/306,632 (entitled FLIGHT MANAGEMENT ANNUNCIATOR PANEL AND SYSTEM, filed Jul. 19, 2001) which is herein incorporated by reference. 
     
    
     
       TECHNICAL FIELD  
         [0002]    This invention relates generally to flight management systems and particularly, but not by way of limitation, to a flight management system and methods for general aviation aircraft.  
         BACKGROUND  
         [0003]    Flight safety is an important consideration for pilots of general aviation aircraft. In addition to safely operating the airplane and maintaining a desired course, pilots are also tasked with monitoring the data presented by the various cockpit instruments. This data can show, for example, that the fuel level remaining in a particular tank is approaching a dangerous state. During solo flight operations in congested airspace under instrument flight conditions, the general aviation pilot is sometimes overburdened with monitoring the various instruments in the typical aircraft.  
           [0004]    What is needed is a system and method that can be installed in new aircraft or retrofitted to existing aircraft that provides the pilot with early notification of a potential problem.  
         SUMMARY  
         [0005]    Signals from a plurality of sensors distributed throughout the aircraft are analyzed by an on-board processor. The processor drives a visual warning display and produces an audible message. The visual warning display includes a first part and a second part. The first part visually annunciates the condition and includes a user operable switch. The second part provides specific system information as to the nature of the condition and is installed in view of the pilot. The pilot can cancel the audible message using an annunciator switch on the first part. For certain warning conditions, activation of the annunciator switch will extinguish the visual warning display message appearing on the second part.  
           [0006]    Set points for alerts triggered by the various sensors are determined by calibration routines executed at the time of configuring the system and determined by aircraft performance data. In one embodiment, a portable computer can be coupled to the present system by an interface connection. The portable computer, which may be a laptop computer, can be used to access stored data, upload programming to the on-board processor, or manage the calibration or operation of the present system. The set points for selected conditions can be adjusted dynamically based on measured parameters.  
           [0007]    An existing aircraft can be retrofitted with the present subject matter without disturbing the type certification issued by the Federal Aviation Administration or other governing body. In addition, the present subject matter can be incorporated in the production of new aircraft and certificated as a package.  
           [0008]    The present subject matter provides the pilot with status and warning information from several aircraft systems and the cabin environment. Status and warning annunciations provided by the present subject matter are advisory only and shall be treated as secondary to the primary aircraft indication systems. The present system is designed to mimic the indications given by those systems. The present subject matter can adapt to a variety of aircraft having different bus voltages and transducer output levels.  
           [0009]    Conditions and parameters monitored by the present system include those that may be potentially hazardous. In one embodiment, a first visible annunciator indicates cautions and warnings and a second visible annunciator indicates system classifications for detected conditions. For example, system notification may include fuel system, cabin integrity, landing gear system, electrical system, navigation system, flight-related systems such as airspeed, stall speed or other systems. The second visible annunciator indicates, for instance, a threatened stall, a low fuel condition or excessive airspeed. An audible voice message, which may be synthesized or recorded voice, is played over the aircraft communication system and is tailored to the potential hazard condition. A user operable button provides condition-specific cancellation authority.  
           [0010]    This summary is intended to provide a brief overview of some of the embodiments of the present system, and is not intended in an exclusive or exhaustive sense, and the scope of the invention is to be determined by the attached claims and their equivalents. 
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0011]    In the drawings, like numerals describe substantially similar components throughout the several views. Like numerals having different letter suffixes represent different instances of substantially similar components.  
         [0012]    [0012]FIG. 1A is a drawing illustrating selected portions of a flight management system.  
         [0013]    [0013]FIG. 1B is a block diagram illustrating an acknowledge switch module.  
         [0014]    [0014]FIG. 2 is a schematic illustrating generally one embodiment of a processor coupled to selected devices and systems.  
         [0015]    [0015]FIG. 3 illustrates a processor coupled to various sensors of a flight management system.  
         [0016]    [0016]FIG. 4A illustrates a fuel gauge coupled to a processor according to one embodiment of the present subject matter.  
         [0017]    [0017]FIG. 4B illustrates an oil temperature gauge coupled to a processor according to one embodiment of the present subject matter.  
         [0018]    [0018]FIG. 5A illustrates a schematic for monitoring current in a fuel pump circuit.  
         [0019]    [0019]FIG. 5B illustrates a schematic for monitoring current in a pitot heater circuit.  
         [0020]    [0020]FIG. 6 illustrates a model of an oil temperature gauge circuit.  
         [0021]    [0021]FIG. 7 illustrates a graph of oil temperature sensor voltage as a function of indicated temperature for different aircraft bus voltages.  
         [0022]    [0022]FIGS. 8A, 8B,  8 C and  8 D illustrate connections to various aircraft instruments according to one embodiment of the present subject matter.  
         [0023]    [0023]FIG. 9 illustrates a flow chart for receiving information from the acknowledge switch.  
         [0024]    [0024]FIG. 10 illustrates a flow chart for receiving pilot selections using the display module switch.  
         [0025]    [0025]FIG. 11 illustrates a flow chart for a method based on actuation of the acknowledge switch in a caution mode.  
         [0026]    [0026]FIG. 12 illustrates a flow chart for a method based on actuation of the acknowledge switch in a warning mode.  
         [0027]    [0027]FIG. 13 illustrates a flow chart for a method based on monitoring the pitot heat sensor.  
         [0028]    [0028]FIG. 14 illustrates a flow chart for a method based on monitoring the fuel pump sensor.  
         [0029]    [0029]FIG. 15 illustrates a flow chart for a method based on the stall warning vane switch.  
         [0030]    [0030]FIG. 16 illustrates a flow chart for a method based on detecting a threatened accelerated stall.  
         [0031]    [0031]FIGS. 17A and 17B illustrate flow charts for a method based on alerting for a waypoint.  
         [0032]    [0032]FIGS. 18A and 18B illustrate flow charts for methods based on fuel quantity indications.  
         [0033]    [0033]FIG. 19 illustrates a flow chart for a method based on remaining fuel in a tank.  
         [0034]    [0034]FIG. 20 illustrates a flow chart for a method based on a measured oil temperature.  
         [0035]    [0035]FIG. 21 illustrates a flow chart for a method based on carbon monoxide levels detected in the aircraft cabin.  
         [0036]    [0036]FIG. 22 illustrates a flow chart for a method based on detected oil pressure levels.  
         [0037]    [0037]FIG. 23 illustrates a flow chart for a method based on detected oil pressure levels and vacuum levels.  
         [0038]    [0038]FIG. 24 illustrates a flow chart for a method based on landing gear position sensor switches.  
         [0039]    [0039]FIG. 25 illustrates a flow chart for a method based on landing gear position sensor switches.  
         [0040]    [0040]FIG. 26 illustrates a flow chart for a method based on aircraft bus voltages.  
     
    
     DETAILED DESCRIPTION  
       [0041]    In the following detailed description, reference is made to the accompanying drawings which form a part hereof, and in which is shown by way of illustration specific embodiments in which the invention may be practiced. These embodiments are described in sufficient detail to enable those skilled in the art to practice the invention, and it is to be understood that the embodiments may be combined, or that other embodiments may be utilized and that structural, logical and electrical changes may be made without departing from the spirit and scope of the present invention. The following detailed description is, therefore, not to be taken in a limiting sense, and the scope of the present invention is defined by the appended claims and their equivalents. In the drawings, like numerals describe substantially similar components throughout the several views. Like numerals having different letter suffixes represent different instances of substantially similar components.  
         [0042]    [0042]FIG. 1A illustrates a flight management system according to one embodiment of the present subject matter. In this example, the system is coupled to, and installed in, general aviation aircraft  12 . Aircraft  12  is powered by a single engine. Display module  300 A and acknowledge switch module  400 A are installed within the cockpit of aircraft  12 . Sensor module  100  is installed in the cabin environment of aircraft  12  and main processor unit  200 A is installed on the airframe of aircraft  12 . Main processor unit (MPU)  200 A is coupled to display module  300 A, acknowledge switch module  400 A and sensor module  100 A by digital data lines. In addition, selected electrical sensors and switches of aircraft  12  are coupled to MPU  200 A by signal lines  50 . Sensor module  100 A includes one or more pressure transducers coupled to selected aircraft systems by air pressure sense, or pneumatic, lines. Sensor module  100 A also includes an accelerometer adapted to provide an electrical signal based on acceleration of aircraft  12 . In one embodiment, MPU  200 A includes a main processor, an input board and a voice board.  
