Abstract:
In a method for stand-alone alignment of an inertial unit for an onboard instrument capable of being mounted in an aircraft, the method includes monitoring the appearance of a movement of the inertial unit during the alignment, suspending the alignment of the inertial unit in the event of the appearance of movement, and resuming the alignment of the inertial unit on the disappearance of the movement.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a National Stage of International patent application PCT/EP2008/066663, filed on Dec. 2, 2008, which claims priority to foreign French patent application No. FR 07 09034, filed on Dec. 21, 2007, the disclosures of which are incorporated by reference in their entirety. 
     FIELD OF THE INVENTION 
     The invention relates to a method for stand-alone alignment of an inertial unit for an onboard instrument capable of being mounted in an aircraft, and an onboard instrument using such a method. The invention relates in particular to a method for stand-alone alignment of an inertial unit for a stand-by instrument generating and displaying information relating to the speed, altitude and attitude of an aircraft. It is particularly useful for the alignment of inertial units mounted in aircraft which can take off from non-stabilized platforms, such as oil platforms, aircraft carriers or helicopter carriers. However, it can also be applied for the alignment of inertial units mounted in aircraft taking off from stable platforms, such as an airstrip, insofar as the aircraft may be subjected to movements, even when stopped, for example due to wind or ground support facilities around the aircraft. 
     BACKGROUND OF THE INVENTION 
     Stand-by instruments are stand-alone onboard instruments which generate and display flight information which is essential to the pilot of an aircraft in the event of failure of primary onboard instruments. This flight information, generally obtained with less precision than that of the primary onboard instruments, essentially relates to the speed, altitude and attitude of the aircraft. In order to guarantee the stand-alone capability of the stand-by instruments in relation to the primary onboard instruments, the stand-by instruments must have their own sensors in order to generate and display the speed, altitude and attitude of the aircraft. In particular, the stand-by instruments normally comprise a static pressure sensor, a total pressure sensor and an inertial unit. 
     The static and total pressure sensors are connected respectively to a static pressure connector and a total pressure connector located on the skin of the aircraft. The static pressure allows the altitude of the aircraft to be determined. The difference between the total pressure and the static pressure allows the speed of the aircraft to be determined in relation to the air. 
     The inertial unit comprises 3 gyrometers and 2 or 3 accelerometers. The gyrometers measure the speed of rotation of the sensor referential, here a system of axes linked to the stand-by instrument, in relation to an inertial referential. Through integration of the rotation speeds, it is possible to identify the position of the stand-by instrument in relation to the inertial referential and therefore, knowing the position of the stand-by instrument in relation to the aircraft and the position of the local geographical frame of reference in relation to the inertial frame of reference, it is possible to identify the position of the aircraft in relation to the local geographical frame of reference. The position of the aircraft in relation to the local geographical frame of reference, referred to as the attitude of the aircraft, is determined in relation to a roll axis, a pitch axis and a yaw axis, and the movements around these axes are referred to respectively as the roll, pitch and yaw. Accelerometers measure non-gravitational forces applied to the aircraft, from which translation accelerations of the sensor referential in relation to the inertial referential are deduced. The combination of gyrometers and accelerometers enables a precise determination of the attitude of the aircraft, the data supplied by the accelerometers being used in preference to the data supplied by the gyrometers in the static or quasi-static flight phases, and the data supplied by the gyrometers being used in preference to the data supplied by the accelerometers during the dynamic phases of the flight. 
     When an aircraft, and in particular a stand-by instrument, is powered up, the inertial unit of the stand-by instrument must be initialized in order to supply the most reliable attitude information possible during the flight. This initialization includes an alignment phase, consisting notably in estimating the drift of the different gyrometers, i.e. the speed of rotation measured by the gyrometers in the absence of any movement of the latter. The gyrometers being electronic sensors, their drift may differ between two different power-ups of the inertial unit, to the point of rendering unusable any measurement carried out by these gyrometers and therefore any attitude displayed by the stand-by instrument. It is therefore necessary to determine the drift of the gyrometers on every power-up. Moreover, the alignment of the gyrometers must be carried out in the absence of any movement of the inertial unit, otherwise a movement of the inertial unit will be integrated into the drift of a gyrometer. 
     To ensure the correct alignment of the gyrometers of an inertial unit, it is known to check for the presence or absence of movements of the inertial unit by means of the accelerometers of the inertial unit. For the entire duration of the alignment, the accelerometers measure the non-gravitational forces of the inertial unit in relation to the inertial referential. In the event of movement of the inertial unit during the alignment, measured by the accelerometers, the stand-by instrument, at the end of the alignment, invalidates the determination of the drift of each gyrometer, displays a message indicating the detection of movement to the pilot and asks the pilot to restart the alignment either by switching off the stand-by instrument then powering it up again, or by pressing a button on the front surface of the stand-by instrument. This restart of the alignment is imperative insofar as the availability of the stand-by instrument, and therefore the alignment of the inertial unit, is a necessary condition for the aircraft take-off authorization. 
     A solution of this type presents a plurality of disadvantages. A first disadvantage is the wait for the end of the alignment in order to indicate the detection of a movement during the alignment. It is therefore only at the end of the alignment of the gyrometers that the pilot is aware of the invalidation of the alignment and can restart it. Consequently, the time elapsed between the detection of movement and the end of the alignment is lost. A second disadvantage is the loss of the estimation of the drifts carried out between the start of the alignment and the detection of a movement. At the end of the invalidated alignment, the entire alignment procedure is restarted, entailing the risk that the estimated drift has been distorted by the movement. Moreover, if the alignment is restarted by a hardware reset, i.e. by switching off the stand-by instrument then powering it up again, there is a risk that the drift of the gyrometers will change, rendering the preceding determination of the drifts obsolete. A third disadvantage is the impossibility, in certain situations, of being able to carry out an alignment. This may notably occur if the aircraft has started up on a moving platform. In most cases, the movement of the platform, for example due to the swell of the sea, cannot be prevented. The aircraft must then wait for the cessation of the movements, in this case a calming of the swell, to be able to take off. An immobilization of this type is indisputably detrimental to the economic efficiency of the aircraft. 
     SUMMARY OF THE INVENTION 
     An object of the invention is notably to overcome all or some of the aforementioned disadvantages. For this purpose, the subject of the invention is a method for stand-alone alignment of an inertial unit for an onboard instrument capable of being mounted in an aircraft. According to the invention, the method includes the following steps:
         monitoring the appearance of a movement of the inertial unit during the alignment,   suspending the alignment of the inertial unit in the event of the appearance of movement,   resuming the alignment of the inertial unit on the disappearance of the movement.       

