Abstract:
A control system having a series of control loops for determining the upper and lower cyclic limit increments due to the contribution of collective pitch in rotorcraft based at least on nacelle angle and airspeed.

Description:
BACKGROUND 
       [0001]    1. Field of the Invention 
         [0002]    The present application relates generally to flight control systems, and more specifically, to an aircraft flight control system for controlling rotor blade flapping introduced by collective pitch. 
         [0003]    2. Description of Related Art 
         [0004]    All rotor systems are subject to dissymmetry of lift in forward flight. During hover, the lift is equal across the entire rotor disk. As the helicopter gains airspeed, the advancing rotor blade develops greater lift because of the increased airspeed. For example, rotor blades at hover move at 300 knots and in forward flight at 100 knots the advancing blades move at a relative speed of 400 knots and while the retreating blades move at 200 knots. This has to be compensated for in some way, or the helicopter would corkscrew through the air doing faster and faster snap rolls as airspeed increased. 
         [0005]    Dissymmetry of lift is compensated for by blade flapping. Because of the increased airspeed (and corresponding lift increase) on the advancing rotor blade, the rotor blade flaps upward. Decreasing speed and lift on the retreating rotor blade causes the blade to flap downward. This induced flow through the rotor system changes the angle of attack on the rotor blades and causes the upward-flapping advancing rotor blade to produce less lift, and the downward-flapping retreating rotor blade to produce a corresponding lift increase. Some rotor system designs require that flapping be limited by flapping stops which prevent damage to rotor system components by excessive flapping. In addition to structural damage, aircraft control can be compromised if the rotor flaps into the stop. Thus it becomes incumbent on the aircraft designer to control flapping and warn of this hazardous condition. This application addresses this requirement. Although the foregoing developments represent great strides in the area of flapping detection and reduction, many shortcomings remain. 
         [0006]    Previous attempts to reduce flapping by limiting cyclic control inputs, such as was disclosed by U.S. Pat. No. 8,496,199, which is hereby incorporated by reference as if fully set forth, only considered rotor flapping and cyclic control positions as inputs. Furthermore, previous attempts have been forced to first measure flapping and then react to the flapping. For example, in forward flight at speeds greater than 40 KCAS in conversion mode, flapping due to collective can be as high as 1 degree per degree of collective pitch input. This flapping contribution can not be acted upon by previous CPMS implementations until it is sensed. 
         [0007]    Equation (1) shows the upper limits of Control Power Management System (CPMS) CPMS-based longitudinal cyclic limits, respectively. 
         [0000]        B   ULIM   =BB   long +√{square root over (( F   MAX   2   −b   1   2 ))}  (1)
 
         [0000]    Equation (2) shows the lower limits of Control Power Management System (CPMS) CPMS-based longitudinal cyclic limits, respectively. 
         [0000]        B   ULIM   =BB   long −√{square root over (( F   MAX   2   −b   1   2 ))}  (2)
 
         [0000]    where B_ULIM=upper CPMS-based longitudinal cyclic command limit, B_LLIM=lower CPMS-based longitudinal cyclic command limit, BB_long is the longitudinal component of blowback, F_max is the design maximum total flapping, and b — 1 is the lateral component of flapping. 
         [0008]    Experience with tiltrotors has shown that more effective flapping control is possible if collective pitch is made available to the CPMS. 
     
    
     
       DESCRIPTION OF THE DRAWINGS 
         [0009]    The novel features believed characteristic of the embodiments of the present application are set forth in the appended claims. However, the embodiments themselves, as well as a preferred mode of use, and further objectives and advantages thereof, will best be understood by reference to the following detailed description when read in conjunction with the accompanying drawings, wherein: 
           [0010]      FIG. 1  is a side view of a rotary aircraft; 
           [0011]      FIG. 2  is an oblique view of a tiltrotor aircraft; 
           [0012]      FIGS. 3A and 3B  are oblique views of a rotary system; 
           [0013]      FIG. 4  is a schematic of the flight control system according to the preferred embodiment of the present application; 
           [0014]      FIG. 5  is a flow chart depicting the preferred method according to the preferred embodiment of the present application; and 
           [0015]      FIG. 6  is a schematic of the collective integration modification of the control power management subsystem (CPMS). 
       