         [0043]    Display  300 A includes visual display  305 , light sensor  310  and mute switch  315 . Visual display  305  includes an array of light emitting diodes (LEDs) arranged in an upper line  305 A and a lower line  305 B adapted to illuminate one or more warning or caution messages. Displayable messages include “GPS,” “WPT,” “FUEL PUMP,” “GEAR,” “CO MOX,” “STALL,” “VAC” and “VOLTS” in upper line  305 A and “NAV,” “MSG,” “PITOT HT,” “CYL TEMP,” “OIL PRES,” “L FUEL” and “R FUEL” in lower line  305 B. Messages appearing in display  305  are selected for illumination based on signals received from MPU  200 A and correspond to sensed conditions, cautions, alerts or warnings. Zero messages, one message or multiple messages may be displayed simultaneously by display  300 A.  
         [0044]    The light intensity of display module  300 A and acknowledge switch module  400 A is adjustable. Light sensor  310 , disposed on the face of display module  300 A, senses the ambient light and generates a signal received by MPU  200 A. Main processor unit  200 A executes a program to adjust the light intensity of both display module  300 A and acknowledge switch module  400 A based on the signal from light sensor  310 .  
         [0045]    Mute switch  315 , disposed on the face of display module  300 A, is a user accessible, momentary contact, push button switch. Switch  315  controls muting of the voice annunciation, controls playback of previously muted and stored alert messages, controls playback of active alert messages, cautions or conditions and controls execution of rebooting and self-testing. After muting, when switch 315 is again pressed, all previously muted and stored alert messages that remain active are presented via audio  450  in the order of detection.  
         [0046]    Acknowledge switch module  400 A includes a two-message display and a user operable push button switch. The two-message display allows selective illumination of a first message such as “WARNING” and a second message such as “CAUTION.” Zero messages, one message or two messages may be displayed simultaneously by acknowledge switch module  400 A. The legend “PUSH” appears in the center of acknowledge switch module  400 A.  
         [0047]    In one embodiment, acknowledge switch module  400 A is of a shape and size conducive to mounting in, or near, the pilot&#39;s instrument scan of aircraft  12 . Display module  300 A is mounted in a location readily viewable by the pilot. For example, in one embodiment, acknowledge switch  400 A is mounted immediately in front of the pilot near the artificial horizon and display module  300 A is mounted in the instrument panel above the radio stack.  
         [0048]    [0048]FIG. 1B illustrates a model of acknowledge switch module  400 C according to one embodiment of the present subject matter. Acknowledge switch module  400 C includes warning light  410 , caution light  415  and switch  420  mounted in housing  420 . Electrical connections to warning light  410 , caution light  415  and switch  420  are provided by a connector.  
         [0049]    When caution light  415  is illuminated, the pilot is instructed to take one of two actions. The pilot acknowledges the cautionary alert by hitting switch  400  once. By depressing switch  420  a single time, caution light  415  is extinguished and a five minute caution reminder timer is started. If the alert condition corresponding to the caution alert remains active at the completion of the timer duration, caution light  415  will again be illuminated as a reminder alert. If the pilot double clicks switch  420 , caution light  415  will be extinguished and will remain extinguished unless the alert condition is repeated or another input illuminates caution light  415 . Double clicking switch  420  suspends the caution reminder timer.  
         [0050]    Warning light  410  an not be manually extinguished. Warning light  410  remains illuminated as long as the warning condition exists.  
         [0051]    Acknowledge switch  400 C provides volume control for audio  450 . If the pilot depresses switch  420  for a period of approximately 3 seconds, audio  450  will annunciate “volume.” The pilot is instructed to depress and release the switch, which will cycle the volume settings from a level of  0  (volume off) to  7  (maximum volume). Audio  450  announces the volume level as it is being changed.  
         [0052]    Acknowledge switch  400 C provides access to a training mode. If the pilot depresses switch  420  for a period of approximately 5 seconds, audio  450  will annunciate “training mode.” In the training mode, toggling switch  420  will toggle between training mode on and training mode off. Audio  450  announces the training mode state as it is changed. When training mode is on, the voice annunciations for stall warning, accelerated stall, gear up, gear down, and check gear down are muted and the gear and stall lights function normally. The system will revert to normal operation at the next power up.  
         [0053]    Acknowledge switch  400 C provides access the software version for various elements. If the pilot depresses switch  420  for a period of approximately 7 seconds, audio  450  will annunciate “software version number” followed by a verbal announcement of the version number for MPU  200 B, display module  300 B and sensor module  100 B.  
         [0054]    [0054]FIG. 2 schematically illustrates a block diagram of one embodiment of the present subject matter. In the figure, MPU  200 B is coupled to a first group of input signals, collectively referred to herein as  505 A, each of which provides a digital signal. Main processor unit  200 B is also coupled to a second group of input signals via sensor module  100 B. In addition, MPU  200 B is coupled to display  300 B, acknowledge switch  400 B, audio  450  and computer  650 .  
         [0055]    Main processor unit  200 B includes brown out detector  205  and watchdog  210 . Brown out detector  205  operates by monitoring the 5 v supply to MPU  200 B. In one embodiment, if the supply voltage drops to 4.5 volts, then brown out detector  205  will power down the system. Brown out detector  205  continues to monitor the supply voltage to MPU  200 B and if the voltage rises to a predetermined level, then a re-boot operation is executed. Software executing on MPU  200 B generates a pulse every 100 milliseconds (mS). If a pulse is missing, then after a delay time period, watchdog  210  triggers a re-boot operation. The delay time period is 4 seconds in one embodiment.  
         [0056]    Those inputs to MPU  200 B directed to navigation functions are collectively referred to as inputs  502 A. Inputs  502 A includes waypoint  510 , navigation  515 A, global positioning system (GPS)  520 A and message  525 A. Each of inputs  502 A receives a digital signal and depending on the signals received, illuminates a segment of display module  300 B.  
         [0057]    For example, if navigation information is received from a GPS receiver in aircraft  12 , then input GPS  520 A will be at a digital low level and the message “GPS” will be illuminated in display module  300 B in the color green. An audible message is not generated with this indication.  
         [0058]    When the navigation information is received from a VORWLOC (very high frequency, VHF, omnidirectional range/localizer) or other non-GPS based navigation instrument, then input NAV  515 A will be at a digital low level and the message “NAV” will be illuminated in display module  300 B in the color green. As with the GPS message, an audible message is not generated with this indication.  
         [0059]    When the navigation information source issues a waypoint alert, then input WPT  510 A will be at a digital low level and the message “WPT” will be illuminated in display module  300 B in the color green. In addition, MPU  200 B causes an audible alert message to be presented. The alert message includes the spoken word “waypoint” played over a headset to be worn by the pilot or played over a cabin speaker. The audible alert message, or voice annunciation, is played via audio  450 . Audio  450 , in one embodiment, includes an aircraft audio panel.  
         [0060]    When the navigation information source issues a message alert, then input MSG  525 A will be at a digital low level and the message “MSG” will be illuminated in display module  300 B in the color amber. As with the GPS message and the NAV message, an audible message is not generated with this indication.  
         [0061]    Input CYL TEMP  530 A is coupled to a spark-plug ring J-type thermocouple and provides an electrical signal corresponding to a cylinder head temperature. When the monitored cylinder head temperature is at the maximum temperature operating range, the message “CYL TEMP” will be displayed in display module  300 B in the color red. In addition, acknowledge switch  400 B illuminates warning light  410  and a voice annunciation alert stating “check cylinder temperature” is presented audibly via audio  450 . Audible presentation entails playing the annunciation message via the pilot&#39;s headset or a cabin speaker. The set point for the cylinder temperature warning is stored in memory accessible to MPU  200 B and is determined by aircraft  12 . In one embodiment, a cylinder head temperature high set point is 430° F.  
