Abstract:
A backup engines-only flight control system accomplishes improved aircraft banking in response to commands by the use of a lead compensated transient differential thrust servo parameter. The lead compensated transient differential thrust servo parameter compensates for the ordinary sluggish banking response during engines-only flight control. Accordingly, because an aircraft equipped with the present invention will respond in a manner expected by pilots, safer backup engines-only flight control is achieved during emergency situations. The improvement comprises a means and method of detecting the magnitude of direction of commanded bank angle and generating a corresponding lead compensated transient differential thrust servo parameter.

Description:
FIELD OF THE INVENTION 
     The present invention relates generally to the field of flight control systems for airplanes. More particularly, to a flight control system that uses engine thrust for backup flight control of a multi-engine airplane. Still more particularly, to a flight control system that eliminates sluggish aircraft banking response. Even more particularly, to the use of lead controllers to eliminate sluggish aircraft banking response. 
     BACKGROUND 
     Standard primary flight control systems are generally reliable. Aircraft designers have successfully integrated complex mechanical and electrically controlled flight control components into these standard primary flight control systems. Such systems rely upon aircraft flight surfaces e.g. rudders, flaps ailerons, to control the heading and pitch of an aircraft. 
     Many standard primary flight control systems have redundant components as part of the standard primary flight control backup system. In the rare event that system fails however, a backup flight control system should be used that is not integral to the primary flight control system. One such flight control system is based entirely on control of the engine thrust. 
     For multi-engine aircraft, one previously disclosed method of controlling the aircraft uses engine thrust to control the aircraft pitch angle and roll. Pitch angle is controlled by concurrent equivalent thrust adjustments upon laterally positioned engines. Aircraft banking is controlled by differential thrust adjustments upon laterally positioned engines. 
     The use of differential engine thrust to control aircraft banking is however inexact and leads to difficulty in flying and hazardous landings. Several factors contribute to the control problems; the pilot&#39;s inexperience with this method of aircraft control, the less exact control method augmented by a sluggishly responding differential engine control, and the stress of an emergency situation. 
     One example of related art disclosed in the art of engines-only flight control uses static control parameters to generate servo command parameters indicative of the desired flightpath. For instance U.S. Pat. No. 5,330,131 issued to Burcham et al. (“Burcham”) discloses an engines-only flight control system based on differential engine thrust. The disclosure in the Burcham patent discloses a pilot controlled input device consisting of either a control stick, thumbwheel or radio frequency receiver. The input device generates command parameters that are then modulated by scaling circuits and amplified by static gains. The outputs from the static gain amplifiers are then applied to the left and right engine servos through a summation function. The Burcham invention therefore discloses a successful system of engines only flight control. However, it is the static control parameters that lead to sluggish and unfamiliar aircraft banking response. 
     Another example of remotely related art is disclosed in U.S. Pat. No. 5,551,402 issued to Dahl (“Dahl”). Dahl&#39;s system discloses the use of wireless transmitters to control the flight control surfaces by the use of receivers positioned within, or on, the engines. However, Dahl&#39;s system of backup flight control depends upon the use of wireless transmitters and receivers, rather than differential engine thrust, to replace the primary engine flight control system. 
     Therefore, in the event that the standard primary flight control components fail to respond to a pilot&#39;s stick commands, a backup engines-only flight control system should be used that does not utilize the standard flight control components, but still permits banking and pitch adjustments. Moreover, this backup engine-only-flight-control-system should preferably respond in a manner that is substantially similar to what the pilot would expect from the primary flight control system. The present invention accomplishes these aims by the incorporation of lead compensation to backup engines-only flight control systems by use of a lead controller. Therefore, the present invention improves the responsiveness and predictability of engines-only flight control, thereby improving the usability, safety and reliability of engines-only flight control. 
     SUMMARY OF THE INVENTION 
     It is an object of the present invention to improve the banking response in engines-only flight control. It is another object of the invention to use at least one lead controller to eliminate sluggish banking response during engine-only flight control. It is another object of the invention to improve banking response by modifying existing engine only backup flight control systems based on lateral-differential engine thrust. It is another object of the invention to provide engine-only flight control systems with responsiveness that is substantially comparable to the banking response expected from stick controlled primary flight control systems. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The novel features characteristic of the invention are set forth with particularity in the appended claims. The invention itself, both as to its structure and its operation together with the additional object and advantages thereof, will best be understood from the following description of the preferred embodiment of the present invention when read in conjunction with the accompanying drawings wherein: 
     FIG. 1. A depiction of the preferred embodiment of the present invention. 