    
    
       [0016]    While the system and method of the present application is susceptible to various modifications and alternative forms, specific embodiments thereof have been shown by way of example in the drawings and are herein described in detail. It should be understood, however, that the description herein of specific embodiments is not intended to limit the invention to the particular embodiment disclosed, but on the contrary, the intention is to cover all modifications, equivalents, and alternatives falling within the spirit and scope of the process of the present application as defined by the appended claims. 
       DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
       [0017]    The system and method of the present application overcomes the abovementioned problems commonly associated with conventional aircraft control systems. The control system comprises a subsystem adapted to modifying predetermined flight control limits for a particular aircraft. The subsystem determines whether the aircraft is operating within or near an impending hazardous flight condition, which, in the exemplary embodiments, are conditions where excessive blade flapping occurs. Further description and illustration of the control system and method is provided in the figures and disclosure below. 
         [0018]    It will of course be appreciated that in the development of any actual embodiment, numerous implementation-specific decisions will be made to achieve the developer&#39;s specific goals, such as compliance with system-related and business-related constraints, which will vary from one implementation to another. Moreover, it will be appreciated that such a development effort might be complex and time-consuming, but would nevertheless be a routine undertaking for those of ordinary skill in the art having the benefit of this disclosure. 
         [0019]    The system and method of the present application will be understood, both as to its structure and operation, from the accompanying drawings, taken in conjunction with the accompanying description. Several embodiments of the system are presented herein. It should be understood that various components, parts, and features of the different embodiments may be combined together and/or interchanged with one another, all of which are within the scope of the present application, even though not all variations and particular embodiments are shown in the drawings. It should also be understood that the mixing and matching of features, elements, and/or functions between various embodiments is expressly contemplated herein so that one of ordinary skill in the art would appreciate from this disclosure that the features, elements, and/or functions of one embodiment may be incorporated into another embodiment as appropriate, unless described otherwise. 
         [0020]    Because previous attempts were limited by first having to measure flapping before reacting to, a new system and method that predicts flapping resulting from collective pitch is required for aircraft flying at airspeeds greater than 40 KCAS. This system and method works by incrementing the cyclic limits computed with the CPMS system with a component derived from collective pitch. The commanded collective pitch input is processed by an empirically determined gain time a lagged washout which multiples the reconstituted flapping per degree collective pitch derivatives to generate the increments to be added to the upper and lower cyclic limits computed by the CPMS system. In order to add the effect of collective pitch into the CPMS algorithm Equation 1 and Equation 2 are differentiated with respect to collective pitch with the assumptions that F_max is a constant and not a function of collective pitch and cyclic pitch in the blowback relation is likewise not a function of collective pitch. Thereby creating 
         [0000]    
       
         
           
             
               
                 
                   