         [0062]    Inputs L FUEL  535 A and R FUEL  535 B are coupled to fluid level transducers installed in the left and right fuel tanks. Main processor unit  200 B receives voltage signals from each transducer. An empty fuel tank is represented by zero volts and the voltage output from a full tank is determined by filling the tank with fuel and storing the output voltage in a memory accessible to MPU  200 B. Software executing on MPU  200 B calculates and establishes alarm set points at 25% and 10% of the range. When MPU  200 B determines that the remaining fuel in any one tank of aircraft  12  corresponds to the 25% set point, caution light  415  is illuminated and display module  300 B displays a corresponding message, such as “L FUEL” or “R FUEL,” and MPU  400 B generates and presents an audible message such as “right tank at twenty-five percent” or “left tank at twenty-five percent” via audio  450 . When MPU  200 B determines that the remaining fuel in all tank corresponds to a value at or below the 10% set point, MPU  400 B generates and presents an audible message such as “right tank at ten percent” or “left tank at ten percent” via audio  450 . When MPU  200 B determines that the remaining fuel in all tanks is below the 10% set point, caution light  415  is extinguished and warning light  410  is illuminated.  
         [0063]    In one embodiment, the maximum transducer output voltage, corresponding to full tanks) is used to calculate the set points at lower fuel quantities. False fuel level alerts are reduced by storing  1024  transducer voltage samples and taking the average over 100 seconds.  
         [0064]    Input PITOT HT CURRENT  550 A is coupled to a solid state current monitor. When the aircraft pitot heat switch is in the off position, no current is drawn and the visual and audible annunciators are off. When the pitot heat switch is in the on position, as determined by a line coupled to the pitot heat panel switch and PITOT HT  545 A, and current flow to the pitot heater is detected, display module  300 B illuminates a message “PITOT HT” in the color green. When the pitot heat switch is in the on position and the current is less than the predetermined set point, then the “PITOT HT” message of display module  300 B is extinguished and caution light  415  is illuminated. In addition, a voice annunciation alert “pitot heat failure” is presented via audio  450 . In one embodiment, the minimum pitot heat current is 3.00 amperes and the maximum pitot heat current is 1.00 amperes.  
         [0065]    Input FUEL PUMP CURRENT  560 A is coupled to a solid state current monitor. When the fuel pump switch is in the off position, no current is drawn and the visual and audible annunciators are off. When the fuel pump switch is in the on position, as determined by a line coupled to the fuel pump panel switch and FUEL PUMP  555 A, and current flow to the fuel pump is detected display module  300 B illuminates a message “FUEL PUMP” in the color green. When the fuel pump switch is in the on position and the current is less than the predetermined set point, then the “FUEL PUMP” message of display module  300 B is extinguished and caution light  415  is illuminated. In addition, a voice annunciation alert “fuel pump failure” is presented via audio  450 .  
         [0066]    Input OIL TEMP  565 A is coupled to a thermal sensor exposed to an engine oil galley. The engine oil temperature thermal sensor may include a resistance temperature detector (RTD) probe configured as a plug or bayonet. In one embodiment, the oil temperature thermal sensor is also used to provide an electrical signal for a panel-mounted temperature gauge. When the measured oil temperature is above the maximum temperature set point, a voice annunciation alert “check oil temperature” is presented via audio  450 . In addition, warning light  410  is illuminated and the voice annunciation “check oil temperature” repeats every 5 seconds unless switch  420  is activated. Activation of switch  420  will delay or cancel the audio alert from audio  450 . The maximum temperature set point is selected based on data provided by the aircraft manufacturer or as determined by the aircraft operators manual and stored in memory accessible to MPU  200 B. In one embodiment, the maximum temperature set point for the oil is 225° F.  
         [0067]    Input OIL PRES  570 A is coupled to pressure transducer exposed to engine oil pressure. When the measured oil pressure is below a minimum oil pressure level, warning light  410  is illuminated and display module  300 B illuminates “OIL PRES” in a red color. In addition, a voice annunciation alert “check oil pressure” is presented via audio  450 . The minimum oil pressure set point is selected based on data provided by the aircraft manufacturer or as determined by the aircraft operators manual and stored in memory accessible to MPU  200 B. In one embodiment, the low oil pressure set point is approximately 30.0 psi.  
         [0068]    Input VOLTS  575 A is coupled to the aircraft supply bus. When the measured voltage exceeds the high set point, warning light  410  is illuminated and display module  300 B illuminates a “VOLTS” message in an amber color. In addition, a voice annunciator alert “bus voltage high” is presented via audio  450 . When the measured voltage is below the low set point, caution light  415  is illuminated and display module  300 B illuminates the “VOLTS” message in an amber color. In addition, a voice annunciator alert “bus voltage low” is presented via audio  450 . The high set point and the low set point are selected based on data provided by the aircraft manufacturer or as determined by the aircraft operators manual and stored in memory accessible to MPU  200 B. In one embodiment, the high set point and low set point for the bus voltage is 14.50 and 12.00 volts DC (VDC), respectively.  
         [0069]    Input STALL  580 A is coupled to a stall vane switch on aircraft  12 . In unaccelerated flight, when the vane switch indicates a stall condition, display module  300 B illuminates the message “STALL” in a red color and warning light  410  is illuminated. In addition, a voice annunciator alert “stall, stall, stall” is presented via audio  450 .  
         [0070]    In accelerated flight, accelerometer  130 A, coupled to MPU  200 B via sensor module  100 B senses the g-loading. When the g-loading of the aircraft, as measured by accelerometer  130 A, and airspeed of the aircraft, as measured by the airspeed transducer of sensor module  100 B, indicates that aircraft  12  is within 5 to 10 knots of an accelerated stall, display module  300 B illuminates a “STALL” message in red color and caution light  415  is illuminated. In addition, a voice annunciation alert “check airspeed” is presented via audio  450 . In one embodiment, to reduce false alarms resulting from turbulence, a delay period of time is introduced before triggering the visual and audible annunciation. The set point for the accelerated stall is selected based on data provided by the aircraft manufacturer or as determined by the aircraft operators manual and is stored in memory accessible to MPU  200 B. In one embodiment, the set point corresponds to the flaps up, maximum weight data. In one embodiment, the accelerated stall speed set point is 60 knots. Programming executing on MPU  200 B of the present system determines the accelerated stall speed of the aircraft based on extrapolation, or interpolation, of accelerated stall speed data. Visible and audible annunciation of accelerated stall warning is disabled for airspeeds below 90% of the unaccelerated stall speed (that is, flaps up, maximum weight, straight and level flight).  
         [0071]    Table 1 illustrates various accelerated stall speeds as a function of g-loading. An accelerated stall can occur at any bank angle, however, for comparison sake, the table below also shows corresponding bank angles for the listed accelerations. For example, an aircraft with a 40 knot stall speed in straight and level flight will stall at 43 knots when accelerated to 1.15 g, typically encountered in a coordinated turn at a 30 degree bank angle. In a turn at 45 degrees, the same aircraft is experiencing an acceleration of 1.44 g and will stall at 48 knots. In a turn at 60 degrees, the same aircraft is experiencing an acceleration of 2.0 g and will stall at 57 knots. The table also presents stall speeds for other aircraft having unaccelerated stall speeds ranging to 85 knots in 5 knot increments.  
                                                         TABLE 1                                   1.0   1.15 g   1.44 g   2.0 g           (0 degrees)   (30 degrees)   (45 degrees)   (60 degrees)                                        40   43   48   57           45   48.5   53   63           50   54   59   70           55   59   65.5   77           60   64.5   72.5   85           65   70   77   92           70   75.5   83   99           75   80.5   89   107           80   86   95   113           85   92   101   120                      
 
         [0072]    Input GEAR UP  585 A is coupled to the gear up position switch of aircraft  12  and provides an electrical signal when the aircraft landing gear is in the up, or raised, position. Input GEAR DN is coupled to the gear down position switch of aircraft  12  and provides an electrical signal when the aircraft landing gear is in the down position. The GEAR message appearing on display module  300 B is in a red color.  
