Abstract:
A method and a device for an aircraft for detecting noise in a signal of LOC type. A first step includes estimating a first lateral speed of the aircraft according to a first set of parameters. Concurrently, at least one second lateral speed of the aircraft is estimated according to at least one second set of parameters, among which at least one parameter is of different nature from each parameter of the first set of parameters. A second step includes comparing the first lateral speed and the at least one second lateral speed according to a threshold. If the difference between the first lateral speed and the at least one second lateral speed is greater than the threshold, the presence of noise in the signal of LOC type is detected.

Description:
BACKGROUND OF THE INVENTION 
     The present invention relates to automatic flight control for an aircraft and more particularly to a method and device for detecting noise on a guide signal of LOC type received by an aircraft. 
     There exist systems for assisting a pilot in landing airplanes under poor visibility conditions. Such systems are commonly known as ILS (Instrument Landing System). 
     An ILS system is generally composed of a ground portion and of a portion on board the airplanes. The ground portion generally comprises radio-frequency transmitters that permit establishment of an imaginary axis of approach to the runway by means of a horizontal radio beam and a vertical radio beam. 
     The horizontal radio beam, known as Localizer or LOC, establishes the axis of the runway, while the vertical radio beam, known as Guide Slope or GS, establishes the slope of descent of the airplane to the edge of the runway. The LOC emits a VHF (Very High Frequency) signal in the 108-118 MHz frequency band. The GS emits a UHF (Ultra High Frequency) signal in the 329-335 MHz frequency band. Thus the LOC signal is used to determine a difference between the axis of displacement of the airplane and the runway axis, and the GS signal is used to determine a difference between the axis of displacement of the airplane and the nominal approach slope. The LOC and GS beams are narrow and sensitive to perturbations. 
     Different incidents encountered by the airline companies or during flight tests undertaken by the manufacturers reveal perturbations of the LOC beam. Particular consequences of such perturbations for the automatic flight controls are untimely alarms about excessive deviation, premature engagement in an LOC capture mode, more or less large excursions of the airplane parameters (such as slip angle or yaw rate, close to the ground if the automatic pilot is already in LOC beam holding mode, or an excursion in lateral trajectory if the airplane is in automatic phase of rolling on the ground. 
     These perturbations can occur in very diverse and sometimes unpredictable situations. For example, these perturbations can occur during an undetected and uncorrected breakdown of the LOC receiver, during an undetected breakdown of an LOC transmitter or during perturbations of the transmitter, especially when an airplane flies over the LOC transmitter on takeoff or stops in front of the LOC transmitter on the runway. In all of these cases, signal reflections cause perturbations while the airplane in automatic landing mode can be at low altitude or on the ground. 
     The problem is accentuated by the diversity of noise profiles (multiple frequency) and by the fact that the airline companies are tending to generalize the use of automatic landing, even in good visibility. 
     SUMMARY OF THE INVENTION 
     A need therefore exists for detecting the perturbations of the LOC signal, especially when the airplane is on the ground. The invention permits at least one of the problems explained hereinabove to be resolved. 
     Thus the invention has as an object a method in an aircraft for detecting the presence of noise in a signal of LOC type, this method comprising the following steps:
         estimation of a first lateral speed of the said aircraft according to a first set of parameters;   estimation of at least one second lateral speed of the said aircraft according to at least one second set of parameters, among which at least one parameter is of different nature from each parameter of the said first set of parameters;   comparison of the said first lateral speed and of the said at least one second lateral speed according to a threshold;   detection of the presence of noise in the said signal of LOC type if the difference between the said first lateral speed and the said at least one second lateral speed is greater than the said threshold.       

     The method according to the invention thus makes it possible, by means of data available in the aircraft, easily to reduce the effects related to perturbations due to noise present on a guide signal of LOC type received by an aircraft. 
     According to a particular embodiment, the method additionally comprises a step of validation of the said detection of the presence of noise in the said signal of LOC type, the said validation step being able to be based on at least one parameter of state of the said aircraft. The validation step makes it possible to control the actions to be taken when noise is detected on a guide signal of LOC type and to avoid taking untimely action. 
     Advantageously, the method additionally comprises a step of relative validation of the said estimates of the said first and at least one second lateral speed, in order to estimate the coherence thereof. 
     According to a particular embodiment, the said first set of parameters comprises parameters of inertial type. 
     According to another particular embodiment, the said at least one second set of parameters comprises parameters of guide type determined on the basis of data obtained from a source external to the said aircraft, such as data determined on the basis of a signal of LOC type. 
     According to a particular embodiment, the method additionally comprises a step of deactivation of an automatic control device of the said aircraft to reduce the differences of trajectory of the aircraft. 
     The invention also has as an object a device comprising means suitable for implementation of each of the steps of the method described in the foregoing as well as an aircraft equipped with such a device. 
     The invention also has as an object a computer program comprising instructions suitable for implementation of each of the steps of the method described in the foregoing. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Other advantages, objectives and characteristics of the present invention become clear from the detailed description hereinafter, given by way of non-limitative example, with reference to the attached drawings, wherein: 
         FIG. 1  schematically represents the device according to the invention; 
         FIG. 2  illustrates the notations used for calculation of the inertial estimate of the lateral speed of an airplane; 
         FIG. 3  schematically presents the module for estimation of the lateral speed by means of parameters obtained from guide data; 
         FIG. 4  illustrates the frequency content of the lateral speed estimated according to the module represented in  FIG. 3 ; 
         FIG. 5  schematically represents the module for detecting noise on a signal of LOC type; and 
         FIG. 6  schematically illustrates an example of a module for implementing a corrective action when the presence of noise on a signal of LOC type has been detected. 
     
