Abstract:
A rotor state sensor system is provided for use with a rotor including a hub, a hub arm and a blade coupled to the hub by the hub arm. The sensor system includes sensors disposed on the hub arm to define a first plane, which emit emissions and receive reflected emissions, and which generate a signal according to the received reflected emissions, reflector plates disposed on the blade which define a second plane at locations where the emissions from the sensors are incident on the reflector plates and from which the reflected emissions are reflected towards the sensors and a computing device which receives the signal from the sensors, determines relative orientations of the first and second planes according to the received signal and determines a condition of the rotor based on the determined relative orientations.

Description:
FEDERAL RESEARCH STATEMENT 
       [0001]    This invention was made with government support under contract no. W911W6-10-2-0004 awarded by the Department of the Army. The government has certain rights to the invention. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    The subject matter disclosed herein relates to a rotor state sensor system and, more particularly, to a three point displacement rotor state system. 
         [0003]    Fly-by-wire (FBW) control systems provide for helicopter stability, response predictability and maneuvering agility while allowing pilots to effectively manage mission and situational awareness. Modern vehicle management systems (VMS) are becoming highly integrated and comprehensive, which effectively protects the aircraft from vibration, provides condition based maintenance, improves maneuvering capability and adapts to mission and environmental demands. At the same time, a unique and critical system on a helicopter is the rotor system, which is highly complex and consists of numerous moving parts. It would be beneficial for the vehicle control and management systems noted above to utilize rotor information, although current production rotors have few (if any) flight control or sensors in the rotating frame. 
         [0004]    The rarity of rotors with sensors in the rotating frame can be attributed to many factors including, but not limited to, input/output (I/O) processing increases, increased numbers of sensors in total and increased redundancy requirements. Additional factors can include the fact that the rotor environment can be harsh for sensor equipment, the fact that the available sensors may have low quality and high costs and the fact that it can be difficult to transmit sensor data from the rotating frame to the non-rotating frame. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0005]    According to one aspect of the invention, a rotor state sensor system is provided for use with a rotor including a hub, a hub arm and a blade coupled to the hub by the hub arm. The sensor system includes sensors disposed on the hub arm to define a first plane, which emit emissions and receive reflected emissions, and which generate a signal according to the received reflected emissions, reflector plates disposed on the blade which define a second plane at locations where the emissions from the sensors are incident on the reflector plates and from which the reflected emissions are reflected towards the sensors and a computing device which receives the signal from the sensors, determines relative orientations of the first and second planes according to the received signal and determines a condition of the rotor based on the determined relative orientations. 
         [0006]    In accordance with further embodiments, the rotor state sensor system further includes a transmission system by which signals are transmittable between the sensors and the flight computer. 
         [0007]    In accordance with further embodiments, the reflector plates are disposed on a radially inward portion of the blade. 
         [0008]    In accordance with further embodiments, the condition includes at least one or more of a lead/lag condition and a flapping condition. 
         [0009]    In accordance with further embodiments, a helicopter is provided and includes a non-rotating frame in which the computing device is disposed and a rotating frame comprising the rotor, the sensors and the reflector plates, wherein the computing device is further configured to adjust commanded pitch angles in accordance with the signals. 
         [0010]    According to another aspect of the invention, a rotor state sensor system is provided and includes a rotor including a hub, a hub arm and a blade coupled to the hub by the hub arm, sensors disposed on the hub arm to define a first plane, reflector plates disposed on the blade such that emissions generated by the sensors define a second plane at locations where the emissions are incident on the reflector plates and a computing device receptive of first and second signals from the sensors, the first and second signals being indicative of relative orientations of the first and second planes with respect to one another. The computing device is configured to determine at least one or more of a lead/lag condition and a flapping of the rotor based on the signals. 
         [0011]    In accordance with further embodiments, the rotor state sensor system further includes a transmission system by which the first and second signals are transmittable between the sensors and the flight computer. 
         [0012]    In accordance with further embodiments, the reflector plates are disposed on a radially inward portion of the blade. 
         [0013]    In accordance with further embodiments, a helicopter is provided and includes a non-rotating frame in which the computing device is disposed and a rotating frame comprising the rotor, the sensors and the reflector plates, wherein the computing device is further configured to adjust commanded pitch angles in accordance with the signals. 
         [0014]    According to yet another aspect of the invention, a method of operating a helicopter is provided and includes receiving pilot commands, converting the received pilot commands into commanded pitch angles for a blade of a rotor, determining whether the commanded pitch angles and actual pitch angles are in line with one another and adjusting the commanded pitch angles in an event the actual pitch angles are different from the commanded pitch angles. 
         [0015]    In accordance with further embodiments, the determining comprises determining relative orientations between a first plane associated with a hub arm of the rotor and a second plane associated with the blade. 
         [0016]    These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0017]    The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which: 
           [0018]      FIG. 1  is a perspective view of an aircraft in accordance with embodiments; 
           [0019]      FIG. 2  is an enlarged view of a portion of a main rotor of the aircraft of  FIG. 1 ; 
           [0020]      FIG. 3  is a schematic diagram of a flight computer of the aircraft of  FIG. 1  in accordance with embodiments; 
           [0021]      FIG. 4  is a plan view of components of the main rotor of the aircraft of  FIG. 1  in accordance with embodiments; 
           [0022]      FIG. 5  is a radial view taken along line A-A in  FIG. 4 ; 
           [0023]      FIG. 6  is a schematic diagram of a first angle formed between first and second planes; 
           [0024]      FIG. 7  is a schematic diagram of a second angle formed between first and second planes; and 
           [0025]      FIG. 8  is a flow diagram illustrating a method of controlling an aircraft in accordance with embodiments. 
       