     FIG. 2. A depiction of the preferred embodiment in greater detail. 
    
    
     DESCRIPTION OF PREFERRED EMBODIMENTS 
     While these descriptions directly describe the above embodiments, it is understood that those skilled in the art may conceive modifications and/or variations to the specific embodiments shown and described herein. Any such modifications or variations that fall within the purview of this description are intended to be included therein as well. It is understood that the description herein is intended to be illustrative only and is not intended to be limiting. Rather, the scope of the invention described herein is limited only by the claims appended hereto. 
     Generally, a multi-engine aircraft will have at least a drive-engine-pair. Each engine of this engine-pair will be positioned equidistant from the fuselage on each wing of the aircraft. During normal flight conditions, the engine-pair provides some or all propulsion to the aircraft. A primary flight control system adjusts the bank and pitch of the aircraft. 
     The primary flight control system generally comprises a pilot&#39;s stick or autopilot output that inputs commands to several standard flight control components, e.g., elevators, ailerons, and rudders. It is also common to provide one or more backup flight control systems. One method of backup flight control uses only the engines to control the pitch and bank angle of the aircraft. This type of backup flight control system will hereinafter be referred to as an engines-only-flight-control system. The present invention, a Lead Controller  1 , is useful as an improvement to engines-only-flight-control systems. 
     A multi-engine aircraft engaged in engines-only-flight-control makes use of main-drive engine-thrust-magnitudes to direct both the pitch and the banking of the aircraft. Thrust magnitudes applied concurrently and equivalently to each laterally positioned main drive engine servo controls the aircraft pitch. Commonly, a proportional plus derivative controller controls the pitch of an aircraft during engine-only-flight, which probably provides the most reliable means of controlling the pitch of the aircraft during engines-only flight. Thus, the primary use of lead compensation is contemplated to be in banking control of engines-only-flight control systems. The use of lead compensation in aircraft pitch angle changes would not be carried out in much the same way as described below for bank angle changes because only a classical control lead would be required. 
     With reference to FIG. 1, the current state of the art of engines-only-flight-control utilizes an input device  2  to initiate aircraft banking changes. The input device  2  is used to issue commands to a prior art engines-only-flight-control system  3  and the system  3  in turn issues banking commands to left and right engine servos  56  and  57 . The banking commands issued to the servos  56  and  57  are statically proportional to the lateral displacement of the input device  2 . The input device  2  disclosed in the prior art is commonly a device that has discrete directional heading selection capability. 
     Generally, in current engines-only flight control systems, the lateral position of the input device  2  generates at least one input parameter  22 . In these prior art systems, the application of the input parameter  22  to the prior art engines-only flight control system  3  results in a static servo parameter value  32 . The static servo parameter value  32  has a static magnitude depending directly upon the lateral position of the input device  2 , and polarity corresponding to right or left banking. 
     The magnitude and polarity of the static servo parameter  32  when the input device  2  is at dead center would have zero magnitude and therefore zero polarity. Accordingly, no banking would result. Ordinarily, the initial or default value of the static servo parameter  32  has a magnitude of zero corresponding to a null lateral thrust differential. A previously known method in the art of engines-only flight control to accomplish the null lateral thrust differential uses roll rate and roll amplitude gyros. 
     When the input device  2  is displaced to the right or left, a nonzero static servo parameter value  32  having a magnitude proportional to the lateral displacement of the input device  2  results. Moreover, the static servo parameter value  32  would have positive or negative polarity depending upon the direction of the input device  2  displacement. The parameter value  32  is then applied to the engine servos  56  and  57  causing a differential lateral thrust to be applied to the engine pair causing aircraft banking. The actual direction of banking would depend on the polarity of the engine servo parameter value  32 . 
     As mentioned previously, statically proportional thrust differentials applied to the engine servos  56 ,  57  result in sluggish aircraft banking response. The addition of the Lead Controller  1  to the engines only flight control system improves the responsiveness of the aircraft to bank commands by the application of a lead compensated transient differential thrust servo parameter  65 . 