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     The system according to the invention is installed on board the airplane to detect the presence of noise on the LOC guide signal when the automatic pilot is engaged. If noise is detected on the LOC guide signal, this information is consolidated and then treated to reduce the effects of this noise on the trajectory of the airplane, especially when the airplane is in the phase of rolling on the ground. 
     The method according to the invention preferably comprises the following three phases: estimation of the lateral speed of the airplane according to at least two different calculation modes, detection of noise on the LOC signal and reduction of the effects of the detected noise. The reduction of the detected noise consists, for example, in disconnecting the automatic pilot. Such an action of resumption of guidance of the airplane on the ground by the pilot can be regarded as the most appropriate for reducing the differences of trajectory of the airplane. 
     According to the invention, noise detection is based on a difference of the lateral speeds of the airplane relative to the runway, the lateral speeds being worked up on the basis of two independent data sources. 
       FIG. 1  illustrates the overall diagram of device  100  according to the invention. The elements represented by dashed ellipses are generally already present in the airplanes. As illustrated, device  100  comprises a module  105  for inertial estimation of the lateral speed of the airplane relative to the runway axis, and its inputs are connected to inertial central unit  110  of the airplane (ADIRS, Air Data Inertial Reference System) and to automatic flight control calculator  115  (FMGEC, Flight Management and Guidance Envelope Computer). 
     Device  100  also comprises a module  120  for estimating the lateral speed determined on the basis of guide information of the LOC signal, known as lateral speed ILS. The inputs of module  120  are connected to inertial central unit  110 , to multi-mode receiver  125  (MMR, Multi-Mode Receiver) and to automatic flight control calculator  115 . 
     A module  130  for detecting noise on the LOC guide signal is connected to modules  105  and  120  as well as to modules  110 ,  115  and  125 , while a decision module, such as module  135  for disengagement of the automatic pilot, is connected to module  130  for noise detection, to module  115  and to a calculator  140  for management of information of the landing gear (LGCIU, Landing Gear Control Interface Unit). 
     The functioning of modules  105 ,  120 ,  130  and  135  is described hereinafter. 
     Module  105  for inertial estimation of the lateral speed of the airplane relative to the runway axis, denoted as VY inertial , uses data obtained from calculators generally present in the airplane. The estimate of the inertial lateral speed of the airplane relative to the runway axis is preferably established according to the following relationship
 
 VY   inertial   =k×GND ×( TTRK−QFU   estimate )+ l   1   ×YAW   rate  
 
where GND represents the component of the speed, relative to the ground, of the airplane in the horizontal plane (expressed, for example, in knots, kts);
 
     TTRK represents the true track angle of the airplane, defined by the speed vector of the airplane, in the plane X-Y of the ground (expressed, for example, in radians, rd); 
     YAW rate  represents the yaw rate of the airplane (expressed, for example, in radians per second, rd/s); 
     I 1  is the algebraic distance between the IRS (Inertial Reference System), or in other words the inertial central unit of the airplane, and the LOC antenna (expressed, for example, in meters, m); 
     QFU estimate  represents the estimated value of the QFU of the runway in geographic axes, or in other words the geographic heading of the runway (expressed, for example, in radians, rd); and 
     k is a unit conversion variable for conversion, for example, of knots (kts) to meters per second (m/s). 
     The values of GND, TTRK and YAW rate  are supplied by the central inertial unit. 
     The value QFU estimate  can be calculated recursively according to the following formula: 
     
       
         
           
             
               
                 QFU 
                 estimate 
               
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                 TTRK 
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                         QFU 
                         estimate 
                       
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                           1 
                         
