Patent Publication Number: US-7225061-B2

Title: Method and device for detecting pilot induced oscillations in an aircraft

Description:
The present invention relates to a method and a device for detecting, in an electric command for flying an aircraft, oscillations corresponding to pilot induced oscillations. 
     It is known that the controls (sticks, rudder) of modern aircraft, particularly those known as ministicks, are easy for the pilot and/or the copilot of said aircraft to handle, it being possible for them to be tilted very quickly. By contrast, the actuators operating the moving aerodynamic surfaces of the aircraft (ailerons, flaps, rudders, etc.) operated from these controls cannot respond instantly to the electric commands generated by these controls. There may therefore, at large control amplitudes, be a significant phase shift between the movement of a control and the movement of the moving aerodynamic surfaces it controls. 
     As a result, the pilot, noticing that the position of the operated mobile aerodynamic surfaces is lagging behind the position chosen for said control member, may attempt to further increase the amplitude by which he tilts said control member. However, the amplitude of turning of the aerodynamic surfaces may then exceed that corresponding to the initial command, which means that the pilot then reduces the tilt of said control member, thus leading, with a delay, to a return of the aerodynamic surfaces, etc. Oscillations, caused by coupling and generally referred to in aeronautical parlance as “pilot induced oscillations” (PIO) therefore occur in the aircraft and may degrade the precision of the flying. 
     In an attempt at solving this problem, it is known practice to increase the size of the actuators of the aerodynamic surfaces controlled and of their electrical and hydraulic supplies, thus increasing the cost and mass of the aircraft. Such increases in cost and in mass may become intolerable in the case of large-sized aircraft. 
     Furthermore, document U.S. Pat. No. 4,298,833 envisions the processing of the command, particularly by filtering it, to convert it into an order that is free of pilot induced oscillations. However, this processing is performed without any true detection of the pilot induced oscillations and without the knowledge of the pilot, who may then continue to generate such oscillations. 
     The object of the present invention is to overcome these disadvantages. The invention relates to a simple method and a simple device for detecting pilot induced oscillations, so as to alert the pilot to their occurrence and/or take action to avoid their effect. 
    
    
     To this end, according to the invention, the method for detecting oscillations, corresponding to pilot induced oscillations, in an electric command δ controlling an aircraft about one of its axes of maneuver (roll, pitch or yaw), said electric command δ being able to adopt a maximum value δo, is notable in that:
         said electric command δ is sampled at a sampling interval Δt to obtain a plurality of N successive samples x k  with k=0, 1, 2, . . . N−1;   the number of samples N and the sampling interval Δt are chosen so that the inverse of their product N.Δt is at least approximately equal to 0.3 Hz;   said sampled command is broken down into a Fourier series in such a way that each sample x k  adopts the form:       

     
       
         
           
             
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              in which expression j is the symbol assigned to the imaginary part of a complex number, r is the number of harmonic components in the Fourier series and A r  are the Fourier coefficients, each of the latter being equal to: 
           
         
       
    
     
       
         
           
             
               A 
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             the moduli |A1| and |A2| of the first two Fourier coefficients A1 and A2, respectively corresponding to r=1 and r=2, are determined; 
             an index I is calculated that is equal to the quotient of the sum |A1|+|A2| of said moduli by the number N of samples; and 
             said index I is compared with a fraction q.δo of said maximum value δo of said electric command δ and it is reckoned:
           that there are no pilot induced oscillations if said index I is lower than said fraction q.δo; and   that there are pilot induced oscillations if said index I is equal to or greater than said fraction q.δo, the value, lying between 0 and 1, of said index I being representative of the amplitude of said oscillations corresponding to said pilot induced oscillations.   
         
           
         
       
    
     The method of the present invention is based on two observations made by the Applicant Company when examining real earlier flight scenarios in which pilot induced oscillations had occurred:
         on the one hand, this examination showed that induced oscillations have frequencies of between 0.3 Hz and 0.6 Hz, which means that the purpose of breaking down the command into a Fourier series is to determine the components of this command that lie between 0.3 Hz and 0.6 Hz;   on the other hand, said examination showed that, when said induced oscillations occur, their energy exceeds a certain fraction, generally at least approximately equal to 15%, of the maximum vale δo of the command δ. Hence, according to a particular feature of the present invention, said coefficient q of said fraction q.δo is determined from a pre-established database grouping together a number of real earlier flight scenarios in which such pilot induced oscillations occurred.       

