Patent Publication Number: US-2023133313-A1

Title: On the Ground Detection for VTOL Aircraft

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     The present application claims the benefit of U.S. Provisional Application No. 63/275,032, filed Nov. 3, 2021. 
    
    
     TECHNICAL FIELD OF THE DISCLOSURE 
     The present disclosure relates, in general, to landing systems on aircraft having a vertical takeoff and landing flight mode and, in particular, to on the ground detection systems that provide an early indication of when the aircraft has transitioned from an in the air condition to an on the ground condition during a landing maneuver. 
     GOVERNMENT RIGHTS 
     This invention was made with U.S. Government support under Agreement No. W9124P-19-9-0001 awarded by the Army Contracting Command-Redstone Arsenal to the AMTC and a related AMTC Project Agreement 19-08-006 with Bell Textron Inc. The Government has certain rights in the invention. 
     BACKGROUND 
     Vertical takeoff and landing (VTOL) aircraft are capable of taking off and landing without the need for a runway. One example of a VTOL aircraft is a helicopter which is a rotorcraft having one or more rotors that provide lift and thrust to the aircraft. The rotors not only enable hovering and vertical takeoff and landing, but also enable forward, backward and lateral flight. These attributes make helicopters highly versatile for use in congested, isolated or remote areas. Another example of a VTOL aircraft is a tiltrotor aircraft that includes a set of proprotors that can change their plane of rotation based on the operation being performed. Tiltrotor aircraft generate lift and forward propulsion using the proprotors that are typically coupled to nacelles mounted near the ends of a fixed wing. The nacelles rotate relative to the fixed wing such that the proprotors have a generally horizontal plane of rotation in a VTOL flight mode and a generally vertical plane of rotation while cruising in a forward flight mode, wherein the fixed wing provides lift and the proprotors provide forward thrust. 
     In modern fly-by-wire VTOL aircraft, the control logic used by the automatic flight control system may vary depending upon the mission and/or the maneuver being performed. For example, the control logic used during the in the air condition of the aircraft may be different from the control logic used during the on the ground condition of the aircraft. Consequently, it is important for the automatic flight control system to know whether the aircraft is in the air or on the ground including when the aircraft transitions from the in the air condition to the on the ground condition during a landing maneuver. Attempts have been made to use various weight on wheel sensors, such as limit switches or proximity switches, to determine when the air to ground transition has taken place. It has been found, however, that conventional weight on wheel sensors require significant displacement of components within the landing gear in order to obtain an on the ground indication which results in an undesirable delay in switching from the in the air control logic to the on the ground control logic. Accordingly, a need has arisen for an improved on the ground detection system that provides an early indication of when the aircraft has transitioned from air to ground during a landing maneuver. 
     SUMMARY 
     In a first aspect, the present disclosure is directed to a landing gear for an aircraft that includes a landing strut having proximal and distal ends with the proximal end couplable to the fuselage of the aircraft. The landing strut includes a gas chamber, a liquid chamber, a cylinder and a piston that is movable relative to the cylinder between extended and retracted positions. A wheel is coupled to the distal end of the landing strut. A force sensor is disposed between an extend stop surface of the piston and the chamber. Pressurized gas in the gas chamber biases the piston to the extended position such that the force sensor experiences a preload force. The force sensor is configured to detect a reduction in the preload force during a landing maneuver responsive to contact between the wheel and a landing surface. 
     In some embodiments, the landing strut may be a thru-piston shock strut wherein the cylinder includes the liquid chamber and the piston includes the gas chamber. In other embodiments, the landing strut may be a gas over liquid shock strut wherein the piston includes the liquid chamber and the cylinder includes the gas chamber. In certain embodiments, the liquid chamber may be an oil chamber. In some embodiments, the gas chamber may be an air chamber or a nitrogen chamber. In certain embodiments, the force sensor may include one or more load cells, one or more linear variable differential transformers, one or more strain gauges, one or more pneumatic pressure sensors and/or one or more hydraulic pressure sensors. In some embodiments, the force sensor may be a compressible ring having a central opening configured to slidably receive the piston therethrough and having a main body with at least one force sensing element disposed therein. 