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                     - 
                     
                       TTRK 
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     The algorithm for calculating the value QFU estimate  is advantageously initiated at a radio-altimetric altitude of 400 feet (approximately 122 meters) when the automatic pilot is engaged. 
       FIG. 2  illustrates the notations used for calculating the inertial estimate of the lateral speed of an airplane. When an airplane  200  is approaching runway  205  for a landing, it detects an LOC guide signal of a transmitter  210  if it is in the envelope of transmitted LOC guide signal  215 . The LOC guide signal makes it possible to align the speed vector of the airplane along the LOC axis. The average track of the airplane therefore constitutes a good estimate of the LOC axis and of the runway axis, and is all the more precise when the automatic pilot is engaged. Reference  220  establishes the estimated runway orientation, the variable QFU estimate  representing the angle between this orientation and geographic north. The airplane axis is established by reference  225 , the airplane heading (HDG) representing the angle between this axis and geographic north. This axis can be used to determine angle  230 , generally referred to as HDG-QFU estimate , formed between the airplane axis and the runway axis determined by the value QFU estimate  establishing the runway orientation. Similarly, the axis of displacement of the airplane, determined by the speed vector GND of the airplane and reference  235 , makes it possible to evaluate the angle  240  formed between axis  235  of displacement of the airplane and the runway axis. This angle is equal to the difference between the true track of the airplane (TTRK) and QFU estimate . 
     Concurrently, a second estimate of the lateral speed of the airplane relative to the runway is determined by module  120  on the basis of the LOC signal and of IRS parameters. 
       FIG. 3  schematically represents module  120 . A first calculation is performed in submodule  300  to determine a lateral speed VY IRS  on the basis of the VNS IRS , VE OIRS  and YAW rate  data obtained from central inertial unit  110  and QFU data supplied by automatic flight control calculator  115 . The lateral speed VY IRS  is calculated according to the following equation:
 
 VY   IRS   =VNS   IRS ×sin( QFU )− VEO   IRS ×cos( QFU )+ l   2   ×YAW   rate  
 
where: VNS IRS  represents the component of the ground speed of the airplane along the geographic north-south axis (expressed, for example, in meters per second, m/s);
 