         [0073]    A voice annunciation alert stating “check gear” is presented via audio  450 , warning light  410  is illuminated, and a “GEAR” message appears on display module  300 B if a landing gear in-transit signal is received for a duration more than 20 seconds.  
         [0074]    Several methods can be employed to determine when the landing gear is in-transit and one method is selected at the time of installation or calibration. One exemplary method entails monitoring for a time when both gear up and gear down lights are extinguished. One exemplary method entails monitoring illumination of a gear unsafe light. One exemplary method entails monitoring landing gear switches showing that the landing gear is not in a locked position. Other methods are also contemplated.  
         [0075]    In addition, a voice annunciation alert stating “check gear” is presented via audio  450 , warning light  410  is illuminated, and a “GEAR” message appears on display module  300 B if aircraft  12  indicates a gear-down and gear-up condition simultaneously for a duration of more than 6 seconds. Also, a voice annunciation alert stating “check gear” is presented via audio  450 , warning light  410  is illuminated and a “GEAR” message appears on display module  300 B if both the gear-down and gear-up indicators remain off simultaneously for a duration of more than 20 seconds.  
         [0076]    When sensor module  100 B signals that the airspeed is below a predetermined set point and the gear position indicator signals that the landing gear is in the up position, warning light  410  is illuminated, display module  300 B illuminates the “GEAR” message and a voice annunciation alert “check gear down” will be presented via audio  450 . In one embodiment, the gear down set point is 90 knots and speeds below this level with the landing gear in the up position will trigger a warning.  
         [0077]    A voice annunciation alert “gear up” will be presented via audio  450  when the landing gear system completes a gear up cycle. A voice annunciation alert “gear down” will be presented via audio  450  when the landing gear system completes a gear down cycle.  
         [0078]    If the pilot selects gear down while at an airspeed, as detected by sensor module  100 B, above a maximum gear extension speed set point, warning light  410  is illuminated, a “GEAR” message is illuminated on display module  300 B and a voice annunciation alert “gear overspeed” will be presented via audio  450 . The maximum gear extension speed set point is stored in a memory accessible to MPU  200 B. In one embodiment, the maximum gear extension speed set point is 130 knots.  
         [0079]    Input BAGGAGE DOOR  595 A is coupled to a door switch adapted to indicate an unsafe baggage or utility door position. When the door switch indicates an unsafe position, caution light  415  is illuminated and remains on until switch  420  is manually operated. In addition, a voice annunciation alert “check baggage door” is presented via audio  450 . The voice annunciation is repeated on five minute intervals. Pressing switch  420  once extinguishes caution light  415  and starts a timer having a duration of approximately five minutes. At the end of the timer period, caution light  415  is again illuminated. Pressing switch  420  twice extinguishes caution light  415  and terminates the voice annunciation without starting a timer. The baggage door alert will be reset and a subsequent detection of a door switch signal will again trigger a caution light and voice annunciation.  
         [0080]    Input CABIN DOOR  600 A is coupled to a door switch adapted to indicate an unsafe cabin door position. When the door switch indicates an unsafe position, caution light  415  illuminates and remains on until switch  420  is manually operated. In addition, a voice annunciation alert “check cabin door” is presented via audio  450 . The voice annunciation is repeated on five minute intervals. Pressing switch  420  once extinguishes caution light  415  and starts a timer having a duration of approximately five minutes. At the end of the timer period, caution light  415  is again illuminated. Pressing switch  420  twice extinguishes caution light  415  and terminates the voice annunciation without starting a timer. The cabin door alert will be reset and a subsequent detection of a door switch signal will again trigger a caution light and voice annunciation.  
         [0081]    Input ENGINE ANALYZER  605 A is coupled to an electronic engine analyzer or monitor having a discrete alarm output signal. When an alarm output signal is received, warning light  410  illuminates and remains on until the engine analyzer or monitor cancels the signal. In addition, a voice annunciation alert “check engine analyzer” is presented via audio  450 . The voice annunciation is repeated on five minute intervals. Pressing switch  420  twice terminates the voice annunciation without starting a timer.  
         [0082]    Programming executing on MPU  200 B monitors for airspeeds approaching V NE  (velocity, never exceed), as determined by data specified by the aircraft manufacturer. When aircraft  12  is at a speed within 5% of V NE , warning light  410  is illuminated and, on 5 second intervals, a voice annunciation of ‘check airspeed” is presented via audio  450 . Actuation of switch  420  will delay or cancel the voice annunciation.  
         [0083]    Programming executing on MPU  200 B monitors for a gear overspeed condition. If the landing gear is not in an up position and airspeed is above a maximum landing gear airspeed set point, as determined by data provided by the aircraft manufacturer, then warning light  410  is illuminated, display module  300 B illuminates a gear message, and a voice annunciation of “gear overspeed” is presented via audio  450 .  
         [0084]    Computer  650 , which may include a desktop, laptop, handheld or other computer, can be coupled to MPU  200 B via connector cable  230 . In one embodiment, cable  230  includes an RS 232  serial cable and is selectable for communicating using port com 1  or com 2 .  
         [0085]    Programming executing on computer  650  communicates with MPU  200 B and provides access to stored data, calibration functions and data writing functions. For example, computer  650  can read the serial number, version number or other data concerning MPU  200 B, display module  300 B and sensor module  100 B. Computer  650  can adjust the volume level of alerts delivered via audio  450 , greetings volume (including recitation of the volume level for voice alerts, training mode and software version), as well as read or write the following parameters and set points: cylinder temperature (high set point), bus voltage (low and high set point), oil pressure (low set point), vacuum (fail, low and high set points), carbon monoxide alert and warning levels, gear down speed set point, gear over-speed set point, accelerated stall speed, minimum pitot current set point, minimum fuel pump current set point and maximum airspeed (velocity never exceed, V NE  ) set point. In addition, computer  650  includes programming to specify or select and input table corresponding to gear lights, cabin door, baggage door, engine analyzer and other functions.  
         [0086]    Furthermore, computer  650  includes programming to initiate, set or read calibrations for the fuel tank (full), airspeed, carbon monoxide (low and high levels), oil temperature range calibration (including bus voltage variance) and accelerometer sensor. Calibration of the present subject matter may be performed in a shop or in the aircraft.  
         [0087]    Calibration of the accelerometer entails noting the output signal from the accelerometer while positioned in two different orientations. The accelerometer is positioned as it will be mounted in the aircraft and a first output signal is stored. A second output signal is stored when the accelerometer is positioned inverted. The combination of the first output signal and second output signal allows MPU  200 B to calibrate the accelerometer.  
         [0088]    Sensor module  100 B includes pressure transducers and is coupled to MPU  200 B by bus  240 . For example, sensor module  100 B interfaces with the aircraft pitot pressure, static pressure, vacuum and pressure altitude (for pressurized aircraft), the accelerometer (g-sensor)  130 A and MPU  200 B. Sensor module  100 B includes pitot input  110 A and static input  115 A, coupled to the aircraft pitot line and static pressure line. One or more pressure transducers of sensor module  100 B provides an electrical signal based on the pressures sensed at pitot input  100 A and static input  115 A. In one embodiment, the pressure transducer includes a silicone diaphragm solid-state device that is temperature compensated and having a range of +/− 1 . 5  psi and a resolution of +/−2 knots. In a failure mode, the pressure transducer will generate a false “check airspeed” alert.  
         [0089]    Inputs to sensor module  100 B includes vacuum input  120 A and ambient input  125 A, coupled to the aircraft vacuum and ambient pressure system. For pressurized aircraft, ambient input  125 A is coupled to a region external to the pressure vessel. Vacuum input  120 A is coupled to the vacuum system at an open port or tee on a vacuum driven instrument. One or more pressure transducers of sensor module  100 B provides an electrical signal based on the pressures sensed at vacuum input  120 A and ambient input  125 A. In one embodiment, the pressure transducer includes a silicone diaphragm solid-state device that is temperature compensated and having a range of +/−14.5 psi and a resolution of +/−2.5%. In a failure mode, the pressure transducer will produce a false vacuum alert. In one embodiment, the vacuum fail, vacuum low and vacuum high set points are set to 1.00, 3.00 and 6.00 in Hg, respectively.  