     In certain embodiments, the force sensor may be configured to detect the reduction in force from the preload force during load balancing of the landing strut as the piston transitions from tension to compression during the landing maneuver. In such embodiments, the force sensor may be configured to detect the reduction in force from the preload force during load balancing of the landing strut before the extend stop surface of the piston translates relative to the chamber during the landing maneuver. In other embodiments, the force sensor may be configured to detect the reduction in force from the preload force during load balancing of the landing strut when the extend stop surface of the piston translates relative to the chamber during the landing maneuver. In some embodiments, the extend stop surface of the piston may be internal to the cylinder. In other embodiments, the extend stop surface of the piston may be external of the cylinder. 
     In a second aspect, the present disclosure is directed to an aircraft that includes a fuselage and a plurality of landing gear coupled to the fuselage. Each of the landing gear includes a landing strut having proximal and distal ends with the proximal end coupled to the fuselage of the aircraft. The landing strut includes a gas chamber, a liquid chamber, a cylinder and a piston that is movable relative to the cylinder between extended and retracted positions. A wheel is coupled to the distal end of the landing strut. A force sensor is disposed between an extend stop surface of the piston and the chamber. Pressurized gas in the gas chamber biases the piston to the extended position such that the force sensor experiences a preload force. The force sensor is configured to detect a reduction in the preload force during a landing maneuver responsive to contact between the wheel and a landing surface. 
     In some embodiments, the aircraft may be a rotorcraft such as a tiltrotor aircraft. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       For a more complete understanding of the features and advantages of the present disclosure, reference is now made to the detailed description along with the accompanying figures in which corresponding numerals in the different figures refer to corresponding parts and in which: 
         FIGS.  1 A- 1 B  are schematic illustrations of a tiltrotor aircraft having landing gear with an on the ground detection system in accordance with embodiments of the present disclosure; 
         FIGS.  2 A- 2 B  are schematic illustrations of a helicopter having landing gear with an on the ground detection system in accordance with embodiments of the present disclosure; 
         FIGS.  3 A- 3 D  are schematic illustrations of a landing gear with an on the ground detection system in various operating configurations in accordance with embodiments of the present disclosure; 
         FIG.  4    is a top view of a force sensor for an on the ground detection system in accordance with embodiments of the present disclosure; 
         FIGS.  5 A- 5 D  are schematic illustrations of a landing gear with an on the ground detection system in various operating configurations in accordance with embodiments of the present disclosure; and 
         FIGS.  6 A- 6 B  are schematic illustrations of a landing gear with an on the ground detection system in various operating configurations in accordance with embodiments of the present disclosure. 
     
    
    
     DETAILED DESCRIPTION 
     While the making and using of various embodiments of the present disclosure are discussed in detail below, it should be appreciated that the present disclosure provides many applicable inventive concepts, which can be embodied in a wide variety of specific contexts. The specific embodiments discussed herein are merely illustrative and do not delimit the scope of the present disclosure. In the interest of clarity, not all features of an actual implementation may be described in this specification. It will of course be appreciated that in the development of any such actual embodiment, numerous implementation-specific decisions must be made to achieve the developer’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 be a routine undertaking for those of ordinary skill in the art having the benefit of this disclosure. 
     In the specification, reference may be made to the spatial relationships between various components and to the spatial orientation of various aspects of components as the devices are depicted in the attached drawings. However, as will be recognized by those skilled in the art after a complete reading of the present disclosure, the devices, members, apparatuses, and the like described herein may be positioned in any desired orientation. Thus, the use of terms such as “above,” “below,” “upper,” “lower” or other like terms to describe a spatial relationship between various components or to describe the spatial orientation of aspects of such components should be understood to describe a relative relationship between the components or a spatial orientation of aspects of such components, respectively, as the device described herein may be oriented in any desired direction. 
     Referring to  FIGS.  1 A- 1 B  in the drawings, a rotorcraft depicted as a tiltrotor aircraft is schematically illustrated and generally designated  10 . Aircraft  10  includes a fuselage  12 , a wing mount assembly  14  and a tail assembly  16  including tail members  16   a ,  16   b  having control surfaces operable for horizontal and/or vertical stabilization during forward flight. A wing member  18  is supported by wing mount assembly  14 . Located at outboard ends of wing member  18  are propulsion assemblies  20   a ,  20   b . Propulsion assembly  20   a  includes a nacelle depicted as fixed pylon  22   a  that houses an engine and transmission. In addition, propulsion assembly  20   a  includes a mast assembly  24   a  that is rotatable relative to fixed pylon  22   a  between a generally horizontal orientation, as best seen in  FIG.  1 A , a generally vertical orientation, as best seen in  FIG.  1 B . Propulsion assembly  20   a  also includes a proprotor assembly  26   a  that is rotatable relative to mast assembly  24   a  responsive to torque and rotational energy provided via a drive system mechanically coupled to the engine and transmission. Likewise, propulsion assembly  20   b  includes a nacelle depicted as fixed pylon  22   b  that houses an engine and transmission, a mast assembly  24   b  that is rotatable relative to fixed pylon  22   b  and a proprotor assembly  26   b  that is rotatable relative to mast assembly  24   b  responsive to torque and rotational energy provided via a drive system mechanically coupled to the engine and transmission. 