     As a preference, said electric command δ is filtered prior to sampling, in order to eliminate from it any oscillations other than those corresponding to pilot induced oscillations. 
     When the aircraft is designed to be able to be flown by a pilot and a copilot, it is advantageous for said electric command δ to be the sum of the corresponding individual electric commands resulting respectively from the action of the pilot and from the action of the copilot. 
     The present invention also relates to a device for implementing the abovementioned method. Such a device, allowing the detection of oscillations corresponding to pilot induced oscillations, in an electric command δ controlling an aircraft about one of its axes of maneuver, is notable in that it comprises:
         first means for sampling said electric command δ;   second means for breaking said sampled command δ down into a Fourier series;   third means for calculating said index I; and   fourth means making the comparison between said index I and said fraction q.δo of said maximum value of said electric command δ.       

     As a preference, said first, second, third and fourth means consist of one and the same computer which, advantageously, may form part of the fly-by-wire computers of said aircraft. 
     The figures of the attached drawing will make it easy to understand how the invention can be embodied. In these figures, identical references denote elements which are similar. 
       FIGS. 1 to 3  show block diagrams of three schematic embodiments of the invention. 
     The control system, depicted schematically in  FIG. 1 , comprises a control  1 , for example of the stick type, which can tilt about an articulation  2  in two opposite directions (double-headed arrow  3 ). The control  1  is associated with a transducer  4  delivering, at its output  5 , an electric command δ representative of the amplitude of the tiling of the control  1 . 
     The electric command δ is transmitted to fly-by-wire computers  6  which generate an order for the actuators  7 , able to turn moving aerodynamic surfaces  8  of an aircraft (not depicted) about their axis of rotation  9 . If the control  1  is a stick, for example a ministick, the moving aerodynamic surfaces  8  control said aircraft either in terms of roll or in terms of pitch. 
     Furthermore, the command δ, in which frequencies lying between 0.3 and 0.6 Hz, representative of any possible pilot induced oscillation induced by the pilot operating said control  1 , are sought, is sent to a filtering system  10  able to eliminate from this command δ any oscillations that do not correspond to said induced oscillations. Thereafter, the filtered command δf is sent to a computer  11 , able to break it down into a Fourier series, so as to detect therein any components having frequencies lying between 0.3 and 0.6 Hz. 
     To allow such a breakdown into a Fourier series, it is absolutely essential for said command δ to be stored temporarily. For this, the computer  11  begins by sampling the command δ at a sampling interval Δt to obtain N successive samples x k  with k=0, 1, 2, . . . N−1. Thus, the continuous command δ is converted into a discrete signal x k . 
     The number N and the interval Δt are chosen, among other technical considerations, so that the inverse of their product is at least practically equal to 0.3 Hz. For example, in practice, Δt may be taken as 125 ms with N=27. Thus, the product N.Δt is equal to 3.375 s and its inverse 1/N.Δt is equal to 0.296 Hz. 
     The discrete signal x k  is then broken down into a Fourier series by the computer  11  so that, regardless of the value of k between 0 and N−1, the following hold true, in a known way: 
     
       
         
           
             
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     In these expressions:
         j is the symbol (equal to √{square root over (−1)}) assigned to the imaginary part of a complex number;   r is the number of harmonic components in the Fourier series; and   Ar are the Fourier coefficients of said breakdown.       

     Thus, the coefficients Ar constitute a periodic signal which is periodic in terms of frequency, with a period 1/Δt, sampled at the frequency Δf=1/N.Δt. 
     Such a breakdown into a Fourier series therefore provides the frequency-based analysis of the signal x k  on a horizon N.Δt with a resolution Δf=1/N.Δt. 
     Furthermore, in a known way, the moduli |Ar| of the Fourier coefficients Ar, for which r is other than 0, measure the energy of the signal X k  at the frequency r.Δf. 
     Now, in this instance, Δf is close to 0.3 Hz, which means that r.Δf is also close to 0.3 Hz for r=1, corresponding to the first Fourier coefficient A1, and close to 0.6 Hz for r=2, corresponding to the second Fourier coefficient A2. 
     The sum |A1|+|A2| of the moduli |A1| and |A2| of said first and second Fourier coefficients A1 and A2 therefore represents the energy of the signal x k  for frequencies lying between 0.3 Hz and 0.6 Hz and corresponding to the pilot induced oscillations. 
     The computer  11  therefore calculates said sum of the moduli |A1|+|A2|, which is consistent with the amplitude of the command δ give or take the factor N. 
     In consequence, in order to determine whether there is any pilot induced oscillation in the command δ, the energy of the looked-for sinusoidal signal, that is to say the sum |A1|+|A2|, can be compared with the maximum value δo of said command δ. 
     To do this, the computer  11  calculates an index I equal to the quotient (|A1|+|A2|)/N and compares this index with a predetermined fraction q.δo of this maximum value δo. 
     The coefficient q of this fraction q.δo is, for example, determined from a database grouping together a number of earlier known flight scenarios in which pilot induced oscillations occurred onboard aircraft. This coefficient is generally close to 0.15. 
     Hence, the computer  11  is able to generate a signal S which is:
         equal to 0 when said index I is lower than q.δo, because then there is no pilot induced oscillation; and   equal to (|A1|+|A2|)/N, lying between 0 and 1, when said index I is equal to or greater than q.δo and there is then some pilot induced oscillation.       

