Abstract:
Methods and apparatus are provided for estimating aircraft dynamic pressure. The load on the flight control surface actuator that is coupled to a flight control surface is measured. An estimate of the aircraft dynamic pressure is calculated from the measured load.

Description:
TECHNICAL FIELD 
     The present invention generally relates to dynamic pressure measurement and, more particularly, to a system and method for calculating an estimate of aircraft dynamic pressure from flight control surface actuator load sensors. 
     BACKGROUND 
     Many modern flight control systems rely on various air data parameters, such as dynamic pressure, to calculate aircraft control commands. It is thus undesirable for the flight control system to receive inaccurate air data signals. Typically, dynamic pressure is sensed using one or more pitot tubes. Most often, a pitot tube is located on the wing or front section of an aircraft, with its opening facing forward. As aircraft airspeed varies, ram air pressure at the pitot tube opening also varies. The pressure variations can thus be used to determine and indicate aircraft airspeed. 
     Because the pitot tube opening faces forward, it can become blocked. The sources of potential blockage are numerous and varied, and include ice, water, volcanic ash, dirt, insects, and various other contaminants. A blocked pitot tube can cause inaccurately sensed dynamic pressure and thus, for example, inaccurate aircraft airspeed being determined and displayed. Presently, there are no back-up dynamic pressure sensors beyond the use of redundant pitot tubes, which are also susceptible to blockage. This blockage can occur on a single sensor or as a result of a common mode event such as, for example, icing or volcanic ash. In the unlikely, but possible, event of multiple pitot probe blockages or contaminations, it would be possible for valid, yet erroneous data to be used by the control laws or displayed to the flight crew. 
     Hence, there is a need for a system and method for determining aircraft dynamic pressure that is not susceptible to blockage by debris and various other sources. The present invention addresses at least this need. 
     BRIEF SUMMARY 
     In one embodiment, a method for estimating dynamic pressure applied to an aircraft having a flight control surface actuator coupled to a flight control surface includes measuring a load on the flight control surface actuator. An estimate of the dynamic pressure is calculated from the measured load. 
     In another embodiment, a system for estimating aircraft dynamic pressure includes a sensor and a processor. The sensor is configured to sense a load on a flight control surface actuator and supply actuator load signals representative thereof. The processor is coupled to receive the actuator load signals and is configured to calculate an estimate of the aircraft dynamic pressure from the sensed load. 
     In yet another embodiment, a system for estimating aircraft dynamic pressure includes a plurality of first flight control surface actuators, a plurality of first load sensors, a plurality of second flight control surface actuators, a plurality of second load sensors, and a processor. Each of the first flight control surface actuators is adapted to be coupled to a first aircraft flight control surface. Each of the first flight control surface actuators is further adapted to receive a controlled flow of hydraulic fluid and is configured, upon receipt of the controlled flow of hydraulic fluid, to move the first flight control surface to a control position. Each of the first load sensors is coupled to one of the first flight control surface actuators and is configured to sense a load thereon and supply first flight control surface actuator load signals representative thereof. Each of the second flight control surface actuators is adapted to be coupled to a second aircraft flight control surface. Each of the second flight control surface actuators is further adapted to receive a controlled flow of hydraulic fluid and is configured, upon receipt of the controlled flow of hydraulic fluid, to move the second flight control surface to a control position. Each of the second load sensors is coupled to one of the second flight control surface actuators and is configured to sense a load thereon and supply second flight control surface actuator load signals representative thereof. The processor is coupled to receive the first flight control surface actuator load signals and the second flight control surface actuator load signals and is configured to calculate an estimate of the aircraft dynamic pressure from the first and second flight control surface actuator load signals. 
     Other desirable features and characteristics of the dynamic pressure determination system and method described herein will become apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and the preceding background. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The present invention will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and wherein: 
         FIG. 1  is depicts a functional block diagram of a portion of an aircraft flight control system; and 
         FIG. 2  depicts a functional block diagram of at least a portion of a flight control module that may be used to implement the system of  FIG. 1 . 
     