     The subject of the invention is also an onboard instrument, capable of being mounted in an aircraft, including an inertial unit comprising means for stand-alone alignment. According to the invention, the onboard instrument includes means for monitoring the appearance of a movement of the inertial unit during the alignment, means for suspending the alignment in the event of the appearance of movement, and means for resuming the alignment of the inertial unit on the disappearance of the movement. 
     The advantage of the invention is notably that it enables a fast and reliable alignment of the inertial unit without increasing the complexity of the drift calculation algorithm. In particular, the alignment may be carried out during all periods when no movement of the inertial unit is detected, but solely during these periods. This results in an optimization of the duration of the alignment. Moreover, the alignment remains protected against movements of the inertial unit. It is even possible to increase the precision of the alignment by lowering the threshold for the detection of movements of the inertial unit, insofar as the alignment duration is optimized. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention will be more easily understood and other advantages will become clear upon reading the detailed description of an embodiment, presented by way of example, and set out with reference to the attached drawings, in which:
           FIG. 1  shows a stand-by instrument capable of being mounted in an aircraft;     FIG. 2  shows an aircraft fitted with the stand-by instrument shown in  FIG. 1 , the aircraft and the stand-by instrument each having a system of axes;     FIG. 3  shows a synopsis of the means implemented by the stand-by instrument for calculating and displaying the attitude of the aircraft;     FIG. 4  shows an example of steps implemented for the initialization of the stand-by instrument;     FIG. 5  shows an example of steps implemented according to the invention for the initialization of the stand-by instrument;     FIG. 6  shows, in the form of a chronogram, an example of the initialization of the stand-by instrument during which a movement is detected;     FIG. 7  shows, in the form of a chronogram, another example of the initialization of the stand-by instrument during which movements are detected;     FIG. 8  shows an illustration of different times involved during the initialization of the stand-by instrument.       