     As depicted in FIG. 1, the preferred embodiment of the Lead Controller  1  uses an input device  2  and the static servo parameter  32  from prior art engines only flight control systems  3 . Moreover, the Lead Controller  1  further comprises a banking polarity and magnitude detector  50 , a washout-accumulator  40 , and a first summation function  60 . 
     During a typical banking command from the input device  2 , the banking polarity and magnitude detector  50  and washout-accumulator  40  generate a lead compensation parameter  51 . The lead compensation servo parameter  51  is applied to the first summation function  60  with the statically proportional servo parameter  32  from the prior art engines only flight control system  3 . The resultant output from the summation function  60  is the lead compensated transient lateral thrust differential servo parameter  65  that is applied to the engine thrust servos  56  and  57  and as a result, improves the banking response of the aircraft. The Lead Controller  1  can also be equipped with supplementary functions that constrain how much, and if the Controller  1  compensation parameter  51  will modulate the static servo parameter value  32 . 
     Referring to FIG. 1, the banking polarity and magnitude detector  50  of the Lead Controller  1  generates a banking polarity and magnitude output parameter  54  proportional to the direction and magnitude of the rate and direction of change of the input device  2  displacement. The preferred embodiment of the Lead Controller  1  generates the banking magnitude and polarity output parameter  31  by computing the difference between subsequent lateral positions of the input device  2 . 
     Subsequent lateral positions of the input device  2  are compared by applying the input parameter  22  to a delay function  71 , multiplying the output with an amplifier having gain of minus one (−1)  53 , and applying the result to a second summation function  52 . The input parameter  22  not subject to a delay comprises the second input to the second summation function  52 . The resultant parameter is the banking polarity and magnitude output parameter  54 . Subsequently, this parameter  54  is coupled to the washout-accumulator  40  to affect the desired characteristic compensation. 
     A more detailed depiction of the preferred embodiment of the Lead Controller  1  is depicted in FIG.  2 . In this depiction one or more logic gates or circuit functions represent each component discussed above. 
     In FIG. 2, a discrete unit delay  501 , an angle difference calculator  504 , and an angle correction switch  505  represent the circuitry comprising the banking polarity and magnitude detector  50  of the depiction of FIG.  1 . Additional functions in FIG. 2 required to control the components representing the banking polarity and magnitude detector  50  is the angle polarity comparators  506 , an angle adjustment switch  507 , and a second summation function  508 . 
     In the depiction of FIG. 2, a input device  2  (e.g., a conventional knob) may be rotated through three hundred and sixty degrees (360°) generates the input parameter  22 . The three hundred and sixty degrees of rotation is suitably quantified into discrete segments,  10  each representing, for example, one degree of angular displacement. Moreover, a separate turn direction calculation determines if the angular displacement represents a left bank turn or a right bank turn. 
     The input parameter  22  of the input device  2  is coupled to a first input  504 . 1  of the heading difference calculator  504  and also to the input  501 . 1  of the discrete unit delay  501 , respectively. The output  501 . 2  of the discrete unit delay  501  is coupled to a second input  504 . 2  of the angle difference calculator  504 . 
     The output  504 . 3  of the angle difference calculator  504  represents the difference between subsequent input parameter  22  values and is coupled to components representing the washout-accumulator  40  through a limiter  310  and an angle correction switch  505 . The angle correction switch  505  is comprised of a logic input  505 . 1 , a adjustment angle input  505 . 2 , an unadjusted angle input  505 . 3 , and an output  505 . 4 . 
     The angle correction switch  505  represents the ability to correct for angle errors generated when the input device  2  is rotated past plus or minus one hundred eighty degrees (±180°) between subsequent sampling periods during a “cold” knob condition. Cold knob refers to a backup flight control condition whereby a pilot first rotates the input device  2  to select a desired heading, and then engages the backup flight control system commanding the aircraft to chase the desired heading. 
     Alternatively, a “hot” knob condition refers to a backup flight condition whereby the aircraft immediately responds to input commands from the input device  2 . During a hot knob condition, the aircraft is never subject to a input device  2  rotation past plus or minus one hundred eighty degrees (±180°) between subsequent sampling periods and correction of the angle difference calculator output  504 . 3  is unnecessary. Therefore, during hot knob conditions, the angle correction switch  505  is enabled only to gate the unadjusted output of the angle difference calculator  504 . 