         [0090]    Accelerometer  130 A includes a microchip that generates an analog voltage when acceleration is detected along one axis. Accelerometer  130 A has a range of +/−5 g&#39;s and has a resolution of 2 mg&#39;s. In a failure mode, accelerometer  130 A will produce a false “check airspeed” alert.  
         [0091]    Carbon monoxide sensor  135 A is coupled to sensor module  100 B. Carbon monoxide sensor is installed in a position to be exposed to cabin air and includes a solid state device that uses chromium titanium oxide as a detecting material. Sensor  135 A has a range of 5−100 ppm and an accuracy of +/−3 ppm at 30 ppm and +/−5 ppm at 70 ppm. In a failure mode, sensor  135 A will provide a signal to indicate an alert. MPU  200 B executes programming to detect a failure of sensor  135 A. If MPU  200 B detects a failure of carbon monoxide sensor  135 A, then a voice annunciation will present “carbon monoxide sensor failure” via audio  450  and the carbon monoxide alert is disabled. MPU  200 B illuminates caution light  415  for carbon monoxide levels between 35 and 50 ppm and illuminates warning light  410  for carbon monoxide levels above 50 ppm. The time difference between illumination of caution light  415  and warning light  410  provides an indication of the rate of increase of carbon monoxide in the cabin. In one embodiment, a digital display in the cockpit provides a numerical value for measured carbon monoxide levels.  
         [0092]    Calibration of carbon monoxide sensor  135 A includes storing (in memory accessible to MPU  200 B) the pulse counts generated by sensor  135 A when exposed to an atmosphere known to have a carbon monoxide concentration of 0 ppm and when exposed to an atmosphere of 50 ppm. A linear relationship between the values of 0 ppm and 50 ppm allows extrapolation of counts for atmospheres having levels greater than 50 ppm.  
         [0093]    In one embodiment, sensor  135 A includes a solid state element sensitive to variations in humidity, temperature and altitude. At the time of system start up, and after sensor  135 A has reached a stable operating temperature, a base line measure of counts is stored in memory accessible to MPU  200 B. Programming executing on MPU  200 B calculates an offset from the baseline to allow accurate measurements of elevated carbon monoxide levels.  
         [0094]    In addition, a raw pulse count is determined at initial start up of the present system. The raw pulse count may be in the range of 0 to 1024 pulses per 100 mS from an analog-to-digital (A-to-D) counter. If the raw count range is less than 375 or greater than 725, then a voice annunciation message of “carbon monoxide failure” is presented via audio  450 .  
         [0095]    A heater provides a stable, warm temperature for sensor  135 A. Heater failure is detected by programming executing on MPU  200 B which changes the electrical power supplied to the heater by means of a switchable resistor divider network. Heater functionality is verified by a change in sensor  135 A counts when the heater power is changed. In one embodiment, a properly functioning sensor  135  will generate a change of 80+/−50 counts when the heater power is changed. In one embodiment, sensor  135  will exhibit an increase in the count when the heater power is reduced.  
         [0096]    In a high temperature, high humidity environment, sensor  135 A clears itself as a function of power on time. Sensor  135 A operates at a nominal temperature of 400° C. when powered. A carbon filter element of sensor  135 A removes impurities at the elevated temperature. When sensor  135 A is unpowered, the sensor absorbs moisture in a high humidity, warm temperature environment. Impurities remaining in the carbon filter interfere with accurate readings from sensor  135 A. After continuously powering sensor  135 A on for approximately 5 to 6 hours, impurities on the carbon filter element are baked out and sensor  135 A again will provide accurate data. If the count rate drops below the start-up baseline count, then programming executing on MPU  200 B will replace the start-up baseline with the new low level.  
         [0097]    If the count is greater than 500, then sensor  135 A calculates a new baseline level every two minutes and verifies heater functionality. If the count is between 450 and 500, then sensor  135 A calculates a new baseline every 20 minutes.  
         [0098]    If the airspeed is greater than 40 knots and the system performs a reboot, then sensor module  100 B uses the last baseline value and disables the start-up baseline calculation.  
         [0099]    Software  
         [0100]    Processors in MPU  200 B, display module  300 B and sensor module  100 B execute instructions to control the operation of the present subject matter.  
         [0101]    An engine start routine is executed to determine whether the aircraft engine has properly started or is malfunctioning. The routine receives data from the oil pressure transducer, via input OIL PRES  570 A, and the aircraft vacuum pressure system, via vacuum input  120 A. The first pressure to reach 50% of its low or minimum set point will arm both annunciators for proper failure annunciation at engine start.  
         [0102]    In addition, an airspeed indication, via inputs  110 A and  115 A, above the stall speed set point indicates the aircraft is in flight and activates gear annunciation and airspeed annunciations. In one embodiment, if either oil pressure or vacuum is above 50% of its minimum set point, then both are armed. Therefore, if oil pressure rises and vacuum does not, then the oil pressure will in turn, arm the vacuum and the vacuum warning will occur.  
         [0103]    Annunciation Manipulation and Configuration  
         [0104]    The message lights of display module  300 B remain illuminated for as long as the alert is active.  
         [0105]    Various conditions will give rise to voice annunciations not accompanied by illumination of an acknowledge switch message or display module message. In these cases, the voice alert is presented once upon alert initiation and if the alert condition were to be corrected and occur again, then a second voice annunciation would be provided. In one embodiment, an alert condition is signaled using acknowledge switch  400 B. In one embodiment, an alert condition is signaled using both acknowledge switch  400 B and display module  300 B.  
         [0106]    Notwithstanding the foregoing, for oil temperature alerts, the voice annunciation is repeated every 5 to 10 seconds (approximately) and warning light  410  remains illuminated. A single actuation of switch  420  will extend the voice annunciation repetition rate to once every 5 minutes. A double actuation of switch  420  will cancel future voice alerts as to oil temperature.  
         [0107]    In addition, for airspeed alerts, the voice annunciation is repeated every 5 to 10 seconds (approximately) and warning light  410  remains illuminated. A single actuation of switch  420  will extend the voice annunciation repetition rate to once every 5 minutes. A double actuation of switch  420  will cancel future voice alerts as to airspeed at the V NE .  
         [0108]    Also, for engine analyzer alerts, audio  450  presents the alarm signal as provided by the analyzer. For example, when a parameter monitored by the analyzer is exceeded, the present subject matter will annunciate the alarm. The voice annunciation is repeated every 5 minutes (approximately) and warning light  410  remains illuminated. A double actuation of switch  420  will cancel future voice alerts as to the engine analyzer.  
         [0109]    For cabin door alerts, the voice annunciation is repeated every 5 minutes (approximately) and caution light  415  remains illuminated. A single actuation of switch  420  will cause caution light  415  to extinguish for the duration of the 5 minutes and illuminate thereafter. A double actuation of switch  420  will cancel future alerts as to the cabin door ajar unless the cabin door is cycled. Cabin door position is determined by an airframe mounted microswitch.  
         [0110]    For baggage or utility door alerts, the voice annunciation is repeated every 5 minutes (approximately) and caution light  415  remains illuminated. A single actuation of switch  420  will extend the voice annunciation repetition rate to once every 5 minutes and cause caution light  415  to extinguish for the duration of the 5 minutes and illuminated thereafter. A double actuation of switch  420  will cancel future voice alerts as to the baggage or utility door ajar. Door position is determined by an airframe mounted microswitch.  
         [0111]    As to navigation, the GPS message alert may be illuminated for extended periods of time. As such, audio  450  does not provide an audible alert but rather the lighted message of display module  300 B remains illuminated. In addition, the WPT alert has a rearming delay of 25 seconds. That is, following a WPT voice alert, a 25 second reset timer is started and the WPT alert will not activate again until the reset timer has completed its cycle.  
         [0112]    Voice annunciations can be muted by the user. A single activation of switch  315  on display module  300 A will silence all voice annunciations and save any further notifications in a queue. Stored notifications can be played audibly, via audio  450  by activating, or pressing, switch  315  once.  