    
    
       [0026]    The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings. 
       DETAILED DESCRIPTION OF THE INVENTION 
       [0027]    As will be described below, a rotor state sensor system is provided and serves to measure blade motion or blade position during flight. The system can be employed in order to improve performances of control and management systems, to reduce loads and to monitor blade health. 
         [0028]    With reference to  FIGS. 1 and 2 , an aircraft  10  is provided and may be a helicopter  11  or another similar type of aircraft. The helicopter  11  includes a non-rotational frame (i.e., an airframe)  12  and rotational frames  13 . One or more of the rotational frames  13  may be provided on a top portion of the non-rotational frame  12  as a main rotor  120  and another may be provided at a rear of the non-rotational frame  12  as a tail rotor  121 . An engine and a transmission are contained within the non-rotational frame  12  to drive rotation of the rotational frames  13  about rotational axes  130  and  131  with respect the non-rotational frame  12 . Such rotation of the main rotor  120  provides for lift and thrust of the helicopter  11  and rotation of the tail rotor  121  provides for attitude control of the helicopter. 
         [0029]    While the aircraft  10  is shown as a coaxial aircraft, it is understood that aspects can be used in non-coaxial aircraft or other types of coaxial aircraft having other configurations. 
         [0030]    The main rotor  120  rotates about rotational axis  130  and includes a hub  14  through which the rotational axis  130  is defined, a plurality of hub-arms  15  and a plurality of blades  16 . Each blade  16  extends radially outwardly from the hub  14  and includes an inner radial portion  160  that is coupled to the hub  14  by way of an associated one of the hub arms  15 . During flight, the blades  16  can be controlled cyclically or collectively to pitch around pitch axes  161  that are defined along longitudinal lengths of each of the blades  16 . However, in practice, each of the blades  16  may tend to pitch around the corresponding pitch axis  161  more or less than desired and additionally may exhibit lead/lag errors in the circumferential dimension  162  as well as flapping errors in the height-wise dimension  163 . 
         [0031]    With reference to FIGS.  1  and  3 - 5 , the helicopter  11  may further include a flight computer  20 , a rotor state sensor system  25  and a transmission system  26 . The flight computer  20  is supportively disposed in the non-rotational frame  12 . The flight computer  20  controls the driving of the main rotor  120  and the tail rotor  121  as well as the cyclic and collective control of the blades  16 . The flight computer  20  is responsive to pilot commands and flight control algorithms some of which will be described below. The rotor state sensor system  25  is provided at least partially in the rotational frame  12  and may be communicative with the flight computer  20  by way of the transmission system  26  such that the flight computer  20  can be receptive of information or signals from the rotor state sensor system  25 . The received information or signals can be employed by the flight computer  20  to alter or govern the flight control algorithms. The transmission system  26  can be a wired and/or wireless data transmission system. 
         [0032]    A connection between a single blade  16  and the hub  14  includes a single hub arm  30  and a connector element  31 . The hub arm  30  extends radially outwardly from the hub  14  and the connector element  31  serves to couple the inner radial portion  160  of the blade  16  with an outer radial portion  300  of the hub arm  30 . As shown in  FIG. 4 , the hub arm  30  may include a narrow portion  301  that is proximate to the hub  14 , a wide portion  302  that extends out toward the outer radial portion  300  and a tapered portion  303  that connects the narrow portion  301  to the wide portion  302 . The connector element  31  may include a spindle element, an elastomeric bearing or an inner nut of an elastomeric bearing. However, it is understood that the hub arm  30  can have other shapes and configurations in other aspects of the invention. 