     Aircraft  10  includes a landing gear system including a pair of forward landing gear  28   a  and an aft landing gear  28   b  each including an on the ground detection system. Each of the on the ground detection systems is preferably linked to a flight control system  30  that executes control logic to provide commands to operate the fly-by-wire control system of aircraft  10 . As the control logic used during the in the air condition is different from the control logic used during the on the ground condition for aircraft  10 , the on the ground detection systems incorporated into landing gear  28   a ,  28   b  provide an early indication of when aircraft  10  has transitioned from the in the air condition to the on the ground condition during landing maneuvers such that the proper control logic is executed by flight control system  30 . Each of forward landing gear  28   a  and aft landing gear  28   b  is coupled to fuselage  12  and is rotatable relative thereto such that forward landing gear  28   a  and aft landing gear  28   b  can be retracted into fuselage  12  during flight, as best seen in  FIG.  1 A . 
       FIG.  1 A  illustrates aircraft  10  in airplane or forward flight mode, in which proprotor assemblies  26   a ,  26   b  are rotating in a substantially vertical plane to provide a forward thrust enabling wing member  18  to provide a lifting force responsive to forward airspeed, such that aircraft  10  flies much like a conventional propeller driven aircraft.  FIG.  1 B  illustrates aircraft  10  in helicopter or VTOL flight mode, in which proprotor assemblies  26   a ,  26   b  are rotating in a substantially horizontal plane to provide a lifting thrust, such that aircraft  10  flies much like a conventional helicopter. It should be appreciated that aircraft  10  can be operated such that proprotor assemblies  26   a ,  26   b  are selectively positioned between forward flight mode and VTOL flight mode, which can be referred to as a conversion flight mode. 
     Referring to  FIGS.  2 A- 2 B  in the drawings, a rotorcraft depicted as a helicopter is schematically illustrated and generally designated  40 . The primary propulsion assembly of helicopter  40  is a main rotor assembly  42  powered by one or more engines via a main rotor gearbox. Main rotor assembly  42  includes a plurality of rotor blade assemblies  44  extending radially outward from a main rotor hub. Main rotor assembly  42  is coupled to a fuselage  46  and is rotatable relative thereto. The pitch of rotor blade assemblies  44  can be collectively and/or cyclically manipulated to selectively control direction, thrust and lift of helicopter  40 . A tailboom  48  extends from fuselage  46  in the aft direction. An anti-torque system  50  includes a tail rotor  52  that is rotatably coupled to the aft portion of tailboom  48 . Anti-torque system  50  controls the yaw of helicopter  40 . Helicopter  40  includes a landing gear system including a forward landing gear  54   a  and a pair of aft landing gear  54   b  each including an on the ground detection system. Each of the on the ground detection systems is preferably linked to a flight control system  56  that executes control logic to provide commands to operate the fly-by-wire control system of helicopter  40 . As the control logic used during the in the air condition is different from the control logic used during the on the ground condition for helicopter  40 , the on the ground detection systems incorporated into landing gear  54   a ,  54   b  provide an early indication of when helicopter  40  has transitioned from the in the air condition to the on the ground condition during landing maneuvers such that the proper control logic is executed by flight control system  56 . Each of forward landing gear  54   a  and aft landing gear  54   b  is coupled to fuselage  46  and is rotatable relative thereto such that forward landing gear  54   a  and aft landing gear  54   b  can be retracted into fuselage  46  during flight. 