    
    
     DETAILED DESCRIPTION 
     The following detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. Furthermore, there is no intention to be bound by any theory presented in the preceding background or the following detailed description. 
     Referring first to  FIG. 1 , a functional block diagram of a portion of an aircraft flight control surface actuation system  100  is depicted. In particular, that portion of a flight control surface actuation system  100  that controls the position of two flight control surfaces—a left aileron  102  and a right aileron  104 —is depicted. The depicted portion of the flight control surface actuation system  100  includes an inceptor  106 , a plurality of actuator controls  108 , a plurality of flight control surface actuators  110 , and a flight control module  112 . 
     The inceptor  106  is configured to move in response to an input force supplied from, for example, a pilot. One or more non-illustrated position sensors sense the position of the inceptor  106 , and supply inceptor position signals representative of the sensed inceptor position. It will be appreciated that the inceptor  106  may be implemented using any one of numerous inceptor configurations including, for example, a side stick, a yoke, or a rudder pedal, just to name a few. It will additionally be appreciated that the inceptor  106  may be implemented as an active inceptor or a passive inceptor. No matter its specific implementation, the inceptor  106  supplies the inceptor position signals to the actuator controls  108 . 
     The actuator controls  108  are each coupled to receive the inceptor position signals from the inceptor  106 , and may also receive flight control augmentation data from the flight control module  112 . The actuator controls  108  are each responsive to the inceptor position signals and the flight control augmentation data to generate flight control surface actuator commands. Although the actuator controls  108  may be variously configured, in the depicted embodiment each actuator control  108  includes actuator control electronics (ACE)  116  and a remote electronics unit (REU)  118 . The ACEs  116  are each coupled to receive the inceptor position signals and the flight control augmentation data and are configured, in response thereto, to generate flight control surface position commands, which are supplied to one of the REUs  118 . 
     The REUs  118  are each coupled to receive the flight control surface position commands from its associated ACE  116  and are configured, in response thereto, to supply appropriate actuator position commands to its associated flight control surface actuator  110 . The REUs  118  are additionally coupled to receive one or more control surface position feedback signals from one or more non-illustrated control surface position sensors, and actuator load signals from one or more actuator load sensors  122 . The REUs  118  transmit data representative of control surface position and actuator load (and information about the health of the actuator and its sensors) back to the associated ACEs  116  for subsequent transmission to the flight control module  112 . 
     The flight control surface actuators  110  are each coupled to one of the flight control surfaces. In the depicted embodiment, two flight control surface actuators  110  are coupled to the left aileron  102  and two flight control surface actuators are coupled to the right aileron  104 . It will be appreciated that this is merely exemplary of a particular embodiment. No matter the specific number of flight control surface actuators  110 , each is coupled to receive the flight control surface actuator commands supplied from one of the actuator controls  108 . The flight control surface actuators  110  are each configured, upon receipt of the flight control surface actuator commands, to move its associated flight control surface to the commanded control position. It will be appreciated that the flight control surface actuators  110  may be variously implemented to carry out this functionality. For example, the flight control surface actuators  110  may be implemented using any one of numerous electric, electromechanical, hydraulic, or pneumatic actuators now known or developed in the future. In the depicted embodiment, however, each flight control surface actuator is implemented using a hydraulic actuator that receives a controlled flow of hydraulic fluid via, for example, a servo-control valve  124  and, in response to the controlled flow of hydraulic fluid, moves the flight control surface to which it is coupled to the control position. 
     The actuator load sensors  122  that were mentioned above are each associated with one of the flight control surface actuators  110 . Although  FIG. 1  depicts a single actuator load sensor  122  associated with each flight control surface actuator  110 , it will be appreciated that two or more actuator load sensors  122  may be associated with each flight control surface actuator  110 . It will additionally be appreciated that the actuator load sensors  122  may also be variously implemented. For example, the actuator load sensors  122  may be implemented using any one of numerous know load cells, force sensors, or pressure sensors, just to name a few. In the depicted embodiment, the actuator load sensors  122  are implemented using pressure sensors. The pressure sensors  122  each sense the pressure of the hydraulic fluid in their associated flight control surface actuators  110 , and supply a pressure signal representative thereof to the associated REU  118 . As described above, REUs  118  transmit the pressure data (which are representative of actuator load) to the associated ACEs  116  for subsequent transmission to the flight control module  112   
     The flight control module  112 , among various other functions, generates the above-described flight augmentation data that are supplied to the actuator controls  108 . The flight control module  112  also receives the data representative of control surface position and actuator load from the ACEs  116 . The flight control module  112  is configured to process the data representative of actuator load to determine the individual load on each flight control surface actuator  110  and, if necessary, supply suitable commands to the actuator controls  108  to equalize the loads on the flight control surface actuators  110  that are coupled to the same flight control surface  102 ,  104 . The flight control module  112  is also configured to process the data representative of actuator load and calculate an estimate of the aircraft dynamic pressure from the sensed actuator load. The manner in which the flight control module  112  makes this calculation will now be described. 
     It is generally known that, for a given aileron position, the airload (lift) on the aft section of a wing is approximately proportional to the dynamic pressure. This is shown as follows: 
                 C   P     =         P   UPPER     -     P   LOWER       q       ,     
     ⁢   or                     P   UPPER     -     P   LOWER       =       C   P     *   q       ,         
where:
         C p  is the local pressure coefficient,   P UPPER  is the local pressure on the upper surface of the wing,   P LOWER  is the local pressure on the lower surface of the wing, and   q is the dynamic pressure.       