                     
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         [0021]    where B — 1LL =lower CPMS-based longitudinal cyclic command limit, B — 1UL=upper CPMS-based longitudinal cyclic command limit, Θ — 0 is the collective pitch, a — 1 is longitudinal flapping, F_max is the design maximum total flapping, and b — 1 is the lateral component of flapping. It should be noted that the upper CPMS-based longitudinal cyclic limit based on collective pitch is defined by the “−” sign on the SQRT function and the lower limit is defined by the “+” sign. 
         [0022]    It should be noted that F_max is a function of aircraft variables (e.g., airspeed, nacelle) and tuned using empirical data and knowledge of the accuracy of the flapping measurements and the flapping stop limit. 
         [0023]    Ideally, the single tuning parameter of the algorithm, Fmax, would be set to the design flapping limit. In practice, however, Fmax must be set to be less than the design limit based on considerations of flapping measurement accuracy and flight test results. In the preferred embodiment, Fmax is generally a function of airspeed. However, it will be appreciated that Fmax could be a function of other flight parameters. With provisions in the developmental flight control system to vary parameters in flight, Fmax can be rapidly and efficiently tuned to accommodate the flapping occurring in the worst case maneuvers expected of the aircraft. 
         [0024]    Referring now to the drawings,  FIGS. 1 and 2  show two different rotary aircraft utilizing the flight control system of the present application.  FIG. 1  shows a side view of a helicopter  101 , while  FIG. 2  shows an oblique view of a tiltrotor aircraft  201 . The flight control system is preferably utilized in tiltrotor aircraft  201  during low speeds and with a fixed lateral cyclic. However, it will be appreciated that the control system is easily and readily adaptable for use with other types of rotary aircraft, i.e., helicopter  101 , operating at various speeds and with or without a fixed lateral cyclic control. 
         [0025]    Helicopter  101  comprises a rotary system  103  carried by a fuselage  105 . One or more rotor blades  107  operably associated with rotary system  103  provide flight for helicopter  101  and are controlled with a plurality of controllers within fuselage  105 . For example, during flight a pilot can manipulate the cyclic controller  109  for changing the pitch angle of rotor blades  107 , thus providing lateral and longitudinal flight direction, and/or manipulate pedals  111  for controlling yaw direction. 
         [0026]    Tiltrotor aircraft  201  includes two or more rotary systems  203  carried by rotatable nacelles. The rotatable nacelles enable aircraft  201  to takeoff and land like a conventional helicopter, thus the rotary systems of tiltrotor  201  are susceptible to excessive flapping of the rotor blades  205  caused by control of the rotor blades, rotor system rotation, and the rotor operating environment such as wind speed and direction. In the preferred embodiment, the control system of the present application is carried within fuselage  207  for assisting the pilot during flight. It should be understood that, like helicopter  101 , tiltrotor aircraft  201  comprises a cyclic controller and pedals for manipulating lateral, longitudinal, and yaw control. 
         [0027]    For ease of description, some of the required systems and devices operably associated with the present control system are not shown, i.e., sensors, connectors, power sources, mounting supports, circuitry, software, and so forth, in order to clearly depict the novel features of the system. However, it should be understood that the system of the present application is operably associated with these and other required systems and devices for operation, as conventionally known in the art, although not shown in the drawings. 
         [0028]    Referring to  FIGS. 3A and 3B  in the drawings, oblique views of rotary system  103  are shown.  FIG. 3A  shows rotary system  103  during normal operation, while  FIG. 3B  shows rotary system  103  during hazardous flight conditions, i.e., the rotary system experiencing excessive flapping. Rotary system  103  comprises a mast  301  rotatably attached to rotor blades  107  via a rotor yoke  303 . One or more restraints  305  and/or other nearby structures are positioned alongside mast  301 . In the exemplary embodiment, restraints  305  are conventional “stops” adapted to restrain the tilting of the hub. It should be understood that both helicopter  101  and tiltrotor  201 , along with other types of rotary aircraft, are susceptible to excessive flapping, which could result in damage to the rotary system. 
         [0029]    During flight, the rotation of mast  301  combined with the pitching of rotor blades  107  causes flapping, as depicted with vertical arrows. Excessive flapping can cause yoke  303  to tilt in direction D 1 , as indicated with the vertical arrow, which in turn could cause the yoke to come into contact with restraint  305 , resulting in damage to components of the rotor system and/or restraint  305 , and in some scenarios, resulting in catastrophic failure. It will be appreciated that one of the novel features of the control system of the present application is to assist the pilot in controlling flight of the aircraft to avoid contact between yoke  303  and restraint  305 . 