     During cold knob conditions, if the input device  2  is rotated less than plus or minus one hundred and eighty degrees, angle correction is unnecessary. Under such conditions, the logic input  505 . 1  will be logic low and the angle correction switch  505  will pass the angle difference values calculated by the angle difference calculator  504  coupled to the uncorrected angle input  505 . 3 . On the other hand, angular rotations of the input device  2  greater than plus or minus one hundred and eighty degrees (±180°) require angle correction. 
     Angle correction is accomplished via that angle correction switch  505  which is controlled via an AND gate  511 . Said AND gate  511  has one input, coupled to an independent cold knob input  512  that is asserted during cold knob conditions. During hot knob conditions the cold knob input  512  is logic low. The generation of the logic level cold knob input  512  is considered beyond the scope of the present invention but could be achieved by any manner ordinary in the art. 
     The other input to the AND gate  511  is coupled to the output of an Exclusive-OR gate  509 . The logic state of the output of the Exclusive-OR gate  509  is dependent upon the state of an independent input, Counter-Clockwise Turn  510 , and the output of the polarity comparator  506 . For the purposes of the embodiment in FIG. 2, the independent input, Counter-Clockwise Turn  510 , is asserted when the input device  2  is rotated to the left to command a counter-clockwise turn. Generation of the independent input, Counter-Clockwise Turn  510 , is also considered beyond the scope of the present invention. However, ordinary means in the art for developing this input are contemplated. 
     During a cold knob operating condition, in the event that the input device  2  is rotated past plus or minus one hundred and eighty degrees (±180°) between sampling periods, the angle difference calculator  504  will report an angle difference of less than plus or minus one hundred eighty degrees (±180°). More specifically, rotations to the right past positive one hundred eighty degrees (+180°) will result in negative reported angles and rotations to the left past negative one hundred eighty degrees (−180°) will result in positive reported angles. To provide angle correction in the correct direction and quantity, three hundred sixty degrees (360°) is added to clockwise rotations to the right past positive one hundred eighty degrees (+180°). Likewise, three hundred sixty degrees (360°) is subtracted from counter-clockwise rotations to the left past negative one hundred eighty degrees (−180°). 
     Angle correction of the angle reported from the angle difference calculator  504  is accomplished with the angle adjustment switch  507 , the polarity comparator  506 , the angle summation function  508 , and the angle correction switch  505 . As previously mentioned, if the input device  2  is rotated less than one hundred and eighty degrees in either direction, angle adjustment is unnecessary. Under such conditions, the angle correction switch merely passes the calculated angle differences as reported by the angle difference calculator  504 . This condition is further characterized by a logic low condition on the logic input  505 . 1  of the angle correction switch  505 . However, if the input device  2  is rotated past one hundred and eighty degrees (180°) in either direction, angle adjustment is necessary. 
     To correct errors reported by the angle difference calculator  504 , the angle adjustment switch  507  and the angle summation function  508  add an angle adjustment value  507 . 5  depending upon the output of the polarity calculator  506  and the independent input, Counter-Clockwise Turn  510 . The output  504 . 3  of the angle difference calculator  504  is coupled to one input  506 . 2  of the polarity comparator  506 . The other input  506 . 1  of the polarity comparator  506  is coupled to a reference angle representing zero degrees (0°). The polarity comparator output  506 . 3 , and accordingly, the logic input  507 . 1  of the angle adjustment switch  507  will be logic high whenever the input  506 . 2  of the polarity comparator  506  detects on the output  504 . 3  of the angle difference calculator  504  an angle difference less than zero. 
     Conversely, the logic comparator output  506 . 3 , and therefore the logic input  507 . 1  of the angle adjustment switch  507  will be logic low whenever the input  506 . 2  of polarity comparator  506  detects on the output  504 . 3  of the angle difference calculator  504  an angle difference greater than zero. Moreover, only when the input device  2  is rotated to the left will the Counter-Clockwise Turn  510  input will be logic high. 