         [0033]    In accordance with embodiments, the rotor state sensor system  25  may include a plurality of sensors  40  and a plurality of reflectors  41 . Each sensor  40  is mounted on or near the outer radial portion  300  of the hub arm  30  and each reflector plate  41  is mounted on or near the inner radial portion  160  of the blade  16 . Each of the sensors  40  can be configured to provide and issue a first signal S 1  (see  FIG. 3 ) that is indicative of a position of the sensor  40  on the hub arm  30 , or such positions can be programmed. The sensors  40  also each generate an emission directed toward an associated one of the reflector plates  41 . The locations on each of the reflector plates  41  (such as the distance and/or angles relative to the sensors  40 ) where the corresponding emissions are incident are detectable or determinable by the sensor  40  and/or by additional sensors in accordance with a direction of reflection of the emissions. Alternatively, the distances between the respective positions of the sensors  40  and the respective positions of the incident locations of the reflector plates  41  are detectable or determinable by the sensor  40  or by the additional sensors. 
         [0034]    In any case, with the incident locations known, each of the sensors  40  is further configured to provide and issue a second signal S 2  (see  FIG. 3 ) that is indicative of respective positions of the incident locations of the reflector plates  41 . The combined first signals of the sensors  40  define a first plane P 1  at or near the outer radial portion  300  of the hub arm  30  and the combined second signals define a second plane P 2  at or near the inner radial portion  160  of the blade  16 . 
         [0035]    As shown in  FIG. 3 , the first and second signals S 1  and S 2  may be transmitted from the sensors  40  to the flight computer  20  by way of the transmission system  26 . The transmission system  26  may include transmitters coupled to the sensors  40  and disposed in the rotational frame  12  and receivers coupled to the flight computer  20  and disposed in the non-rotational frame  13 . The transmitters may be components of the sensors  40  and the receivers may be components of the flight computer  20 . 
         [0036]    With the first and second signals S 1  and S 2  transmitted to the flight computer  20  by way of the transmission system  26 , the flight computer  20  is configured to recognize the first and second planes P 1  and P 2  and their respective orientations and to ascertain from such recognition the relative orientations of the first and second planes P 1  and P 2  with respect to one another. With reference to  FIGS. 6 and 7  and, from the relative orientations of the first and second plane P 1  and P 2 , a lead/lag condition of the blade  16  can be identified (see  FIG. 6 ) in accordance with an angle α 1  defined by the first and second planes P 1  and P 2  in the plane of the main rotor  120  and a flapping condition of the blade  16  can be identified (see  FIG. 7 ) in accordance with an angle α 2  defined by the first and second planes P 1  and P 2  relative to the rotational axis  130 . 
         [0037]    As an example, an operation of the plurality of sensors  40  and the plurality of reflectors  41  in accordance with embodiments will now be explained. The plurality of sensors  40  includes sensor  1 , sensor  2  and sensor  3  whose measurements are d 1 , d 2  and d 3 , respectively, and whose sensor points are as follows. 
         [0000]        s   p 1=[0, y 1, z 1] T    
         [0000]        s   p 2=[0, y 2, z 2] T    
         [0000]        s   p 3=[0, y 3, z 3] T    
         [0038]    The plurality of reflectors  41  includes reflector  1 , reflector  2  and reflector  3 , whose reflector points are as follows. 
         [0000]        r   p 1=[ d 1, y 1, z 1] T    
         [0000]        r   p 2=[ d 2, y 2, z 2] T    
         [0000]        r   p 3=[ d 3, y 3, z 3] T    
         [0039]    The next step is to determine a vector normal to the first plane P 1  (s normal ) and normal to the second plane P 2  (r normal ). Since the first plane P 1  is aligned with the coordinate axis, the vector normal to the first plane P 1  is assumed to be s normal =[1, 0, 0] T . The cross-products between two vectors in the second plane P 2  are used to determine the reflector normal vector, r normal . 
         [0000]        r   normal =( r   p 1− r   p 2)×( r   p 1− r   p 3)
 