         [0113]    Upon a single activation of switch  315 , the present system enters a mute mode and all voice annunciations are muted. Warning or caution conditions detected after the single activation continue to cause the appropriate light to be illuminated (either warning light  410  or caution light  415 ) however voice annunciations are not immediately presented via audio  450 .  
         [0114]    While in the mute mode, if switch  315  is again activated, then the present system returns to a normal mode and the voice annunciations are again presented via audio  450  upon detection. In addition, the stored contents of a queue are presented by voice annunciation via audio  450 . For example, if a cabin door ajar condition is detected while in a mute mode, and the door remains ajar upon return to normal mode, then a voice annunciation will be presented.  
         [0115]    Double clicking switch  315  will cause a voice annunciation to be repeated. If, as in the example above, the door is no longer ajar, then double clicking switch  315  will not cause the voice annunciation to be repeated. If the cabin door condition is no longer detected when the switch  315  is double clicked, then the door ajar annunciation will not be presented. To double click, the operator activates switch  315  twice in a period of approximately 2 or 3 seconds.  
         [0116]    To reduce false or nuisance alarms, hysteresis is provided. In one embodiment, the hysteresis is 5% of the value established as a set point. For example, if an airspeed set point is established at 40 knots, then for the condition to be cleared, the airspeed has to be raised to 42 knots. Airspeeds below 42 knots will continue to generate an alert.  
         [0117]    As to fuel level alarms, the software analyzes  1024  transducer voltage samples and averages those sample over a period of 100 seconds.  
         [0118]    As to stall warnings, a 1.20 second delay is introduced to prevent false warnings due to buffeting and vane flutter.  
         [0119]    As to landing gear warnings, the gear is given a 20 second period of time to transition from up to down or down to up. In addition, the present system provides approximately 5 second delay before annunciating up and down lights on simultaneous.  
         [0120]    All alerts are provided as herein described in the order of activation. Flight critical annunciation, including stall warning, accelerated stall warning and V NE  warning, will take priority over all other annunciations. In other words, if a cabin door alert is being annunciated at a time when an accelerated stall is detected, the cabin door voice annunciation will be interrupted and the accelerated stall warning voice annunciation will be presented.  
         [0121]    The bus voltage during certain phases of aircraft operations may drop harmlessly. For example, before take-off and after landing the engine RPM may not be continuously high enough to maintain the minimum bus voltage requirements. To prevent false low voltage alerts the present subject matter starts a timer for a 5 minute delay when the bus voltage drops below the low set point. If the bus voltage remains below the low set point after the time delay, the system will initiate the low voltage alert.  
         [0122]    A self-test of the entire system is executed upon powering the system. A voice annunciation confirming satisfactory results of the self-test is provided via audio  450 .  
         [0123]    In addition, a watchdog continuous system test operates in the background and conducts self-checks. Also, carbon monoxide sensor  135 A is self-tested every 20 minutes or 2 minutes, depending on the raw pulse count. For example, if the pulse count is less than 500, then the self-test is performed every 20 minutes and if the pulse count is greater than 500, then self-test is performed every 2 minutes. The self-test of sensor  135 A entails adjusting the heater power level and monitoring for a change in count.  
         [0124]    [0124]FIG. 3 illustrates a portion of the present subject matter. In the figure, MPU  200 C is coupled to cylinder head temperature thermocouple  530 B. Cylinder head temperature (CHT) thermocouple  530 B provides an electrical voltage signal based on a measured temperature. CHT thermocouple  530 B is illustrated as a ring J-type thermocouple and is sized for installation in lieu of a spark plug gasket in a particular cylinder.  
         [0125]    In one embodiment, the junction temperature of the thermocouple wire with the aircraft wire within the wiring harness is measured. The junction temperature is measured with a thermistor and the thermistor is co-located at the junction point.  
         [0126]    The cylinder temperature signal corresponds to a voltage signal input having a selectable range from 0 to 15 millivolts (mV) with a scale factor of 300° C.=4.5 v and with a gain of 300. Analog to digital accuracy is 10 bit, sample rate is 100 mS and latency is 100 mS.  
         [0127]    Left fuel gauge  535 C and right fuel gauge  535 D are coupled to MPU  200 C. For example, a fuel level transducer in the left fuel tank is coupled to left fuel gauge  535 C and a parallel wire connection is established to MPU  200 C. For both the left and right fuel gauge, the input signal from the fuel gauge is a DC voltage and is selectable in the range of 0-100 mV, 0-1 v, 0-14 v and 0-28 v with scale factors of gain equal to 50, 5, 0.32 and 0.16, respectively. Analog to digital accuracy is 10 bit, sample rate is 100 mS and latency is 100 mS.  
         [0128]    Pitot heat switch  545 B is the source for the pitot heat sensor input and is coupled to MPU  200 C as illustrated in the figure. An input signal is received from a 0-20 A current sensor and from the pitot heat switch. The current sensor input signal is a 0-200 mV signal with a scale factor of gain  20 . The pitot heat switch position sensor supplies 0-28 VDC.  
         [0129]    Fuel pump switch  555 B is the source for the fuel pump sensor input and is coupled to MPU  200 C as illustrated in the figure. An input signal is received from a 0-20 A current sensor and from the fuel pump switch. The current sensor input signal is a 0-200 mV signal with a scale factor of gain  20 . The fuel pump switch position sensor supplies 0-28 VDC.  
         [0130]    Oil pressure sending unit  570 B is the source for the engine oil pressure input and is coupled to MPU  200 C as illustrated in the figure. The input signal is hardware selectable via dip switches and measures 0-100 psi. The input signal is selectable as a 0-100 mV or a 0-5 VDC signal with a scale factor of gain 50 or 1, respectively. Sending unit  570 B includes a threaded portion that couples with a source of engine oil pressure.  
         [0131]    Oil temperature gauge  565 B is coupled to MPU  200 C. An oil temperature transducer is coupled to oil temperature gauge  565 B and a parallel wire connection is established to MPU  200 C. In one embodiment, the oil temperature gauge includes an RTD element. The RTD element which drives the oil temperature gauge is also used to provide an electrical signal to MPU  200 C. The present system does not interfere with an existing circuit of aircraft  12  to the extent of changing how the circuit operates, thus preserving the original certification of aircraft  12  issued by the Federal Aviation Administration (FAA).  
         [0132]    Variations in aircraft supply bus voltage can affect the temperature indicated on the oil temperature gauge. To compensate for variations in bus voltage, a software routine is executed using the data received from the oil temperature transducer. A calibration harness is used to calibrate the temperature circuit.  
         [0133]    The oil temperature transducer output voltage varies with changes in the aircraft bus voltage. To calibrate for variations in bus voltage, a calibration harness is used temporarily. When connected, the calibration harness replaces the existing oil temperature transducer with a manually adjustable potentiometer. The oil temperature circuit can be modeled with the circuit shown in FIG. 6. Bus voltage at  660 C is set to 14 VDC (or 28 VDC depending on the aircraft bus requirement). Potentiometer  610  is adjusted to cause aircraft oil temperature gauge  565 D to indicate full scale deflection (FSD), or redline. The operator activates a switch coupled to MPU  200 C to trigger the processor to receive and store bus voltage and the potentiometer voltage. In various embodiments, the switch coupled to MPU  200 C includes a momentary contact switch, a soft switch on computer  650 , or a key on the keyboard or mouse of computer  650 .  
         [0134]    Next, bus voltage at  660 C is set to 12 VDC (or 24 VDC, as appropriate) and again, the potentiometer is manually adjusted to cause aircraft oil temperature gauge  565 D to indicate FSD. The operator activates a switch to trigger the processor to gather bus voltage data and the potentiometer voltage. In various embodiments, the switch includes a momentary contact switch, a soft switch on computer  650 , or a key on the keyboard or mouse of computer  650 . Using these data points, the processor determines a set point for the alert as a function of bus voltage.  
         [0135]    The set points define those conditions for which the pilot is to be warned. The processor executes instructions to determine the oil temperature set point based on the measured bus voltage. The software provides an alert set point that varies with changes in the bus voltage.  