     Therefore, if the input device  2  is rotated to the right past positive one hundred eighty degrees (+180°), the angle difference calculator  504  will report a negative angle and the polarity comparator  506  will generate a logic high on the output  506 . 3 . In this condition the Counter-Clockwise Turn  510  input will be logic low. Conversely, if the input device  2  is rotated to the left past negative one hundred eighty degrees (−180°), the angle difference calculator  504  will report a positive angle and the polarity comparartor  506  will generate a logic low on the output  506 . 3 . In this condition the Counter-Clockwise Turn  510  input will be logic high. Only under these two conditions will the Exclusive-OR gate  509 . 3  assert a logic high on the logic input  505 . 1  of the angle correction switch  505  thereby enabling angle correction. 
     The output  507 . 4  of the angle adjustment switch  507  is coupled to the first input  508 . 1  of the angle summation function  508 . The output  504 . 3  of the angle difference calculator  504  is connected to the second input  508 . 2  of the angle summation function  508 . The output  508 . 3  of the summation function  508  is coupled to the second input  505 . 2  of the angle correction switch  505 . If the logic input  505 . 1  of the angle correction switch  505  is logic high, the angle correction gate output  505 . 4  will equal the output of the angle summation function  508 . 3  that represents the corrected angle from the angle difference calculator  504 . However if the logic input  505 . 1  of the angle correction switch  505  is logic low, the angle correction gate output  505 . 4  will equal the output  504 . 1  of the angle difference calculator  504 . 
     The limiter  310  couples the output  505 . 4  of the angle correction gate switch  505 . 4  to a third summation function  415 . The limiter  310  has an input  310 . 1  and an output  310 . 2 . The input of the limiter  310  is coupled to the output  505 . 4  of the angle correction gate  505  and the output  310 . 2  of the limiter  310  is coupled to a third summation block  415 . The limiter  310  restricts the output magnitude of the angle correction gate  505 . In the preferred embodiment, the actual restricted magnitude would depend upon the aircraft and would be based on safety concerns. 
     A accumulator summation function  415 , an accumulator enable gate  402  and a second unit delay  403  comprise the recursion system of accumulating angle difference calculator  504  output values used by the preferred embodiment of the Lead Controller  1 . The accumulator summation function  415  has a first input  415 . 1 , a second input  415 . 2 , and an output  415 . 3 . The first input  415 . 1  of the accumulator summation function  415  is coupled to the output  310 . 2  of the limiter  310 . The second input  415 . 2  of the accumulator summation function  415  is coupled to the third input  402 . 3  of the accumulator enable gate  402  and the output  415 . 3  of the accumulator summation function  415  is coupled to the second input  402 . 2  of the accumulator enable gate  402 . The output  402 . 4  of the accumulator enable gate  402  is coupled to the input  403 . 1  of the second unit delay  403 . The output  403 . 2  of the second unit delay  403  is coupled to the second input  415 . 2  of the accumulator summation function  415  and the third input  402 . 3  of the accumulator enable gate  402 . 
     The first input  402 . 1  of the accumulator enable gate  402  is coupled to an output  404 . 4  of an AND gate  404 . The AND gate  404  operates to enable the accumulator enable gate  402  to begin the accumulation of angular displacement units. The AND gate  404  has a first  404 . 1  and second  404 . 2  input. The first input  404 . 1  of the AND gate  404  is coupled to the output  405 . 3  of a logic comparator  405 . The second input of the logic comparator  405 . 2  is coupled to an angle reference parameter value  406  corresponding to an angular difference less than a single angular displacement unit  11 . 
     The first input  405 . 1  of the logic comparator  405  is coupled to the output  407 . 2  of an absolute value function  407 . The first input  407 . 2  of the absolute value function  407  is coupled to the output  504 . 3  of the angle difference calculator  504 . 
     The absolute value function  407  of the lead controller  1  is used to signify to the second logic comparator  405  that the angle difference calculator  504  is greater than zero and that the input device  2  has been displaced and changed subsequent values of the input parameter  22 . If the output  504 . 3  of the angle difference calculator  504 , and therefore the first input  405 . 1 , of the logic comparator is greater than the reference value  405  on the second input  405 . 2  of the logic comparator  405 , the output  405 . 3  of the logic comparator  405  will result in a logic high at the first input  404 . 1  of the AND gate  404 . If the other input of the AND gate  404  are also at a logic high level, the accumulator summation function  415 , the accumulator enable gate  402 , and the second unit delay  403  are enabled to accumulate the discrete angle difference calculator  504  output values. The accumulation of angle difference calculator  504  output values from the angle correction switch  505  is coupled to the components representing the washout-filter  40  of the lead controller  1  of the present invention. 