     Referring now to  FIGS.  3 A- 3 D  of the drawings, a land gear that is representative of landing gear  28   a ,  28   b ,  54   a ,  54   b  is schematically illustrated and generally designated  100 . In the illustrated embodiment, landing gear  100  includes a thru-piston shock strut depicted as having an outer cylinder  102  that defines a liquid chamber referred to herein as oil chamber  104  that contains a hydraulic fluid such as oil therein as indicated by wavy lines  106 . Slidably and sealingly received within and extending through cylinder  102  is a piston  108  that defines a gas chamber referred to herein as air chamber or nitrogen chamber  110  that contains a gas such as air or nitrogen therein as indicated by stippling  112 . Piston  108  also defines a liquid chamber referred to herein as oil chamber  114  that contains a hydraulic fluid such as oil therein as indicated by wavy lines  116 . Fluid communication between nitrogen chamber  110  and oil chamber  114  is prevented by a fluid separator  118  that is positioned between nitrogen chamber  110  and oil chamber  114 . Fluid separator  118  is a floating separator that not only isolates the gas in nitrogen chamber  110  from the liquid in oil chamber  114  but also enables pressure balancing between nitrogen chamber  110  and oil chamber  114 . Oil chamber  114  is in fluid communication with oil chamber  104  via an orifice  120  that regulates the flow of liquid therebetween. 
     In the illustrated embodiment, the lower portion of piston  108  extends through a lower end of cylinder  102  with the lower distal end of piston  108  coupled to a wheel assembly including an axle  122  and a wheel  124  that is rotatably coupled to axle  122 . The upper portion of piston  108  extends through an upper end of cylinder  102  with the upper end of piston  108  coupled to a flange  126  that has a lower surface configured to hold piston  108  in the fully extended position relative to cylinder  102  that is referred to herein as extend stop surface  128  and that is external of cylinder  102 . An on the ground detection system depicted as force sensor  130  is disposed between an upper surface of cylinder  102  and extend stop surface  128  of piston  108 . Force sensor  130  is preferably in communication with the aircraft’s flight control system, such as flight control system  30  of aircraft  10  or flight control system  56  of helicopter  40 , to provide a signal that the aircraft is on the ground such as to provide an early indication of when the aircraft has transitioned from an in the air condition to an on the ground condition during a landing maneuver. 
     The operation of landing gear  100  will now be described. In  FIG.  3 A , landing gear  100  is in the fully extended position. Landing gear  100  would be in this position, for example, when the aircraft has deployed landing gear  100  during flight in preparation for a landing maneuver. In the illustrated embodiment, the pressure of the gas in nitrogen chamber  110  creates the force that causes piston  108  to be in the fully extended position relative to cylinder  102  with extend stop surface  128  preventing any further movement of piston  108  relative to cylinder  102 . In this configuration, force sensor  130  which is disposed between an upper surface of cylinder  102  and extend stop surface  128  of piston  108 , experiences a compressive force referred to herein as a preload force. The preload force is a result of the reaction force between piston  108  and cylinder  102  at extend stop surface  128  created by the gas pressure in nitrogen chamber  110  and the oil pressure in oil chambers  104 ,  114  which places piston  108  in tension. When wheel  124  comes in contact with a landing surface  132 , such as the ground, a landing force is exerted on wheel  124  by the landing surface  132 . In response to the external landing force applied to wheel  124 , landing gear  100  experiences a load balancing process that involves reduction in the tensile force in piston  108 , an increase in the pressure in oil chamber  104  as piston  108  begins to move upwardly relative to cylinder  102  and an increase in the gas pressure in nitrogen chamber  110  as the liquid from oil chamber  104  enters oil chamber  114  through orifice  120 . 
     More specifically,  FIG.  3 B  depicts a time period in which a landing force, as indicted by force arrow  134 , is sufficient to reduce the tension in piston  108  but not yet sufficient to cause displacement of piston  108  relative to cylinder  102  due to the magnitude of landing force  134  and/or the duration of landing force  134 , such as upon the initial contact between the aircraft and landing surface  132 .  FIG.  3 C  depicts a time period in which the landing force has increased as indicted by a larger force arrow  136 , such as during the landing maneuver as more of the weight of the aircraft is supported by landing surface  132 . In this time period, piston  108  is displaced relative to cylinder  102  causing an increase in the pressure in oil chamber  104 . When the pressure in oil chamber  104  exceed that in oil chamber  114 , liquid from oil chamber  104  passes through orifice  120  into oil chamber  114  which causes the volume of oil chamber  114  to increase and the volume of nitrogen chamber  110  to decrease, thereby further compressing the gas in nitrogen chamber  110  and increasing the pressure thereof. The rate at which liquid from oil chamber  104  passes through orifice  120  into oil chamber  114  is determined by the magnitude of landing force  136 , the viscosity of the liquid, the size of orifice  120  and other factors known to those having ordinary skill in the art.  FIG.  3 D  depicts a time period in which the force on wheel  124  represents the landed weight of the aircraft supported by landing gear  100  as indicted by a larger force arrow  138 . In this time period, piston  108  is in a steady state condition, no longer displacing relative to cylinder  102 , and the pressures in oil chamber  104 , oil chamber  114  and nitrogen chamber  110  have equalized, thus completing the load balancing process. 