         [0136]    In one embodiment, a power supply is coupled to the aircraft bus and provides a supply voltage at 14 VDC. The potentiometer is adjusted to drive the oil temperature gauge to FSD and the potentiometer output voltage is measured and stored by the processor. Next, the 14 VDC supply is removed and the bus voltage is allowed to drift down to nominal battery voltage, approximately 12 VDC. Again, the potentiometer is adjusted to put the temperature gauge needle at FSD and the potentiometer voltage is measured and stored. Sample calibration data is as follows:  
                                       gauge indication   bus voltage, VDC   RTD voltage, VDC                   100° F.   14.0   0.913       100° F.   12.0   0.783       200° F.   14.0   1.024       200° F.   12.0   0.878                  
 
         [0137]    In one particular aircraft, it has been observed that for temperature variations of approximately 5° F. in the range of 100° F., the RTD voltage varies by approximately 0.13 VDC. For temperature variations of approximately 5° F. in the range of 200° F., the RTD voltage varies by approximately 0.146 VDC.  
         [0138]    The measured voltages can be modeled by a non-linear equation and the processor of MPU  200 B calculates the proper voltage levels for which to signal an alert to the pilot.  
         [0139]    Returning to FIG. 3, gear up light  585 B and gear down light  590 B provide electrical connections for coupling landing gear data to MPU  200 C.  
         [0140]    Aircraft navigation data, symbolized by block  502 B, is provided to MPU  200 C. Waypoint data is provided on line  510 B. For example, when an on-board GPS receiver indicates a waypoint passage alert, an active low signal is generated by the GPS receiver on line  510 B. The waypoint passage is annunciated as described herein. In addition, when a VOR/LOC is selected for the primary navigation, an active low is generated by the GPS receiver on line  515 B and when the GPS receiver is selected for primary navigation, an active low is generated by the GPS receiver on line  520 B. When the primary navigation system issues a message, a active low signal is generated on line  525 B.  
         [0141]    [0141]FIG. 4A illustrates the electrical connections between MPU  200 D, fuel gauge  535  and fuel tank transducer  536 . The functionality of fuel gauge  535  is not affected by MPU  200 D.  
         [0142]    [0142]FIG. 4B illustrates the electrical connections between MPU  200 E, oil temperature gauge  565 C and oil temperature probe  565 D. The functionality of oil temperature gauge  535 C is not affected by MPU  200 D.  
         [0143]    [0143]FIG. 5A illustrates current sensor  557  adapted to sense current flow between fuel pump switch  555 C and fuel pump  556  as supplied by aircraft bus  660 A. Current sensor  557  provides a signal to MPU  200 F indicative of current in fuel pump  556  without affecting the operation of the fuel pump system.  
         [0144]    [0144]FIG. 5B illustrates current sensor  547  adapted to sense current flow between pitot heat switch  545 B and pitot heat mast  546  as supplied by aircraft bus  660 B. Current sensor  547  provides a signal to MPU  200 G indicative of current in pitot heat mast  546  without affecting the operation of the pitot heat system.  
         [0145]    [0145]FIG. 7 graphically illustrates measured data corresponding to oil temperature RTD voltage as a function of temperature for output bus voltages at 12.0, 13.0 and 14.0 VDC. The data lines of the graph have different slopes with the lines diverging at higher temperatures. A best fit calculation is performed by software of the present subject matter to determine the calibrated temperature at different bus voltage levels.  
         [0146]    [0146]FIGS. 8A, 8B,  8 C and  8 D illustrate air line connections to various instruments. FIG. 8A illustrates airspeed indicator  702  coupled to sensor module  100 B by line  704  and to a pitot mast by line  705  via tee connector  703 . FIG. 8B illustrates artificial horizon or directional gyroscope  706  coupled to sensor module  100 B by line  709  and to a vacuum source by line  708  via tee connector  707 . FIG. 8C illustrates airspeed indicator  710  coupled to sensor module  100 B by line  713  and to a static port by line  712  via tee connector  711 . In one embodiment, static port is coupled to sensor module  100 B by way of a vertical speed indicator (VSI) or other static port instrument. FIG. 8D illustrates an installation in a pressurized aircraft having vacuum gauge  714  coupled to sensor module  100 B by line  717  and to pressure dump line or gyros  716  via tee connector  715 . Line  717  is coupled to an ambient air source external to the pressurized cabin.  
         [0147]    Annunciated Checklist  
         [0148]    In one embodiment, a checklist of procedures is presented via audio  450 . Each checklist item is presented via voice annunciation. A user operable switch will cause MPU  200 B to annunciate the next item on the checklist. In one embodiment, the user operable switch is switch  420 . In one embodiment, the user operable switch is switch  315 . In one embodiment, the checklist items are annunciated in sequential order with each item separated from the last by a predetermined time delay.  
         [0149]    The checklist may relate to procedures to be executed by airmen in an emergency situation. For example, an emergency engine out procedure is performed when the present subject matter determines that the airspeed is greater than a stall speed and within a predetermined period of time, the vacuum, engine oil pressure and bus voltage indicate a drop, then the present system provides an alert to signal an engine failure. In one embodiment, the predetermined time is approximately ten seconds in duration. In one embodiment, an engine failure alert includes illuminating warning light  410 , voice annunciating “engine out, left” or “engine out, right” via audio  450 , and illuminating a message “ENGINE OUT—L” or “ENGINE OUT—R” on display module  300 B. Following the voice annunciation signaling an engine outage, the present system then presents the engine out procedures for that particular aircraft.  
         [0150]    Consider an example with a left engine failure immediately after take-off. In this case, the engine out procedure calls for application of opposite side rudder and retract gear if the gear is down, feathering the dead engine propeller followed by other procedures. Main processor unit  200 B indicates that the airspeed is greater than 40 knot stall speed and that within a ten second period of time, the left engine has lost vacuum, the engine oil pressure has fallen and the bus voltage has dropped. Main processor unit  200 B illuminates warning light  410 , voice annunciates “engine out, left” and illuminates message “ENGINE OUT—L” on display module  300 B. In one embodiment, switch  410  is used to transition to the checklist annunciation routine. In one embodiment, MPU  200 B automatically transitions to the checklist annunciation routine. Main processor unit  200 B presents voice annunciation “apply right rudder” via audio  450 . Upon application of right rudder, the pilot is instructed to activate switch  410 . Activation of switch  410  causes MPU  200 B to present the next voice annunciation. If data accessible to MPU  200 B indicates that the landing gear is in the down position, then, the next voice annunciation includes “retract landing gear” presented via audio  450 . If data accessible to MPU  200 B indicates that the landing gear is in the up position, then, the next voice annunciation includes “feather left engine” presented via audio  450 . Subsequent procedures are presented via voice annunciation in an order as determined by stored data provided by the aircraft manufacturer and the conditions detected by MPU  200 B.  
         [0151]    As another example, in one embodiment, a landing gear manual extension procedure is presented to the pilot via audio  450  in the event of a landing gear malfunction as determined by MPU  200 B.  
         [0152]    Double clicking switch  315  will cause the present checklist item to be repeated aloud via audio  450 .  
         [0153]    Self-Diagnostics  
         [0154]    The present subject matter includes programming to execute one or more self-diagnostic routines as noted below.  
         [0155]    If carbon monoxide sensor  135 A indicates an abnormal condition during a self-test, an annunciated message such as “carbon monoxide sensor failure” is presented via audio  450 . Additional explanatory information can also be provided by audio  450 .  
         [0156]    If carbon monoxide sensor  135 A successfully passes a self-test, then an annunciated message such as “carbon monoxide sensor ready” is presented via audio  450 . Notification of successful testing can occur anytime after a carbon monoxide sensor failure message.  
         [0157]    If accelerometer  130 A indicates an abnormal self-test, an annunciated message such as “please calibrate g sensor” is presented via audio  450 . Additional explanatory information, such as circuit breaker identification and technical support information, can also be provided by audio  450 .  
         [0158]    If the present system detects a communication failure between MPU  200 B and sensor module  100 B, an annunciated message such as “system failure” is presented via audio  450 . Additional explanatory information can also be provided by audio  450 .  
