Patent Publication Number: US-8541091-B2

Title: Composite leg for landing gear assembly

Description:
This is a continuation of U.S. Ser. No. 12/341,885 filed Dec. 22, 2008 now U.S. Pat. No. 8,163,368. U.S. Ser. No. 12/341,885 is a continuation-in-part of U.S. Ser. No. 11/145,058 filed 3 Jun. 2005, now U.S. Pat. No. 7,467,763 and U.S. Ser. No. 12/340,631 filed 19 Dec. 2008, now U.S. Pat. No. 7,807,249. U.S. Ser. No. 12/340,631 is a continuation-in-part of U.S. Ser. No. 11/096,727 filed 31 Mar. 2005, now abandoned. 
    
    
     This invention was made with Government support under contract number MDA972-98-9-0004 awarded by the Defense Advanced Research Projects Agency. The government has certain rights in this invention. 
    
    
     BACKGROUND 
     A traditional landing gear assembly for an aerospace vehicle includes components made of steel forgings. However, the steel forgings are heavy, and they have high tooling costs. 
     Composite material may be used instead due to its high strength and light weight. A landing gear leg, for example, may be made from a material such as fiberglass. 
     SUMMARY 
     According to an embodiment herein, a composite leg for a landing gear assembly includes a composite compression cap including reinforcing fibers that provide strength in compression, and a tension cap including reinforcing fibers that provide strength in tension. At least some of the fibers are oriented at +α and −α with respect to a longitudinal axis of the leg, where α is between 2 and 8 degrees. 
     According to another embodiment herein. a landing gear assembly leg comprises a composite compression cap including reinforcing fibers oriented at +α and −α with respect to a longitudinal axis of the leg, and a composite tension cap including reinforcing fibers oriented at +α and −α with respect to the longitudinal axis of the leg. The angle α is between 2 and 8 degrees. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is an isometric view of an aircraft having a landing gear assembly. 
         FIG. 2  is a front elevational schematic view of a portion of the aircraft of  FIG. 1 . 
         FIG. 3  is an enlarged, partial elevational view of a portion of the landing gear assembly of  FIG. 1 . 
         FIG. 4  is an enlarged, partial isometric view of a portion of aircraft of  FIG. 1 . 
         FIG. 5  is an isometric, exploded view of a portion of the landing gear assembly of  FIG. 2 . 
         FIG. 6  is an isometric, cross-sectional view of a composite leg of the landing gear assembly of  FIG. 1   
         FIG. 7  is a representative graph of stress versus strain for carbon, a glass, and a plastic material 
         FIG. 8  is a representative table of various material properties for various fiber materials. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  is an isometric view of an aircraft  100 . In this embodiment, the aircraft  100  includes a fuselage  110  and a rotary lift and propulsion system  150  having a main rotor  152  and a tail rotor  154 . A landing gear assembly  160  projects outwardly from the fuselage  110  and includes a pair of composite legs  162 . A landing wheel  164  is operatively coupled to each composite leg  162  and a tail landing gear  166  projects downwardly from an aft portion of the fuselage  110 . Alternately, the aircraft  100  may be another type of aircraft having a landing gear assembly. 
       FIG. 2  is a front schematic view of the aircraft  100  and landing gear assembly  160  of  FIG. 1 .  FIG. 3  is an enlarged, partial elevational view of a portion of the landing gear assembly  160  of  FIG. 2 . In this embodiment, each of the composite legs  162  is slideably disposed into a box-like receiving member  168  coupled to the fuselage  110 . A pair of attachment members  170  are disposed through each receiving member  168  and composite leg  162 , securing each composite leg  162  to its corresponding receiving member  168 . In a first (or non-loaded) position  172 , such as just prior to landing or shortly after takeoff, each of the composite legs  162  projects downwardly in a relaxed (or non-loaded) orientation to ensure that the landing wheels  164  are in a proper position for landing or take off. In a second (or loaded) position  174 , however, each of the composite legs  162  may be bent upwardly by the forces associated with landing the aircraft  100 . The amount of upward bending of each composite leg  162  depends upon several variables, including the weight of the aircraft  100 , the vertical landing velocity, and the flexibility of the composite legs  162 . 
     As shown in  FIG. 2 , the upward bending of each composite leg  162  may be characterized by a radius of curvature p of the composite leg  162 , and also by a deflection distance d which is the distance the landing wheel  164  is deflected upwardly between the non-loaded position  172  and the loaded position  174 . The radius of curvature p of the composite leg  162  may be estimated by the following equation:
 
ρ= h   B /(ε t +ε c )
 
where h B  is a dimension (or beam height) of the composite leg  162 , ε t  is a strain in tension of the composite leg  162 , and ε c  is a strain in compression of the composite leg  162 .
 