     Conventional weight on wheel sensors used to determine when the air to ground transition of a VTOL aircraft has taken place, such as limit switches or proximity switches, would typically provide an indication of the on the ground condition when landing gear  100  is in the position depicted in  FIG.  3 C , wherein significant displacement of piston  108  relative to cylinder  102  has occurred. It has been found, however, that in modern fly-by-wire VTOL aircraft, waiting until this significant displacement has occurred results in an undesirable delay in switching from the in the air control logic to the on the ground control logic. The present embodiments solve this problem by positioning force sensor  130  between an upper surface of cylinder  102  and extend stop surface  128  of piston  108  such that, in the fully extended position of landing gear  100 , force sensor  130  experiences the preload force. In this unique configuration, when landing force  134  is sufficient to reduce the tension in piston  108 , as best seen in  FIG.  3 B , force sensor  130  detects a reduction in the preload force at which time force sensor  130  can provide an early indication to the flight control system of the aircraft that the aircraft has transitioned from an in the air condition to an on the ground condition, thereby enabling the flight control system to transition from the in the air control logic to the on the ground control logic in a more timely manner. 
     Referring additionally to  FIG.  4    of the drawings, force sensor  130  will be described in greater detail. In the illustrated embodiment, force sensor  130  has the shape of a ring suitably sized to be positioned between an upper surface of cylinder  102  and extend stop surface  128  of piston  108  and having a central opening  132  configured to receive piston  108  therethrough. Preferably, force sensor  130  is coupled to the upper surface of cylinder  102  such that piston  108  is slidable relative to central opening  132 . Alternatively, force sensor  130  could be coupled to piston  108  and movable therewith relative to cylinder  102 . The main body  134  of force sensor  130  may be formed from a compressible material such as a polymer or an elastomer. Disposed within main body  134  are one or more force sensing elements  136  that are configured to detect the force applied to main body  134  and/or the change in force applied to main body  134  such as to detect a reduction in the preload force during load balancing of the landing gear. Suitable force sensing elements  136  include load cells, linear variable differential transformers, strain gauges, hydraulic pressure gauges and pneumatic pressure gauges. 
     For example, when force sensing elements  136  are transducers selected from the group of load cells, linear variable differential transformers and strain gauges, the transducers would respond to small displacements within main body  134  as the tensile force within piston  108  decreases and the compressive force in force sensor  130  decreases. These small displacements can be detected in the axial and/or circumferential directions of main body  134  depending upon the type and/or orientation of the transducers. When force sensing elements  136  are transducers selected from the group of hydraulic pressure gauges and pneumatic pressure gauges, the transducers would respond to pressure changes measured in fluid capsules or a fluid network within main body  134  as the tensile force within piston  108  decreases and the compressive force in force sensor  130  decreases. Even though force sensor  130  has been described and depicted as having four force sensing elements  136  distributed circumferentially within main body  134 , it should be understood by those having ordinary skill in the art that a force sensor of the present disclosure could have other numbers of force sensing elements both greater than and less than four, including a single force sensing element. In addition, it should be understood by those having ordinary skill in the art that the force sensing elements within a force sensor of the present disclosure could be disposed in any uniform or nonuniform configuration within main body  134 . 
     Referring now to  FIGS.  5 A- 5 D  of the drawings, a land gear that is representative of landing gear  28   a ,  28   b ,  54   a ,  54   b  is schematically illustrated and generally designated  200 . In the illustrated embodiment, landing gear  200  includes an air over oil shock strut depicted as having an outer cylinder  202  that defines a gas and liquid chamber referred to herein as air over oil chamber  204  that contains a gas such as air or nitrogen therein as indicated by stippling  206  and a hydraulic fluid such as oil therein as indicated by wavy lines  210 . Slidably and sealingly received within cylinder  202  is a piston  208  that defines a liquid chamber referred to herein as oil chamber  112  that contains a hydraulic fluid such as oil therein as indicated by wavy lines  214 . Oil chamber  112  is in fluid communication with air over oil chamber  204  via an orifice  216  that regulates the flow of liquid therebetween together with a metering pin  218  that adjusts the size of orifice  216 . 