         [0030]    Referring to  FIG. 4  in the drawing, a schematic view of flight control system  400  is shown. System  400  further comprises a flight control subsystem (FCS)  401  and a control power management subsystem  403  (CPMS). Both FCS  401  and CPMS  403  are operably associated with one another to assist the pilot to avoid excessive flapping. 
         [0031]    Box  405 , labeled as flight control laws (CLAW), depicts the outcome flight control limits generated by both FCS  401  and CPMS  403 . As is shown, a solid line represents the original flight control limits, while the dashed line represents the modified flight control limits, i.e., the solid line being lowered with application of CPMS  403 . It should be understood that CPMS  403  only limits the flight control limits while the aircraft is flying in or near impending hazardous flight conditions, i.e., excessive blade flapping. Optionally, the modified flight control limits may be thereafter displayed to the pilot via a MFD or other suitable display. 
         [0032]    In the preferred embodiment, pilot controller commands  407 , i.e., from cyclic controller  109  and/or pedal  111 , along with automatic aircraft controls  409 , are received by FCS  401 , then relayed to aircraft actuators  411 . 
         [0033]    CPMS  403  is preferably operably associated with a first sensor  413  adapted to sense the angle of the nacelles, a second sensor  415  adapted to sense airspeed, and a third sensor  417  adapted to sense both lateral and longitudinal flapping of the rotor. CPMS  403  is provided with a flapping limiting algorithm, which receives sensed data from sensor  413 , sensor  415 , sensor  417 , and commanded collective pitch  419  from FCS  401  to generate control limit envelopes. As discussed, the nacelle angle and airspeed changes during flight, thereby changing the amount of flapping introduced by collective pitch, thus resulting in changing control limits generated by CPMS  403 . 
         [0034]    Referring to  FIG. 5  in the drawings, a flowchart  501  depicting the preferred method is shown. Box  503  shows the first step, which includes generating control limits for the aircraft, which are predetermined control limits for the particular aircraft. In the preferred method, the combined commanded pilot controls and the automatic aircraft controls are limited by the flight control margins. Box  505  depicts the next step, which includes modifying the control limits to avoid impending hazardous conditions, i.e., excessive flapping. This step is achieved with CPMS via a flapping limiting algorithm operably associated with the aircraft rotary system and the aircraft actuators. The next step morphing the envelope as the aircraft approaches impending hazardous flight conditions, as depicted in box  509 . 
         [0035]    Referring now also to  FIG. 6  in the drawings, a schematic depicting the preferred incorporation of collective pitch into the CPMS module for the left rotor is illustrated.  FIG. 6  (subsystem  600 ) provides a detailed view of the algorithm utilized with subsystem  403 . In particular, the algorithm is implemented in the flight control system software and receives data such as airspeed, nacelle angle, longitudinal and lateral flapping, amount of collective pitch, and the position of the lateral and longitudinal cyclic actuators as inputs. Thereafter, the algorithm modifies the CPMS-based cyclic control limits which may in turn limit the cyclic control commands of the flight control system. It should be appreciated that the algorithm is repeated for each rotor when implemented on a tiltrotor aircraft. 
         [0036]    First control loop  603  determines the lower CPMS limit increment due to collective pitch  605  as shown in Equation 4. The lower CPMS limit increment due to collective pitch  605  once determined, because it is an increment would then be added to the conventional lower CPMS limit to form the adjusted lower CPMS limit. 
         [0037]    The lower CPMS limit increment due to collective pitch  605  is the product  607  of sum  609  and product  611 . Sum  609  is the sum of product  613  together with sum  615 . Sum  615 , the differentiated longitudinal flapping divided by the collective pitch, is the sum of product  617  and the result of constant one lookup table  619 . Constant one lookup table  619  is based on the result of index lookup table  621 . Index lookup table is based upon the nacelle angle  623 . Nacelle angle  623  in the preferred embodiment is the based upon the measured nacelle angle from the first sensor  413 , however it should be apparent that commanded nacelle angle from the FCS  401  is a suitable alternative. Product  613  is the result of multiplying the division  625  with the sum  631 . Division  625  is the result of dividing the lateral flapping  629  by the input one  627 . Lateral flapping  629  in the preferred embodiment is lagged, however other embodiments utilize a non-lagged lateral flapping. Furthermore, lateral flapping  629  is the result of a transducer measuring actual lateral flapping such as the third sensor  417 . 
         [0038]    Input one  627  is the result of Equation 5. 
         [0000]      Inputone=√{square root over (( F   MAX   2   −b   1   2 ))}  (5)
 