         [0040]    The flapping angle β and lead-lag angle ζ are determined by the angles of r normal  relative to s normal . To determine β, the reflector normal vector r normal  is projected into the x-z-plane, and to determine, the r normal  is projected into the x-y-plane. The equations for β and ζ are as follows. 
         [0000]      β=tan −1 ( r   normal (3)/ r   normal (1))
 
         [0000]      ζ=tan −1 ( r   normal (2)/ r   normal (1))
 
         [0041]    It should also be noted that when a hub already has “pre-cone” and “pre-lag” (also known as “thrust offset”), that should be added to the flapping and lead-lag values. Also, the distance between the P 1  and P 2  planes (d s→r ) can be determined from the average of the sensor readings as follows: 
         [0000]        d   s→r =Σ(from i=0 to  N   s ) d   i =( d   1   +d   2   +d   3 )/3
 
         [0000]    where i is the sensor index, N s  is the number of sensors and d i  is the distance measurement for sensor i. In order to determine the extension of the thrust bearing d ext , the nominal distance d nom  between planes P 1  and P 2  would need to be subtracted from d s→r . 
         [0000]        d   ext   =d   s→r   −d   nom . 
         [0042]    In accordance with embodiments, the mounting of the sensors  40  on or near the outer radial portion  300  of the hub arm  30  and then mounting of the reflector plates  41  on or near the inner radial portion  160  of the blade  16  (i.e., the root of the blade  16 ) allows for accuracy in the measurements of the lead/lag and the flapping conditions. That is, the measurements of the conditions include little to no effects of high-order blade bending and torsional deflection due to the reflector plates  41  being located at the blade  16  root, for example. 
         [0043]    As shown in  FIGS. 4 and 5  and, in accordance with further embodiments, rotor state sensor system  25  may further include accelerometers  43  and strain gages  44 . The accelerometers  43  may be disposed proximate to each of the sensors  40  and each of the reflector plates  41  and can generate measurement results indicative of first order motion of the blade  16  relative to the hub arm  30  that can serve as a steady reference point for comparison with and verification of the sensor  40  measurements. The strain gages  44  may be disposed proximate to the tapered portion  303  of the hub arm  30  such that strain may be measured on the hub arm  30  inboard of the sensors  40 . Strain in the hub arm  30  may be proportional to blade  16  lagging and should be correlated with the sensor  40  measurements. 
         [0044]    With reference to  FIG. 8 , a method of controlling the aircraft  10  will now be described. As shown in  FIG. 8 , pilot commands (i.e., cyclic and collective commands) are received by the flight computer  20  in operation  50  and those commands are converted into commanded pitch angles for each of the blades  16  in operation  51 . At this point, the rotor state sensor system  25  is engaged to determine whether the actual pitch angles for the blades  16  are in-line with the commanded pitch angles in operation  52 . In an event the actual pitch angles are different from the commanded pitch angles by a predefined degree, the flight computer  20  readjusts the commanded pitch angles in operation  53  to refine the pitch of the blades  16 . 
         [0045]    In accordance with further embodiments and, as shown in  FIG. 8 , operation  52  may include a determination of sensor  40  positions and an additional determination of distances between the sensors  40  and the positions of the corresponding emissions incidence locations at operation  521 . Operation  52  may further include, at operation  522 , a determination of a vector normal to plane P 1 , a determination of a vector normal to P 2 , a projection of the normal vector to plane P 2  into x-y and x-z planes and a resolution of lead/lag and flapping conditions from the projections. Operation  52  is completed with the resolved lead/lag and flapping conditions being output at operation  523  and, for example, used by the flight computer  20  to adjust the commanded pitch angles for the blades  16  in operation  53 . 
         [0046]    While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. By way of example, aspects can be used on fixed wing aircraft, wind turbine blade control, maritime blade control, or any other implementation where blade position needs to be accurately assessed. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.