     
    
    
     DETAILED DESCRIPTION 
     The description which follows is provided in relation to a stand-by instrument. It is obviously possible to implement the invention on the basis of any onboard instrument comprising an inertial unit. 
       FIG. 1  shows a stand-by instrument  1  capable of being mounted in an aircraft. The stand-by instrument  1  comprises a display  2 , for example a liquid-crystal screen. The display  2  displays flight information essential to the pilot for controlling the aircraft. This information concerns notably the air speed, altitude and attitude of the aircraft. The air speed and altitude of the aircraft are shown in the form of rotating vertical graduated scales, a scale  3  indicating the air speed of the aircraft and a scale  4  indicating the altitude of the aircraft. The attitude of the aircraft is symbolized by a horizon line  5  movable in relation to a fixed silhouette  6  representing the aircraft. The air speed and altitude information is obtained from anemo-barometric sensors connected on the one hand to pressure connectors disposed on the skin of the aircraft and, on the other hand, to a calculator. The anemo-barometric sensors supply a static pressure Ps and a total pressure Pt of the air surrounding the aircraft on the basis of which the calculator determines the air speed and altitude of the aircraft. The attitude of the aircraft is obtained from an inertial unit including gyrometers and accelerometers, as explained below. The anemo-barometric sensors, the inertial unit and the calculator form means for determining the flight parameters. These determination means are stand-alone, as they belong to the stand-by instrument and can function with no external information other than the information originating from the pressure connectors. 
       FIG. 2  shows an aircraft  20  equipped with the stand-by instrument  1  and  FIG. 3  shows a synopsis of the means implemented by the stand-by instrument  1  to calculate and display the attitude of the aircraft  20 . The inertial unit of the stand-by instrument  1  generally includes 3 gyrometers  30  and  3  accelerometers  31 . The gyrometers  30  measure angular speeds {right arrow over (Ω is )} of a frame of reference R is ({right arrow over (X)} is ,{right arrow over (Y)} is ,{right arrow over (Z)} is ) linked to the stand-by instrument  1  in relation to an inertial frame of reference. For the sake of readability of the description, the movements of the earth can be ignored and the local geographical frame of reference point, referred to as the terrestrial frame of reference R t ({right arrow over (X)} t ,{right arrow over (Y)} t ,{right arrow over (Z)} t ) can be considered in the description below as being the inertial frame of reference. However, for an implementation of the invention, it will be possible to take account of the movements of the earth in relation to the inertial frame of reference. As shown in  FIG. 3 , the angular speeds {right arrow over (Ω is )} of the inertial unit are corrected by means of an operator  32  of the internal drifts {right arrow over (dΩ c )} of the inertial unit. The internal drifts {right arrow over (dΩ c )} are, for example, stored in a RAM memory  33 . Means for determining the internal drifts {right arrow over (dΩ c )} will be described below. The angular speeds thus corrected and denoted {right arrow over (Ω c )} in the frame of reference R t ({right arrow over (X)} t ,{right arrow over (Y)} t ,{right arrow over (Z)} t ) are transformed to obtain the angular speeds {right arrow over (Ω a )} of a frame of reference linked to the aircraft  20  and denoted R a ({right arrow over (X)} a ,{right arrow over (Y)} a ,{right arrow over (Z)} a ) in relation to the frame of reference R t ({right arrow over (X)} t ,{right arrow over (Y)} t ,{right arrow over (Z)} t ). Similarly, the accelerometers  31  measure translation accelerations {right arrow over (γ is )} of the frame of reference R is ({right arrow over (X)} is ,{right arrow over (Y)} is ,{right arrow over (Z)} is ) linked to the stand-by instrument  1  in relation to the terrestrial frame of reference R t ({right arrow over (X)} t ,{right arrow over (Y)} t ,{right arrow over (Z)} t ). These translation accelerations {right arrow over (γ is )} are also transformed to obtain the translation accelerations {right arrow over (γ a )} of the frame of reference R a ({right arrow over (X)} a ,{right arrow over (Y)} a ,{right arrow over (Z)} a ) linked to the aircraft  20  in relation to the frame of reference R t ({right arrow over (X)} t ,{right arrow over (Y)} t ,{right arrow over (Z)} t ). The angular speeds {right arrow over (Ω a )} and the translation accelerations {right arrow over (γ a )} allow the attitude of the aircraft  20  to be determined in relation to the terrestrial frame of reference R t ({right arrow over (X)} t ,{right arrow over (Y)} t ,{right arrow over (Z)} t ) by means of a calculator  34  belonging to the inertial unit. Advantageously, the calculator  34  also carries out the transformations of angular speeds and of translation accelerations. In one particular embodiment, only the translation accelerations {right arrow over (γ a )} are used to determine the attitude of the aircraft  20  when it is in stabilized flight. Conversely, only the angular speeds {right arrow over (Ω a )} are used to determine the attitude of the aircraft  20  when it is in dynamic flight. Other embodiments are possible. In particular, it is possible to use a weighting of the translation accelerations {right arrow over (γ a )} and of the angular speeds {right arrow over (Ω a )} to determine the attitude of the aircraft  20 , said weighting being variable according to the flight conditions. The attitude of the aircraft  20  is displayed on the display  2  of the stand-by instrument  1 . 