         [0000]    where F_max is the design maximum total flapping, and b — 1 is the lagged lateral component of flapping. 
         [0039]    Product  617  is the result of multiplying slope one lookup table  633  by the airspeed  635 . Airspeed  635  in the preferred embodiment is the FCS  401  airspeed, which is based on conditioning of a transducer measuring actual airspeed. Slope one lookup table  633  is based on the result of index lookup table  621 . 
         [0040]    Sum  631 , the differentiated lateral flapping divided by the collective pitch, is the sum of product  637  and the result of constant two lookup table  639 . Constant two lookup table  639  is based on the result of index lookup table  621 . Product  637  is the result of multiplying slope two lookup table  641  by the airspeed  635 . Slope two lookup table  641  is based on the result of index lookup table  621 . 
         [0041]    Product  611  is the result of multiplying gain lookup table  643  with filter  645 . Gain lookup table  643  is based on the results of index lookup table  621 . Filter  645  filters the washedout left collective pitch command  647 . Filter  645  is optional as need to filter out high frequencies. Washedout left collective pitch command  647  is the commanded left collective pitch command  649  from the FCS  401 . The commanded left collective pitch command  649  is washed out to reduce or eliminate trim collective conditions such that only dynamic collective pitch inputs are reacted to by the system  600 . Washedout left collective pitch command  647  has a bias  651  to reset the amount of washout, when the FCS  401  initial condition discrete  653  is TRUE. 
         [0042]    Second control loop  657  determines the upper CPMS limit increment due to collective pitch  659  as also shown in Equation 3. The upper CPMS limit increment due to collective pitch  659  is the product  661  of difference  663  multiplied by product  611 . Difference  663  is product  613  subtracted from sum  615 . The upper CPMS limit increment due to collective pitch  657  once determined, because it is an increment would then be added to the conventional upper CPMS limit to form the adjusted upper CPMS limit. 
         [0043]    The system of lookup tables  619 ,  621 ,  639 ,  641 , and  643  combined are based upon the influence of nacelle angle relative to the amount of longitudinal and lateral flapping per degree of collective pitch. These lookup tables are tuned using empirical data and knowledge of the accuracy of the flapping measurements and the flapping stop limit. 
         [0044]    This system and method provides several benefits to rotorcraft that experience flapping resulting from collective pitch. First, the system and method allow the aircraft to predict flapping resulting from collective pitch. Second, because the aircraft can predict the flapping resulting from collective pitch the reaction to the flapping resulting from collective pitch is with less delay then those systems that must measure the flapping resulting from collective pitch and then react. Third, this system and method does not impact aircraft performance such as lowering the rate limits on collective pitch inputs or limiting power lever input rates, and therefore this system and method does not degrade performance. 
         [0045]    It is apparent that a system and method with significant advantages has been described and illustrated. The particular embodiments disclosed above are illustrative only, as the embodiments may be modified and practiced in different but equivalent manners apparent to those skilled in the art having the benefit of the teachings herein. It is therefore evident that the particular embodiments disclosed above may be altered or modified, and all such variations are considered within the scope and spirit of the application. Accordingly, the protection sought herein is as set forth in the description. Although the present embodiments are shown above, they are not limited to just these embodiments, but are amenable to various changes and modifications without departing from the spirit thereof.