         [0159]    [0159]FIG. 9 illustrates a flow chart for receiving information from the acknowledge switch. In the figure, the acknowledge switch controls the volume control, training mode and provides voice annunciation of software versions.  
         [0160]    [0160]FIG. 10 illustrates a flow chart for receiving pilot selections using the display module switch. The display module switch permits control of voice annunciation, replay of active messages and re-boots MPU  200 B.  
         [0161]    [0161]FIG. 11 illustrates a flow chart for a method based on actuation of the acknowledge switch. The method executed depends on the number of times the acknowledge switch has been pushed and whether a caution message is being alerted.  
         [0162]    [0162]FIG. 12 illustrates a flow chart for a method based on actuation of the acknowledge switch. The method executed depends on the number of times the acknowledge switch has been pushed and whether a warning message is being alerted.  
         [0163]    [0163]FIG. 13 illustrates a flow chart for a method based on monitoring the pitot heat sensor. The present subject matter determines if the pitot heat current is within predetermined limits.  
         [0164]    [0164]FIG. 14 illustrates a flow chart for a method based on monitoring the fuel pump sensor. The present subject matter determines if the fuel pump current is within predetermined limits.  
         [0165]    [0165]FIG. 15 illustrates a flow chart for a method based on the stall warning vane switch. A warning message is presented if a stall is threatened.  
         [0166]    [0166]FIG. 16 illustrates a flow chart for a method based on detecting a threatened accelerated stall. The method includes calculating an accelerated stall set point.  
         [0167]    [0167]FIGS. 17A and 17B illustrate flow charts for a method based on alerting for a waypoint. The methods include annunciating depending on the status of a waypoint flag.  
         [0168]    [0168]FIGS. 18A and 18B illustrate flow charts for methods based on fuel quantity indications. In FIGS. 18A and 18B, the fuel quantity remaining in the left fuel tank and right fuel tank, respectively, can trigger a caution alert.  
         [0169]    [0169]FIG. 19 illustrates a flow chart for a method based on remaining fuel in a tank. For sufficiently low fuel quantities, a warning alert is generated.  
         [0170]    [0170]FIG. 20 illustrates a flow chart for a method based on a measured oil temperature. The oil temperature alert set point is based on the aircraft bus voltage.  
         [0171]    [0171]FIG. 21 illustrates a flow chart for a method based on carbon monoxide levels detected in the aircraft cabin. For detected levels of carbon monoxide, a warning alert is issued with voice annunciation.  
         [0172]    [0172]FIG. 22 illustrates a flow chart for a method based on detected oil pressure levels. For oil pressures below a predetermined level, a warning alert is issued with voice annunciation.  
         [0173]    [0173]FIG. 23 illustrates a flow chart for a method based on detected oil pressure levels and vacuum levels. For oil pressures and vacuum levels below a predetermined level, a warning alert message is annunciated.  
         [0174]    [0174]FIG. 24 illustrates a flow chart for a method based on landing gear position sensor switches. Audible messages are presented to the pilot depending on the switch positions.  
         [0175]    [0175]FIG. 25 illustrates a flow chart for a method based on landing gear position sensor switches.  
         [0176]    [0176]FIG. 26 illustrates a flow chart for a method based on aircraft bus voltages. Caution and warning alerts are issued based on measured voltage levels.  
         [0177]    Alternative Embodiments  
         [0178]    Variations of the above embodiments are also contemplated, including the following:  
         [0179]    In one embodiment, if messages appearing on display module  300 B cycle in a rotating pattern, communications between MPU  200 B and display module  300 B have likely failed. An annunciator message, including data as to possible connection failures, is presented via audio  450 .  
         [0180]    In one embodiment, different messages, different colors and different notification systems can be utilized. For example, in one embodiment, one or more turbine engines power the aircraft and annunciated conditions and parameters include those selected from the following: ignition; beta; chip detector; inlet heat; fuel selector; fuel bypass; generator; fuel pressure; start (start cycle); torque; ITT (inlet turbine temperature); turbine RPM; fan rpm; N 1 , N 2 ; rotor RPM; anti-icing; de-ice; function of navigation and anti-collision lights; auxiliary fuel tanks; hydraulic pressure; fuel filter; and fuel heaters. In addition to those conditions and parameters presented earlier, for piston powered aircraft, the following may be included for annunciation: TIT (turbine inlet temperature); flap setting; anti-icing; de-ice; function of navigation and anti-collision lights; auxiliary fuel tanks. Furthermore, aircraft engaged in agricultural or fire fighting operations may include annunciation of conditions and parameters related to hopper level and spray pressure. Other conditions and parameters can be monitored, including avionics messages, trim position or trim failures and navigation signal loss. Those conditions and parameters presented herein are not to be taken in a limiting sense but are presented for example only.  
         [0181]    In one embodiment, a voice only annunciation is provided and no lights or visual messages are presented.  
         [0182]    In one embodiment, the present subject matter monitors flight, engine, and navigation systems for a multi-engine aircraft.  
         [0183]    In one embodiment, acknowledge switch  400 A and display  300 A are installed within a single housing.  
         [0184]    In one embodiment, acknowledge switch  400 A includes a user-operable switch in a first location and a visual display portion in a second location. In one embodiment, acknowledge switch  400 A includes a first and second panel-mounted light and a yoke-mounted or panel-mounted switch.  
         [0185]    In one embodiment, switch  315  is not included and acknowledge switch  400 A performs functions described herein as being performed by switch  315 .  
         [0186]    In one embodiment, switch  315  and acknowledge switch  400 A include a soft switch, a touch-sensitive switch or other user operable switch. The switches may be momentary contact or latching type contact switches.  
         [0187]    In one embodiment, display module  300 A and acknowledge switch  400 A include plastic housings and are adapted to receive mounting hardware for installation in, or on, an instrument panel of aircraft  12 .  
         [0188]    In one embodiment, a different warning light, caution light and display light color is used. In one embodiment, set point levels and timer durations and delay durations are user or operator selectable.  
         [0189]    In one embodiment, programming described herein is executed on a processor of MPU  200 B, a processor of display module  300 B or sensor module  100 B.  
         [0190]    In addition to J-type or K-type thermocouples and RTDs, other types of temperature sensors are also contemplated. For example, in one embodiment, a temperature probe is adapted for threading into a tapped bore of a engine cylinder or other component. In one embodiment, an exhaust gas temperature sensor is used. A J-type thermocouple is one fabricated of iron and constantan and K-type thermocouple is one fabricated of chromel and alumel.  
         [0191]    In one embodiment, engine oil temperature is monitored by an alternative configuration. For example, in one embodiment, the processor is coupled to the bus voltage and the voltage at the potentiometer and the processor receives and stores potentiometer voltages at different bus voltages for selected gauge indications. The calibration routine in the software then establishes set points.  
         [0192]    In one embodiment, the calibration routine in the software establishes a best-fit curve, based on selected measured values, to determine set points. In one embodiment, the software accesses a stored look-up table of data to determine set points. In one embodiment, the software interpolates data between measured values to determine set points.  
         [0193]    In one embodiment, the processor regulates the bus voltage and the operator adjusts a potentiometer to set the gauge to FSD. In one embodiment, the operator activates a button to cause the processor to receive and store the bus voltage and the potentiometer voltage. In one embodiment, potentiometer voltages are measured at bus voltages set to 12.0 VDC and 14.0 VDC. In one embodiment, potentiometer voltages are measured at bus voltages set to value less than 12.0 VDC and greater than 14.0 VDC.  
         [0194]    In one embodiment, the relationship between sensor output voltage and bus voltage is approximated by a stored non-linear equation. In one embodiment, the relationship between sensor output voltage and bus voltage is approximated by a stored linear equation.  
         [0195]    In one embodiment, digital logic levels different than those described herein are used.  
         [0196]    In one embodiment, monitoring of a parameter or condition is suspended temporarily for conditions outside of a predetermined range.  
         [0197]    In one embodiment, the carbon monoxide sensor includes a solid chemical sensor element.  
       Conclusion  
       [0198]    The above description is intended to be illustrative, and not restrictive. Many other embodiments will be apparent to those of skill in the art upon reviewing the above description.