       FIG. 4  shows steps implemented for the initialization of a stand-by instrument  1 . On the power-up of the stand-by instrument  1 , a rough estimation, referred to as the setup  42 , of the internal drifts {right arrow over (dΩ c )} of the inertial unit is carried out. This setup  42  allows a minimum value and a maximum value to be determined for each drift of the gyrometer  30 . At the end of the setup  42 , a fine alignment, also referred to as the alignment  43 , is carried out to determine precisely the drift of each gyrometer  30 . The alignment  43  comprises notably a step of measurement of the instantaneous drifts and a step of filtering of these drifts to obtain precise values of the internal drifts {right arrow over (dΩ c )}. The setup  42  and the alignment  43  are referred to as the global alignment  40 . During this global alignment  40 , the accelerometers  31  monitor the movements of the inertial unit. At the end of the alignment  43 , a control step  44  examines whether the movements detected by the accelerometers  31  have not exceeded a predefined threshold and if the internal drifts {right arrow over (dΩ c )} are between the minimum and maximum values determined during the setup  42 . If so, the internal drifts {right arrow over (dΩ c )} are recorded in the RAM memory  33  of the stand-by instrument  1 . The stand-by instrument  1  is ready for the navigation  45  and displays the information relating to the attitude of the aircraft  20 . In all other cases, the internal drifts {right arrow over (dΩ c )} are invalidated and the global alignment  40  is restarted according to the reference sign  46 . 
     The steps implemented in  FIG. 4  do not enable a fast alignment of the inertial unit if movements are detected. In particular, the time elapsed between the detection of a movement and the end of the alignment is lost, the global alignment  40  being invalidated. Similarly, the time elapsed between the start of the setup  42  and the detection of the movement is also lost, the intermediate drifts not being re-used for the subsequent global alignment  40 . Moreover, these steps do not enable a global alignment  40  of the gyrometers  30  if the movements are repeated and, in particular, if the time elapsing between two movements is each time less than the duration necessary for the global alignment  40 . 
     According to the invention and as shown in  FIG. 5 , following the power-up  41  of the stand-by instrument  1 , the movements of the inertial unit are monitored during the setup  42  and the alignment  43 . If movement is detected, the setup  42  or the alignment  43  of the inertial unit is suspended. If no more movement is detected, the setup  42  or the alignment  43  of the inertial unit  43  is resumed. The detection of a movement, the suspension of the global alignment  40  and its resumption are shown by the reference sign  51 . At the end of the alignment  43 , the internal drifts of the gyrometers  30  are recorded, for example in the RAM memory  33  of the stand-by instrument  1 , and the stand-by instrument  1  can be used to determine and display the attitude of the aircraft  20 . 
       FIG. 6  shows, in the form of a chronogram, an example of the initialization of the stand-by instrument  1  during which a movement is detected by an accelerometer  31  according to an axis of the stand-by instrument  1 . The time is shown on the x-axis and the movements are shown on the y-axis. For this example, the amplitude of the acceleration of the movement is considered. However, other types of movement can be monitored. On the power-up  41  of the stand-by instrument  1  at a time t 0 , the setup  42  of the inertial unit is carried out for a period T 1 , generally around ten seconds. This setup  42  allows a minimum value and a maximum value to be defined for each drift of the gyrometers  30 . At the end of the setup  42 , from a time t 1  and for a duration T 2 , the fine alignment  43  of the inertial unit is carried out until a time t 2  from which a movement is detected. For an entire duration T 3  when the movement is detected, i.e. between the times t 2  and t 3 , the alignment  43  is suspended. Advantageously, only the movements exceeding a determined amplitude, referred to as the threshold  61 , suspend the global alignment  40 . For the duration T 3 , the determination of the internal drifts {right arrow over (dΩ c )} is suspended. In