     In the illustrated embodiment, the lower portion of piston  208  extends through a lower end of cylinder  202  with the lower distal end of piston  208  coupled to a wheel assembly including an axle  222  and a wheel  224  that is rotatably coupled to axle  222 . The upper portion of piston  208  is disposed within cylinder  202  with the upper end of piston  208  coupled to a cylindrical bearing and seal assembly  226  that has a lower surface configured to hold piston  208  in the fully extended position relative to cylinder  202  that is referred to herein as extend stop surface  228  and that is internal to cylinder  202 . An on the ground detection system depicted as force sensor  230  is disposed between an upper surface of a cylindrical bearing and seal assembly  232  of cylinder  202  and extend stop surface  228  of piston  208 . Force sensor  230  is preferably in communication with the aircraft’s flight control system, such as flight control system  30  of aircraft  10  or flight control system  56  of helicopter  40 , to provide a signal that the aircraft is on the ground such as to provide an early indication of when the aircraft has transitioned from an in the air condition to an on the ground condition during a landing maneuver. 
     The operation of landing gear  200  will now be described. In  FIG.  5 A , landing gear  200  is in the fully extended position. Landing gear  200  would be in this position, for example, when the aircraft has deployed landing gear  200  during flight in preparation for a landing maneuver. In the illustrated embodiment, the pressure of the gas in air over oil chamber  204  creates the force that causes piston  208  to be in the fully extended position relative to cylinder  202  with extend stop surface  228  preventing any further movement of piston  208  relative to cylinder  202 . In this configuration, force sensor  230  experiences a compressive force referred to herein as a preload force. The preload force is a result of the reaction force between piston  208  and cylinder  202  at extend stop surface  228  created by the gas pressure in air over oil chamber  204  and the oil pressure in oil chamber  212  which places piston  208  in tension. When wheel  224  comes in contact with a landing surface  234 , such as the ground, a landing force is exerted on wheel  224  by the landing surface  234 . In response to the external landing force applied to wheel  224 , landing gear  200  experiences a load balancing process that involves a reduction in the tensile force in piston  208 , an increase in the pressure in oil chamber  112  as the liquid in oil chamber  112  is forced through orifice  216  and an increase in the pressure in air over oil chamber  204  as piston  208  moves upwardly relative to cylinder  202 . 
     More specifically,  FIG.  5 B  depicts a time period in which a landing force, as indicted by a force arrow  236 , is sufficient to reduce the tension in piston  208  but not yet sufficient to cause displacement of piston  208  relative to cylinder  202  due to the magnitude of landing force  236  and/or the duration of landing force  236 , such as upon the initial contact between the aircraft and landing surface  234 .  FIG.  5 C  depicts a time period in which the landing force has increased as indicted by a larger force arrow  238 , such as during the landing maneuver as more of the weight of the aircraft is supported by landing surface  234 . In this time period, piston  208  is displaced relative to cylinder  202  decreasing the volume of air over oil chamber  204  which further compress gas  206  therein. In addition, the displacement of piston  208  relative to cylinder  202  causes liquid in oil chamber  212  to passes through orifice  216  which regulates the speed at which piston  208  displaces relative to cylinder  202 . The rate at which liquid passes through orifice  216  is determined by the magnitude of landing force  238 , the viscosity of the liquid, the size of orifice  216 , the shape of meter pin  218  and other factors known to those having ordinary skill in the art.  FIG.  5 D  depicts a time period in which the force on wheel  224  represents the landed weight of the aircraft supported by landing gear  200  as indicted by a larger force arrow  240 . In this time period, piston  208  is in a steady state condition, no longer displacing relative to cylinder  202 , and the pressures in air over oil chamber  204  and oil chamber  212  have equalized, thus completing the load balancing process. 
     In the illustrated embodiment, by positioning force sensor  230  between an upper surface of cylinder  202  and extend stop surface  228  of piston  208 , in the fully extended position of landing gear  200 , force sensor  230  experiences the preload force. In this unique configuration, when landing force  236  is sufficient to reduce the tension in piston  208 , as best seen in  FIG.  5 B , force sensor  230  detects a reduction in the preload force at which time force sensor  230  can provide an early indication to the flight control system of the aircraft that the aircraft has transitioned from an in the air condition to an on the ground condition, thereby enabling the flight control system to transition from the in the air control logic to the on the ground control logic. 
