Abstract:
A method for sensing a force applied to an aircraft includes receiving a derivative of the acceleration of a motion of a portion of the aircraft, determining whether the derivative of the acceleration of the motion of the portion of the aircraft exceeds a threshold, and outputting an indication that a force has been applied to the portion of the aircraft responsive to determining that the derivative of the acceleration of motion of the portion of the aircraft exceeds the threshold.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims the benefit of U.S. Provisional Application No. 61/255,547, filed Oct. 28, 2009. 
     
    
     BACKGROUND OF THE INVENTION 
       [0002]    The subject matter disclosed herein relates to detecting impact forces on aircraft, and in particular to detecting landing gear impact on aircraft. 
         [0003]    Aircraft such as, for example, rotary wing aircraft and fixed wing aircraft use a variety of sensors to provide feedback to aircraft control systems. Detecting when a force, such as weight, is applied to the landing assemblies or other portions of an aircraft provides useful feedback to aircraft systems. Previous systems used sensors located on each landing assembly to determine whether weight was applied to a landing assembly. The use of these sensors increased the weight and complexity of the aircraft, and had limited fidelity in sensing actual weight applied to a landing assembly. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0004]    According to one aspect of the invention, a method for sensing a force applied to an aircraft includes receiving a derivative of the acceleration of a motion of a portion of the aircraft, determining whether the derivative of the acceleration of the motion of the portion of the aircraft exceeds a threshold, and outputting an indication that a force has been applied to the portion of the aircraft responsive to determining that the derivative of the acceleration of motion of the portion of the aircraft exceeds the threshold. 
         [0005]    According to another aspect of the invention, a method for sensing a takeoff of an aircraft includes receiving a rate of change in the vertical motion of the aircraft, determining whether the rate of change in the vertical motion of the aircraft exceeds a first threshold, integrating the rate of change in the vertical motion of the aircraft and outputting a virtual altitude signal, responsive to receiving the indication that the portion of the aircraft is contacting a surface, delaying the virtual altitude signal through a discrete low pass filter and outputting the delayed virtual altitude signal, subtracting the delayed virtual altitude signal from the virtual altitude signal to output an altitude perturbation signal, determining whether the altitude perturbation signal exceeds a second threshold value, and outputting an indication that the portion of the aircraft is not contacting the surface responsive to determining that the rate of change in the vertical motion of the aircraft exceeds the first threshold and determining that the altitude perturbation signal exceeds the second threshold value. 
         [0006]    According to yet another aspect of the invention, a system for sensing a force applied to an aircraft includes a sensor, and a processor operative to receive a signal indicative of an acceleration of a motion of the aircraft, apply a kinematic equation to the first signal to transform the indication of the acceleration of the motion of the aircraft to indicate an acceleration of a motion of a portion of the aircraft, calculate a derivative of the acceleration of the motion of the portion of the aircraft, determine whether the derivative of the acceleration of the motion of the portion of the aircraft exceeds a threshold, output an indication that a force has been applied to the portion of the aircraft responsive to determining that the derivative of the acceleration of the motion of the portion of the aircraft exceeds the threshold. 
         [0007]    These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWING 
         [0008]    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: 
           [0009]      FIG. 1  illustrates a block diagram of an exemplary embodiment of an aircraft  100 . 
           [0010]      FIG. 2  illustrates a block diagram of an exemplary embodiment of logic performed by the processor of  FIG. 1 . 
           [0011]      FIG. 3  illustrates an example of the geometric relationship between a sensor and a nose landing assembly of  FIG. 1 . 
           [0012]      FIG. 4  illustrates a block diagram of exemplary impact detection logic of  FIG. 2 . 
           [0013]      FIG. 5  illustrates a block diagram of exemplary takeoff detection logic of  FIG. 2 . 
       
    
    