other words, the measurements of the drifts for this duration T 3  are not taken into account in determining the internal drifts {right arrow over (dΩ c )}. Conversely, the intermediate values of drifts obtained between the times t 0  and t 2  are stored, for example, in the RAM memory  33 , to be re-used on resumption of the alignment  43 . If, at time t 3 , the amplitude of the movement again falls below the threshold  61 , the alignment  43  is resumed where it had been suspended, with the intermediate values of drifts obtained between the times t 0  and t 2 . More generally, the step of suspension of the alignment  40  of the inertial unit may include a sub-step consisting in recording current values used for the alignment  40 , and the step of resumption of the alignment  40  may include a sub-step consisting in recovering the recorded values for the continuation of the alignment  40 . 
     According to a particular embodiment, the alignment  43  has a fixed, parameterizable duration referred to as ALN_Duration. The duration ALN_Duration is generally around several tens of seconds, for example 80 seconds, and may be parameterized according to the latitude at which the aircraft  20  is located. According to this embodiment, the alignment  43  continues at time t 3  for a duration T 4  in such a way that the addition of the durations T 2  and T 4  is more or less equal to the duration ALN_Duration. A difference in duration may be explained notably by the duration necessary for the resumption of the alignment  43 . 
     In this example, it is considered that the movement is detected during the alignment  43 . However, the same method can be applied during the setup  42 . Similarly, the global alignment  40  can be interrupted and resumed an unlimited number of times. 
     In the aforementioned example, described with reference to  FIG. 6 , the detection of movement is considered as the exceeding of a threshold by an acceleration amplitude according to an axis of the stand-by instrument  1 . The invention is not limited to this form of detection and encompasses any form of movement of the stand-by instrument  1 . In particular, it is possible to detect the movements either by means of an accelerometer or by means of a gyrometer, or by a combination of accelerometers and gyrometers. Advantageously, the accelerometers and/or gyrometers of the inertial unit are used. The stand-alone capability of the stand-by instrument  1  is thus maintained. However, it can be envisaged to use sensors outside the stand-by instrument  1 . According to the instruments used, it is possible to monitor a translation acceleration and/or an angular speed of the stand-by instrument  1 . It is understood that the movement can also be monitored through the observation of a translation speed of the stand-by instrument  1 , i.e. the observation of an integrated translation acceleration. The movements can be referenced in a frame of reference R is ({right arrow over (X)} is ,{right arrow over (Y)} is ,{right arrow over (Z)} is ) linked to the stand-by instrument  1  or in a frame of reference R a ({right arrow over (X)} a ,{right arrow over (Y)} a ,{right arrow over (Z)} a ) linked to the aircraft  20 . It is possible to pass from one frame of reference to another by a simple change of frame of reference, the stand-by instrument  1  being fixed in the aircraft  20 . In a particular embodiment, the monitored movements of the stand-by instrument  1  include an angular speed around a yaw axis of the aircraft and translation speeds according to the yaw axis of the aircraft and the roll and pitch axes of the aircraft. 
     According to a particular embodiment, the method according to the invention displays on the display  2  of the stand-by instrument  1  a countdown of the remaining duration before the end of the alignment  43 . The countdown of the remaining duration is started at the start of the alignment  43 , but it can also be envisaged to start the countdown of the remaining duration at the start of the setup  42 . For this embodiment, the following are considered: 
     a fixed duration “ALN_Duration” to carry out the alignment  43  in the absence of movement, 
     a time variable “Tps_ALN_actual” representing the duration during which the drifts were estimated during the alignment  43 , 
     a time variable “Tps_ALN_remaining” corresponding to the remaining duration necessary for the alignment  43  in the absence of movements, 
     a Boolean variable “B_OTM” assuming the value “true” if a movement is detected and the value “false” if not, 
     a Boolean variable “B_ALN_Complete” assuming the value “true” if the alignment  43  has ended and the value “false” if not. 
     All the durations and time variables contain integers representing a number of seconds. At the end of the setup  42 , the variable “Tps_ALN_actual” is initialized to the value zero and the variable “B_ALN_Complete” is initialized to the value “false”. During the alignment  43 , the variable “Tps_ALN_actual” is incremented by one unit every second. The variable “Tps_ALN_remaining” is determined by the following relation:
 