     Referring now to  FIGS.  6 A- 6 B  of the drawings, a land gear that is representative of landing gear  28   a ,  28   b ,  54   a ,  54   b  is schematically illustrated and generally designated  300 . In the illustrated embodiment, landing gear  300  includes a thru-piston shock strut depicted as having an outer cylinder  302  that defines a liquid chamber referred to herein as oil chamber  304  that contains a hydraulic fluid such as oil therein as indicated by wavy lines  306 . Slidably and sealingly received within and extending through cylinder  302  is a piston  308  that defines a gas chamber referred to herein as air chamber or nitrogen chamber  310  that contains a gas such as air or nitrogen therein as indicated by stippling  312 . Piston  308  also defines a liquid chamber referred to herein as oil chamber  314  that contains a hydraulic fluid such as oil therein as indicated by wavy lines  316 . Fluid communication between nitrogen chamber  310  and oil chamber  314  is prevented by a fluid separator  318  that is positioned between nitrogen chamber  310  and oil chamber  314 . Fluid separator  318  is a floating separator that not only isolates the gas in nitrogen chamber  310  from the liquid in oil chamber  314  but also enables pressure balancing between nitrogen chamber  310  and oil chamber  314 . Oil chamber  314  is in fluid communication with oil chamber  304  via an orifice  320  that regulates the flow of liquid therebetween. 
     In the illustrated embodiment, the lower portion of piston  308  extends through a lower end of cylinder  302  with the lower distal end of piston  308  coupled to a wheel assembly including an axle  322  and a wheel  324  that is rotatably coupled to axle  322 . The upper portion of piston  308  extends through an upper end of cylinder  302  with the upper end of piston  308  coupled to a flange  326  that has a lower surface configured to hold piston  308  in the fully extended position relative to cylinder  302  that is referred to herein as extend stop surface  328 . An on the ground detection system  330  is coupled the lower distal end of piston  308  opposite the wheel assembly. On the ground detection system  330  is preferably in communication with the aircraft’s flight control system, such as flight control system  30  of aircraft  10  or flight control system  56  of helicopter  40 , to provide a signal that the aircraft is on the ground such as to provide an early indication of when the aircraft has transitioned from an in the air condition to an on the ground condition during a landing maneuver. 
     Landing gear  300  operates substantially similar to landing gear  100  described herein with the exception of on the ground detection system  330 . In  FIG.  6 A , landing gear  300  has been deployed in preparation for a landing maneuver and is approaching a landing surface  332 . In the illustrated embodiment, on the ground detection system  330  includes a rod  334  that extends below wheel  324  and comes in contact with landing surface  332  before wheel  324  comes in contact with landing surface  332 , as best seen in  FIG.  6 B . Rod  334  may be a flexible rod that bends upon contact with landing surface  332  changing electrical properties within rod  334  to provide an early indication of when the aircraft has transitioned from an in the air condition to an on the ground condition. In such embodiments, rod  334  may have uniform flexibility along its length or may have one or more flexible sections that preferentially bend upon contact with landing surface  332 . Alternatively, rod  334  may be a substantially rigid rod that pivots and/or tilts relative to housing  336  to provide the early indication of when the aircraft has transitioned from an in the air condition to an on the ground condition. In some embodiments, rod  334  may be part of a static electricity discharge system that discharges static electricity from the aircraft upon ground contact wherein rod  334  acts as a static discharge wick to provide a grounding path for the static discharge in addition to acting as an on the ground detection sensor. 
     The foregoing description of embodiments of the disclosure has been presented for purposes of illustration and description. It is not intended to be exhaustive or to limit the disclosure to the precise form disclosed, and modifications and variations are possible in light of the above teachings or may be acquired from practice of the disclosure. The embodiments were chosen and described in order to explain the principals of the disclosure and its practical application to enable one skilled in the art to utilize the disclosure in various embodiments and with various modifications as are suited to the particular use contemplated. Other substitutions, modifications, changes and omissions may be made in the design, operating conditions and arrangement of the embodiments without departing from the scope of the present disclosure. Such modifications and combinations of the illustrative embodiments as well as other embodiments will be apparent to persons skilled in the art upon reference to the description. It is, therefore, intended that the appended claims encompass any such modifications or embodiments.