       [0014]    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 
       [0015]      FIG. 1  illustrates a block diagram of an exemplary embodiment of an aircraft  100 . The aircraft  100  includes a nose landing assembly  101 , a left landing assembly  103 , and a right landing assembly  105 . The landing assemblies may include, for example, a landing gear assembly that includes an inflatable wheel, or any other device that is operative to contact a landing surface. For example a skid assembly may be used, and portions of the skid assembly may be designated as contact points similar to the gear described above. The aircraft  100  includes a processor  102  that is communicatively connected to flight controls  104  and sensors  106  that may include, for example, a gyro sensor, one or more accelerometers, a global positioning system (GPS), or any other inertial sensors. The processor  102  may also be communicatively connected to a memory  110  and a display  108 . 
         [0016]      FIG. 2  illustrates a block diagram of an exemplary embodiment of logic performed by the processor  102 . In this regard, the processor  102  receives input data from the sensors  106 . The input includes acceleration (ax, ay, and az) from, for example, an accelerometer, velocity (U, V, W) from, for example, a GPS or derived from an accelerometer, orientation (pitch, roll, yaw; θ, σ, φ) from for example, a gyroscope (gyro), and a rate of change in orientation (P, Q, R) from, for example, a gyro. In block  202  the signals are processed to mathematically transform vectors associated with the signals at the location of the sensors to positions associated with each gear. For example, the accelerometer may be located close to the center of mass of the aircraft  100 , however the gear are located geometrically in different locations. The geometric relationship between the accelerometer and a particular gear may be measured or known, allowing the input from the accelerometer to be mathematically transformed using a kinematic relationship such that the transformed inputs represent acceleration at a particular gear. The processed sensor data is sent to impact detection logic  204 , for landing evolutions, or takeoff detection logic  206 , for takeoff evolutions. The impact detection logic  204  and takeoff detection logic  206  output a signal to the force on gear logic  208  that outputs a force on gear signal  210 . The force on gear signal  210  indicates that a weight on wheel force has been applied to a gear. The indication provides information to the aircraft  100  operator and/or automatic control systems of the aircraft  100  that assists in operating the aircraft. Particularly, the weight on wheel force may indicate that the aircraft has landed or has taken off from a landing area. 
         [0017]      FIG. 3  illustrates an example of the geometric relationship between a sensor  106  and the nose landing assembly  101  including an example of coordinate systems that are associated with the sensor  106  and the nose landing assembly  101 . A kinematic transform may be used to mathematically associate the data collected by the sensor  106  to the nose landing assembly  101 . Thus, for example, a movement sensed by the sensor  106  in the X 1  direction, may be kinematically transformed to an associate the movement with a force applied to the nose landing assembly  101 . A vector representing the force applied to the nose landing assembly  101  may be plotted on the X 2 , Y 2 , Z 2  coordinate system. 
         [0018]      FIG. 4  illustrates a block diagram of exemplary impact detection logic  204  (of  FIG. 2 ) used to determine if a force has been applied to a gear on the aircraft  100 . The logic  204  may be applied in a similar manner to each gear. For exemplary purposes, the description below will describe logic used to determine whether a force or weight has been applied to the nose landing assembly  101  (of  FIG. 1 ), however the logic may be applied simultaneously to any landing assembly or portion of a landing assembly. In this regard, if the aircraft landing is expected, a signal  408  is output that cues the impact detection logic  204 . Vertical jerk data  402  is compared to a vertical jerk threshold value  401 . The vertical jerk data is the derivative of the acceleration in a vertical direction. If the vertical jerk data  402  is greater than the vertical jerk threshold value  401  a signal indicating that the threshold is exceeded is output. Rolling jerk data  404  is compared to a rolling jerk threshold value  403 . Rolling jerk data  404  is a derivative of the acceleration of the roll. If the rolling jerk data  404  is greater than the rolling jerk threshold value  403 , a signal indicating that the threshold is exceeded is output. Pitching jerk data  406  is compared to a pitching jerk threshold  405 . The pitching jerk data  406  is a derivative of the acceleration of the pitch. If the pitching jerk data  406  is greater than the pitching jerk threshold value  405 , a signal indicating that the threshold is exceeded is output. The signals are output to an AND logic that determines whether each of the three thresholds have been exceeded. The force on gear logic  208  outputs a force on gear signal  210 , set to true, that indicates that a force has been applied to the nose landing assembly  101 . If a takeoff signal  410  is output (true) by the takeoff detection logic  206 , it resets the force on gear signal  210  to false. 
         [0019]    The illustrated embodiment above describes the logic associated with the nose landing assembly  101 , however the logic may be used to determine an impact, force, or weight that is applied to any gear, or location on the aircraft  100 . Regarding the nose landing assembly  101 , a force from the ground (or weight) creates positive pitching signals and negative vertical jerk signals. A force (or weight) on the left gear  103  creates a positive rolling jerk signal and negative pitching jerk and negative vertical jerk signals. A force (or weight) on the right gear  105  creates negative rolling jerk, negative pitching jerk, and negative vertical jerk signals. The thresholds may be determined by design parameters, and the geometry of the aircraft  100 . 
         [0020]      FIG. 5  illustrates a block diagram of exemplary embodiment of takeoff detection logic  206  (of  FIG. 2 ). A vertical rate of acceleration of the gear  502  is compared with a vertical rate threshold  501 . A signal  503  is output if the vertical rate of acceleration of the gear  502  is greater than the vertical rate threshold  501 . The forces on the gear signals  508  are compared using OR logic, if either gear force signal is true, indicating ground contact condition, the vertical rate of the gear  502  is integrated to output a virtual altitude signal  510 . The signal  510  is delayed through a discrete low pass filter in block  506  outputting a delayed virtual altitude signal  511 . Altitude perturbation signal  512  is computed by subtracting the delayed virtual altitude signal  511  from the virtual altitude signal  510  and then is compared to an altitude threshold  505 . If the value of the altitude perturbation signal  512  is greater than the altitude threshold  505 , a signal  507  is output. If the signals  503  and  507  are received at the force on ground logic  208 , takeoff signals  509  and  410  are output (set to true). If neither gear force signal is true, indicating in air condition, the integrator input is set to zero thus disabling the take off detection. 
         [0021]    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. 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.