 Tps   —   ALN _remaining= ALN _Duration− Tps   —   ALN _actual
 
     An example of an algorithm allowing the remaining duration necessary for the alignment  43  to be determined is set out below: 
     
       
         
               
             
           
               
                   
               
             
             
               
                 While B_ALN_Complete = false 
               
               
                   If B_OTM = false then 
               
               
                     Tps_ALN_actual ← Tps_ALN_actual + 1 
               
               
                   End if 
               
               
                   Tps_ALN_remaining ← ALN_Duration − Tps_ALN_actual 
               
               
                   If Tps_ALN_remaining &lt;= 0 then 
               
               
                     B_ALN_Complete ← true 
               
               
                   End if 
               
               
                 End while 
               
               
                   
               
             
          
         
       
     
     At the end of the alignment  43 , the display  2  of the stand-by instrument  1  can display a message indicating to the pilot that the global alignment  40  has ended. The display  2  may also directly display the attitude of the aircraft  20 . 
     In a particular embodiment, the global alignment  40  of the inertial unit is cancelled if the time elapsed since the start of the alignment  43  added to the remaining duration before the end of the alignment  43  “Tps_ALN_remaining” is greater than a maximum determined duration. For this embodiment, the following are considered in addition to the fixed duration “ALN_Duration” and the previously defined variables: 
     a fixed duration “Max_Duration” corresponding to the maximum authorized duration for the alignment  43  of the inertial unit, 
     a variable “Tps_ALN_total” representing the time elapsed since the start of the alignment  43 . This time corresponds to the duration “Tps_ALN_actual” plus the time during which movements were detected, 
     a Boolean variable “B_ALN_TooLong” assuming the value “true” if the duration necessary for the alignment  43  is greater than the maximum duration authorized for the alignment  43  of the inertial unit (Max_Duration). 
     The preceding algorithm is modified in the following manner: 
     