     Input  404 . 2  prevents the banking lead compensation from being applied in the direction of desired bank when the aircraft is up against a bank angle limit in that direction. Input  404 . 2  is developed by a two input OR gate  408 , with two inputs,  408 . 1  and  408 . 2  respectively. The first input  408 . 1  of the OR gate  408  is coupled to the output of a shift register  409 . The second input  408 . 2  of the OR gate  408  is coupled directly to the output of a NOR gate  410 . The first and second inputs of the NOR gate  410  are coupled to left and right bank angle limit inputs generated in the ordinary manner of the art. If the aircraft bank angle limit inputs detect that the aircraft is approaching an unsafe left or right bank angle, the corresponding input  408 . 1  or  408 . 2  of the NOR gate  410  will be asserted and the accumulation of bank angle displacement units disabled. 
     Referring again to FIG. 1, the washout-filter  40  comprises a transfer function  41  and operates on the accumulated banking polarity and magnitude output parameter values  54  for the desired characteristic output. Preferably, the characteristic response of the transfer function  40  of the lead controller  1  will be the impulse response of a lead network generally of the form;            Z        (   s   )       =     K   ·         d                   s   2       +     s                 g           c                   s   2       +     b                 s     +   a           ,       where                 s     =     j   ·   2   ·   π   ·   f       ,       for              -   ∞     &lt;   f   &lt;   ∞     ,   and                          
     K, a, b, c, d, and g, represent real numbers. 
     More specifically, the preferred transfer function  40  is generally,              Z        (   s   )       =     K   ·       s                 g         b                 s     +   a           ,       where                 s     =     j   ·   2   ·   π   ·   f       ,     for   -   ∞              &lt;   f   &lt;   ∞     ,   and                            
     K, a, b, and g, represent real numbers. 
     In the depiction of FIG. 2, a discrete washout-filter  400  samples the accumulated calculated angle difference values sampled at a rate  402 , ƒ s =1/T. The discrete transfer function  401  used in this depiction of the preferred embodiment is;            Z        (   n   )       =       K   ·   s       s   +   a         ,       where                 s     =       j   ·   2   ·   π       n   ·   T         ,       for                 n     =   0     ,   1   ,   2   ,   3   ,     4                 …                 and     ,                          
     K and a represent real numbers dependent upon the aircraft design and the lead response desired. Typical values for K and a are on the order of one (1). The sampling rate of the discrete transfer function  401  depicted in FIG. 2 suitably has a period of approximately 0.05 seconds. Alternatively, another preferred discrete transfer function  401  is;            Z        (   z   )       =       K   ·   s       s   +   a         ,       where                 s     =       (     2   /   T     )     *       (     z   -   1     )     /     (     z   +   1     )           ,   and                          
     K and a represent real numbers dependent upon the aircraft design and the lead response desired and T is the period of the continuous update rate, for example, approximately 0.05 seconds. 
     Additional inputs to the discrete washout-filter  400  are the sampling input  400 . 4 , an enable input  400 . 3 , an initial condition input  400 . 2 , and the discrete filter input  400 . 1 . The initial condition input  400 . 2  provides the discrete filter output  400 . 5  with a start up and default value that is benign to the engine servos  56  and  57 . The value of the initial condition input  400 . 2  ensures that no banking command will be applied to the engine servos  56  and  57  when the engines-only flight control system is enabled for flight control. The enable input  400 . 3  switches the output  400 . 5  of the discrete filter  400  between the value on the initial condition input  400 . 2  and the accumulated angle difference values. 
     When the enable input  400 . 3  is logic high, the output  400 . 5  of the discrete filter  400  is the value on the initial condition input  400 . 2 . Else, if the enable input  400 . 3  is at a logic low level, then the output  400 . 5  of the discrete filter  400  is the result of the discrete transfer function  401  operation on the accumulated angle difference values. The accumulated angle difference values are input to the discrete filter  400  at the discrete filter input  400 . 1 . 
     The result of the discrete filter  400  operations on the accumulated angle difference values is available at the filter output  400 . 5 . This output  400 . 5  is summed with the prior art output servo parameter  32  in the first summation block  60  and the aggregate servo parameter is applied to the engine servos  56  and  57  as depicted in FIG.  1 .