     As mentioned hereinabove, the index I is consistent with the amplitude of the command δ, in the frequency range 0.3 Hz–0.6 Hz. This amplitude has a maximum value δo when the command contains only frequencies lying within that range. Thus, according to the invention, pilot induced oscillation is deemed to have been detected if the amplitude of the command δ, in said frequency range 0.3 Hz–0.6 Hz, exceeds q % of the maximum amplitude δo. 
     The signal S may be transmitted to a display device  12 , for example provided with a pointer  13  associated with a scale  14 , the relative position of said pointer with respect to said scale being representative of the amplitude of said signal S and therefore of the existence and amplitude of any such induced oscillations corresponding to pilot induced oscillations. 
     Furthermore, said signal S is available on a line  15 , for any worthwhile use. 
     In order to break the discrete signal x k  down into a Fourier series, the computer  11  may employ algorithms known by the name of “Fast Fourier Transforms” or any other known method, such as a recurrent method for example. 
       FIG. 2  schematically illustrates the case where the control system comprises two controls  1 . 1  and  1 . 2  in parallel, respectively similar to control  1  and intended for a pilot and for a copilot. In this case, the respective transducers  4 . 1  and  4 . 2  (similar to the transducer  4 ) are able to deliver, on their output  5 . 1  or  5 . 2 , individual commands  61  and  62  originating respectively from the action of the pilot and that of the copilot. These individual commands  61  and  62  are sent to the fly-by-wire computers  6  which process them to control the moving aerodynamic surfaces  8 . 
     Furthermore, according to the present invention, the individual commands δ 1  and δ 2  are sent to filters  10 . 1  and  10 . 2  (similar to the filter  10 ) and summed in a summer  16 , before being transmitted to the computer  11 . In this case, the filtered command δf sent to the computer  11  for similar processing to the processing described hereinabove, is therefore the filtered sum of the two individual commands δ 1  and δ 2 . 
     In the control system depicted schematically in  FIG. 3 , the control  1 . 3  tilts not only in the two opposite directions of the arrow  3 , but also in two opposite directions (see the double-headed arrow  17 ) directed at right angles to the first ones, which means that the associated transducer  4 . 3  can deliver two electric commands δA and δB. One of these may, for example, be intended for control in roll and the other for control in pitch. The fly-by-wire computers  6  which receive the commands δA and δB generate respective orders for actuators  7 A and  7 B, tasked with actuating moving aerodynamic surfaces  8 A and  8 B which themselves allow the aircraft to be flown about the corresponding axis of maneuver (roll axis and pitch axis, for example). 
     Furthermore, according to the present invention, each command δA and δB is processed in a similar way to the signal δ of the system of  FIG. 1 , that is to say that it is filtered by a filter  10 A or  10 B (similar to the filter  10 ) then processed by a computer  11 A or  11 B (similar to the computer  11 ), the result of said processing being displayed on a display device  12 A or  12 B (similar to the display device  12 ) comprising a pointer  13 A or  13 B and a scale  14 A or  14 B and/or available on lines  15 A or  15 B (similar to said line  15 ). 
     It will be readily understood that it may be advantageous for the computers  11 ,  11 A and  11 B to form part of the fly-by-wire computers  6  even though, in the drawings, they are depicted as distinct from the latter. 
     Furthermore, by association of  FIGS. 2 and 3 , it is easy to construct an embodiment provided with two controls, each of which is able to fly the aircraft about two axes of maneuver.