     It has been shown that the pressure coefficient (Cp) at an aileron  102 ,  104  is primarily impacted by the aileron position and is impacted very little by wing angle of attack (“alpha”) near the aft section of the wing, where the ailerons  102 ,  104  are located. Hence, the local pressure difference (P UPPER −P LOWER ) is proportional to the dynamic pressure (q) and the aileron position. Moreover, if the local pressure difference is integrated over the aileron surface, the resultant is the aileron load. Therefore, since the actuator load (LOAD ACTUATOR ) is proportional to the aileron load, it may be shown that:
 
 LOAD   ACTUATOR   =k ( C   h0   +C   hδ *δail)* q,  
 
and thus:
 
               q   =       LOAD   ACTUATOR       k   ⁡     (       C     h   ⁢           ⁢   0       +       C     h   ⁢           ⁢   δ       *     δ   ail         )           ,         
where:
         k is the mechanical relationship between the pressure on the aileron and load in the actuator,   C h0  is the hinge moment coefficient at zero degrees of aileron deflection,   C hδ  is the hinge moment gradient as a function of aileron position, and   δail is the aileron position.       

     Since the aileron load is sensed, via the actuator load sensors  122 , the flight control module  112  may be readily configured to calculate the estimate of aircraft dynamic pressure using these data. It will be appreciated that in the depicted embodiment the flight control module  112  is additionally configured to determine the aileron load from the sensed hydraulic fluid pressure. 
     Turning now to  FIG. 2 , a functional block diagram of at least a portion the flight control module  112  is depicted and will be described. Before doing so, however, it is noted that the flight control module  112 , including each of the functional blocks depicted in  FIG. 2 , may be implemented using one or more general purpose processors, content addressable memory, digital signal processors, application specific integrated circuits, field programmable gate arrays, any suitable programmable logic devices, discrete gate or transistor logic, discrete hardware components, or any combination thereof, designed to perform the functions described herein. A processor may be realized as a microprocessor, a controller, a microcontroller, or a state machine. A processor may also be implemented as a combination of computing devices, e.g., a combination of a digital signal processor and a microprocessor, a plurality of microprocessors, one or more microprocessors in conjunction with a digital signal processor core, or any other such configuration. 
     No matter how the flight control module  112  is specifically implemented, it is seen that it receives the actuator load signals from each of the load sensors  122  associated with the flight control surface actuators  110  on the left aileron  102  and from each of each of the load sensors  122  associated with the flight control surface actuators  110  on the right aileron. After suitable signal processing  202 , a first summing function  204  sums the actuator load signals from the load sensors  122  associated with the left aileron  102  to determine the left aileron load, and a second summing function  206  sums the actuator load signals from the load sensors  122  associated with the right aileron  104  to determine the right aileron load. An averaging function  208  then averages the left and right aileron loads and, after suitable filtering by a low-pass filter  210 , the average aileron load is supplied to a conversion function  212 . 
     The conversion function  212  calculates the estimate of aircraft dynamic pressure (q) from the average aileron load. As  FIG. 2  also depicts, the conversion function  212  may also receive a signal representative of aileron position. This is because the load on the aileron is also a function of the control surface position. Since the ailerons move in opposite directions to produce an aircraft rolling moment the load on one control surface increases and the load on the other surface decreases. Therefore the average load is not impacted by the surface position. However, in the event that one surface is inoperable, the dynamic pressure can be calculated based on the load and position from only a single healthy aileron. 
     The aircraft dynamic pressure estimate is then supplied to a comparison monitor function  214 . The comparison monitor function  214  compares the aircraft dynamic pressure estimate to a directly sensed aircraft dynamic pressure value supplied from a non-illustrated pressure sensor, such as a pitot probe, via the aircraft avionics. The flight control module  112 , at least in some embodiments, is further configured to supply an output signal that prohibits the use of the directly sensed dynamic pressure by at least portions of aircraft flight control software if the directly sensed dynamic pressure differs from the calculated estimate of the dynamic pressure by a predetermined magnitude. Moreover, in some embodiments, this signal may also command at least portions of the aircraft flight control software to use the calculated estimate of the dynamic pressure if the directly sensed dynamic pressure differs from the calculated estimate of the dynamic pressure by the predetermined magnitude. As  FIG. 2  depicts via a dotted line, in some aircraft the dynamic pressure estimate may be used as the sole source for use by the control laws. 
     While at least one exemplary embodiment has been presented in the foregoing detailed description of the invention, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing an exemplary embodiment of the invention. It being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope of the invention as set forth in the appended claims.