     VEO IRS  represents the component of the ground speed of the airplane along the geographic east-west axis (expressed, for example, in meters per second, m/s); 
     I 2  is the algebraic distance between the IRS, or in other words the central inertial unit of the airplane, and the LOC antenna (expressed, for example, in meters, m), I 1 =I 2 ; and 
     YAW rate  is the yaw rate of the airplane (expressed, for example, in radians per second, rad/s). 
     Simultaneously, the LOC signal representing the difference between the axis of the LOC radio beam and the position of the receiving antenna of the airplane is combined with the signal SENS representing the sensitivity of the LOC radio beam in multiplier  305 . The LOC signal can be expressed, for example, in microamperes, while the SENS signal can be expressed in meters per microampere. The LOC signal is obtained from multi-mode receiver  125 , while the SENS signal is supplied by automatic flight control computer  115 . 
     A low-pass filter  310  having a time constant τ 3  is applied to the signal obtained from multiplier  305 . A gain 1/τ 1  is then applied to the filtered signal in submodule  315  before this is added to the lateral speed VY IRS  in adder  320 . The signal obtained from adder  320  is filtered in a high-pass filter  325  having time constant τ 1 . The output of high-pass filter  325  represents a filtered drift VY IRS  of the lateral position of the LOC receiving antenna of the airplane, corresponding to a first estimate of the lateral speed ILS (VY ILS ). 
     Similarly, a gain 1/τ 2  is applied in submodule  330  to the signal filtered by low-pass filter  310 , before this is added to the signal obtained from high-pass filter  325  in adder  335 . The signal obtained from adder  335  is filtered in a high-pass filter  340  having the time constant τ 2  to form the lateral speed signal VY ILS . 
     The two high-pass filters  325  and  340  associated with the two submodules  315  and  330  are complementary filters, mounted in cascade, for estimating the lateral speed ILS. 
     The values of the time constants τ 1 , τ 2  and τ 3  are preferably optimized such that the lateral speed VY ILS  is representative in frequency of the response of the airplane being guided on the ground by the automatic pilot. 
     The lateral speed VY ILS  calculated in this way is therefore the result of two complementary filters applied to inertial data and to information obtained from the multi-mode receiver. 
     As illustrated in  FIG. 4 , which shows a schematic representation, it should be noted here that, at low frequency, the lateral speed VY ILS  is equivalent to the drift of the LOC signal, whereas, at high frequency, the lateral speed VY ILS  behaves as an inertial lateral speed (τ represents a composite variable related to the time constants τ 1 , τ 2  and τ 3 ). 
     Noise detection module  130  algebraically compares the two values of lateral speed obtained in modules  105  and  120 . The algebraic difference of these lateral speeds is compared to a predetermined threshold, typically a threshold fixed at 1 meter per second. If the difference is larger than or equal to the predetermined threshold, a first condition for detection of noise on the LOC signal is achieved. This condition is preferably validated by a complementary mechanism such as described hereinafter. 
     It should be noted here that this threshold is the result of a compromise between, on the one hand, the need to detect noise levels on the LOC signal that have effects the trajectory of the airplane in automatic rolling phase that may cause it to stray off the runway and, on the other hand, the requirement that the automatic pilot not be disengaged in untimely manner. 
     Advantageously, noise detection is validated only if the phase of approach of the airplane has been achieved with an automatic pilot engaged sufficiently soon (the convergence and precision of the calculation of the QFU of the runway necessitate that the average track of the airplane be close to QFU) and if the data necessary for calculation of the inertial lateral speeds and ILS are valid, or in other words sufficiently precise and mutually coherent. In particular, since the precision of the TTRK parameters becomes poorer at low speed, the detection of LOC noise is preferably inhibited for GND speeds slower than 80 knots, or in other words approximately 150 kilometers per hour. 
       FIG. 5  illustrates the algorithm implemented in noise detection module  130 . A submodule  500  is used to calculate the algebraic difference between the two estimated lateral speeds and to compare this difference with a predetermined threshold. Concurrently, submodule  505  verifies, by means of data obtained from automatic flight control calculator  115 , that the automatic pilot has been engaged for a sufficient time, for example for longer than 40 seconds, before the LANDTRK guide phase, corresponding to a radio-altimetric altitude of lower than 400 feet (approximately 122 meters), is tripped. Similarly, submodule  510  verifies, by means of data obtained from multi-mode receiver  125  and from the central inertial unit of the airplane, that the estimated lateral speeds are coherent. For example, submodule  510  checks the difference of the estimated lateral speeds over time and invalidates their value when the GND speed of the airplane is slower than 80 knots (approximately 150 kilometers per hour). If the three conditions determined by submodules  500 ,  505  and  510  are verified, for example by means of a logical AND  515  on the signals obtained from these submodules, an indication according to which noise is detected on the LOC signal is emitted by submodule  520 . 
     According to a particular embodiment, the indication according to which noise is detected on the LOC signal is used to disengage the automatic pilot. Preferably this disconnection can take place only if the airplane is on the ground in 3-point position, meaning that the main landing gear and the nose landing gear are in compressed position. This condition makes it possible to ensure that the pilot can resume manual control of the airplane in a comfortable configuration. 
       FIG. 6  schematically illustrates the algorithm of module  135  for controlling disengagement of the automatic pilot. As illustrated, a submodule  600  verifies that the action related to detection of noise on the LOC signal can be effected. By means of data obtained from calculator  140  for management of information related to the landing gear, submodule  600  verifies here that the airplane is positioned on the ground and more particularly that the main landing gear and the nose landing gear have been positioned on the ground for longer than one second. If noise was detected on the LOC signal in module  130 , and if the airplane is positioned on the ground, the action related to detection of noise on the LOC signal can be effected. This verification can be achieved by means of logical AND  605 . If the conditions are verified, the action is then effected. In this case a command to disengage the automatic pilot is generated by submodule  610 . 
     The device according to the invention therefore makes it possible to detect noise on the LOC guide beam when the automatic pilot is engaged, so as to reduce the effects during the rolling phase. In addition, the device according to the invention makes it possible to achieve a low disturbance rate, meaning that it does not detect noise unduly, since this would have the consequence of disconnecting the automatic pilot in untimely manner during the rolling phase. Similarly, the device according to the invention makes it possible to obtain a sufficient availability factor, taking into account system conditions that inhibit the noise detector. 
     Although the system for detecting noise on the LOC signal in the manner described hereinabove acts on the automatic pilot by commanding its disconnection, it is possible to use the information on differences of lateral speeds for other purposes, especially to modify the guide orders of the automatic pilot, to display corresponding information or to forewarn the crew by means of a specific alarm. 
     It also should be noted that the inertial lateral speed can be calculated differently. In particular, the estimate of the QFU of the runway can be achieved by a different algorithm or can be obtained from another source of the airplane. 
     Similarly, the lateral speed ILS can be calculated differently. In particular, the information originating from central inertial units may not be used. 
     It is also possible to implement a different logic for detecting noise on the LOC signal. In particular, the threshold value used to compare the lateral speeds can be variable and can depend on parameters related to the airplane. 
     Similarly, the logic for disengagement of the automatic pilot can be different. In particular, the duration of confirmation of the information that the airplane is on the ground in 3-point position may depend on parameters related to the airplane instead of being fixed. Other conditions may be added, especially the speed of the wheels of the landing gears. 
     Naturally a person competent in the field of the invention will be able to apply modifications in the foregoing description in order to satisfy specific needs.