       
         
               
             
           
               
                   
               
             
             
               
                 While (B_ALN_Complete = false) and (B_ALN_Complete = false) 
               
               
                   Tps_ALN_total ← Tps_ALN_total + 1 
               
               
                   If B_OTM = false, then 
               
               
                     Tps_ALN_actual ← Tps_ALN_actual + 1 
               
               
                   End if 
               
               
                   Tps_ALN_remaining ← ALN_Duration − Tps_ALN_actual 
               
               
                   If Tps_ALN_remaining &lt;= 0 then 
               
               
                     B_ALN_Complete ← true 
               
               
                   End if 
               
               
                   If (Tps_ALN_remaining + Tps_ALN_total) &gt; Max_Duration 
               
               
                 then 
               
               
                     B_ALN_TooLong ← true 
               
               
                   End if 
               
               
                 End while 
               
               
                   
               
             
          
         
       
     
       FIGS. 7 and 8  illustrate this particular embodiment in which a maximum duration “Max_Duration” is authorized for the alignment  43  of the inertial unit. In the following example, it is considered that the alignment  43  is interrupted for a total duration sufficiently long so that the duration of the alignment  43  exceeds the maximum duration authorized for the alignment  43  (Max_Duration). 
       FIG. 7  shows, in the form of a chronogram, an example of the initialization of the stand-by instrument  1  during which movements are detected. The time is shown on the x-axis and the movements are shown on the y-axis. On the power-up  41  of the stand-by instrument  1  at a time t 10 , the setup  42  of the inertial unit is carried out for a period T 10 . At the end of the setup  42 , from a time t 11  and for a duration T 12 , the fine alignment  43  of the inertial unit is carried out until a time t 12  from which a movement is detected. The alignment  43  resumes after a duration T 13  at a time t 13  if no further movement is detected until a time t 14  when movements are again detected, i.e. for a duration T 14 . The alignment  43  resumes a second time at a time t 15  if no further movement is detected, i.e. after a duration T 15 . 
       FIG. 8  shows the different times used in the preceding algorithm and taken at a time t 16 , i.e. a duration T 16  after the time t 15 . These different times are shown on the horizontal axes. The duration of actual alignment “Tps_ALN_actual” corresponds to the sum of the durations T 12 , T 14  and T 16  and the time elapsed since the start of the alignment  43  (Tps_ALN_total) corresponds to the sum of the durations T 12  to T 16 . The remaining alignment duration “Tps_ALN_remaining” is obtained by subtracting the duration “Tps_ALN_actual” from the duration “ALN_Duration”. This remaining alignment duration is added to the time “Tps_ALN_total” and the sum is compared with the duration “Max_Duration”. At the time t 16 , the sum (Tps_ALN_total+Tps_ALN_remaining) is slightly greater than the duration “Max_Duration”. The Boolean variable “B_ALN_TooLong” then changes to the value “true”. Consequently, the alignment algorithm  43  is abandoned. A message can be displayed on the display  2  of the stand-by instrument to inform the pilot that the inertial unit is not aligned. The pilot can then restart the global alignment  40  or the fine alignment  43 . The global alignment  40  or the fine alignment  43  can also be restarted automatically. 
     It is of course possible to consider other embodiments for limiting the duration of the alignment  43 . In particular, the duration “Tps_ALN_actual” and the duration “Tps_ALN_total” can be incremented as from the setup  42 .