     As further shown in  FIG. 3 , in this embodiment, the receiving member  168  includes a structurally-weakened portion (or “plow field”)  171  proximate each of the attachment members  170 . The structurally-weakened portions  171  may be adapted to provide the necessary strength and rigidity characteristics for the nominal loads experienced during normal flight and normal landing operations. The structurally-weakened portions  171  may be further adapted, however, to intentionally “fail,” deform, bend, or otherwise “give way” under certain non-nominal load conditions, such as those that may be experienced during an abnormally hard landing. In one particular embodiment, for example, the structurally-weakened portions  171  are adapted to fail at a design point that is selected based on the load limit of the composite leg  162 . In other words, the structurally-weakened portions  171  may be adapted to fail prior to a failure point of the composite leg  162 , thereby absorbing some of the energy of impact associated with an abnormally hard landing, and possibly preventing breakage of the composite leg  162 . 
       FIG. 4  is an enlarged, partial isometric view of a portion of the aircraft  100  of  FIG. 1 . In this embodiment, the fuselage  110  of the aircraft  100  includes a plurality of fuselage components  112  that are coupled to a frame assembly  120 . The frame assembly  120  includes a plurality of frame members  122 . In one embodiment, at least some of the fuselage components  112  and frame members  122  are formed of composite materials. 
     In some embodiments, the landing gear assembly  160  may be moved between a landing (or deployed) position  176  and a flight (or stowed) position  178 . For example,  FIG. 5  is an isometric, partially-exploded view of a portion of the landing gear assembly  160  of  FIG. 2 . More specifically,  FIG. 5  is a view of one of the box-like receiving members  168  and one of the frame members  122  to which it is normally coupled (shown separated from the frame member  122  in the exploded view in  FIG. 5 ). As shown in  FIGS. 3 and 5 , in this embodiment, each receiving member  168  is pivotably coupled to one of the frame members  122  by a pair of bearings  180 . The bearings  180  enable the receiving members  168 , and thus the composite legs  162 , to be moved between the landing position  176  and the flight position  178  ( FIG. 4 ) by a suitable actuator (not shown). 
     Reference is now made to  FIG. 6 , which is an isometric, cross-sectional view of the composite leg  162  of  FIG. 2 . The composite leg  162  includes a pair of sidewalls  140 , and upper wall  142 , and a lower wall  144 . In this embodiment, the composite leg  162  has an approximately square cross-sectional shape. In other embodiments, the cross-sectional shape may be rectangular, elliptical, tubular, etc. 
       FIG. 6  shows a leg  162  that is hollow. In other embodiments, however, the leg may be solid, it may be filled (e.g., with foam, a honeycomb core, or other filler), etc. 
     A reference plane  134  bisects the composite leg  162 , and a longitudinal axis  136  extends along a centerline of the composite leg  162 . When the leg  162  is in the loaded position (e.g. during landing as shown in  FIG. 2 ), an upper portion  130  of the leg experiences compression loads, a lower portion  132  of the leg  162  experiences tension loads, and the sidewalls  140  experience shear loads. 
     Consequently, the upper wall  142  and the upper portions of the sidewalls  140  are under compression, while the lower wall  144  and the lower portions of the sidewalls  140  are under tension. The upper and lower portions  130 ,  132  are not necessarily separated by the reference plane  134 . For example, in the embodiment shown in  FIG. 6 , the lower portion  132  extends from below the reference plane  134  to slightly above the reference plane  134 . A “compression cap” is formed by the upper wall  142  and those portions of the sidewalls  140  above the plane  134 . A “tension cap” is formed by the lower wall  144  and those portions of the sidewalls  140  below the plane  134 . 
     The compression cap includes plies  146  of reinforcing fibers  148  that provide strength in compression. The tension cap includes plies  150  of reinforcing fibers  152  that provide strength in tension. The fibers  148  and  152  in the caps extend lengthwise along the leg  162 . 
     In some embodiments, the cap fibers  148  and  152  may be oriented at zero degrees with respect to the longitudinal axis  136 . Thus, compression and tension loads will be transmitted primarily along the fibers  148  and  152 . 
     In other embodiments, the cap fibers  148  and  152  are oriented at +α and −α with respect to the longitudinal axis  136 . For the cap fibers  148  and  152  in these embodiments, a value of α between 2 and 8 degrees will suppress or delay ply splitting. Fibers oriented at angles of +α and −α are described in greater detail in pending U.S. Ser. No. 12/340,631, which is incorporated herein by reference. 
     The compression cap may have a greater number of plies  146  than the tension cap, making it thicker than the tension cap. Making the compression cap 20-40% thicker than the tension cap can result in as much as a 3% improvement in strength-to-weight ratio of the leg  162 . 
     The composite leg  162  further includes a torque box  154 . The torque box  154  is formed on or in the sidewalls  140  and the upper and lower walls  142  and  144 . The torque box  154  includes plies of carbon reinforcing fibers  156  oriented at angles of +ν and −ν with respect to the longitudinal axis  136  of the leg  136  (the carbon fibers  156  in a ply may be cross-woven). The value of ν is between 35 and 55 degrees. The carbon fibers  156  provide torsional rigidity and carry shear flow between the caps. For example, the torque box  154  may be designed to provide sufficient torsional rigidity to substantially prevent wobble in order to allow the landing wheels  164  to track properly during landing of the aircraft  100 . 
     Because the carbon fibers are oriented at 35≦ν≦55, they do not see the same strain as the cap fibers  148  and  152 . Instead, the carbon fibers  156  tend to “scissor” while the cap fibers are undergoing large deflections. This mechanism allows the carbon fibers  156  to survive while the leg  162  is undergoing large deflections (e.g., 20,000 microinches). 
     In some embodiments, the cap fibers  148  and  152  are formed of a glass material (GS) such as fiberglass, while the torque box  156  are formed from carbon (CB). In other embodiments, the compression cap fibers  148  may be formed of a glass material (GS) while the tension cap fibers  152  are formed from a plastic (or organic) material (P). Examples of plastic fibers  152  include, but are not limited to, aramid fibers, polyethylene fibers (e.g., high molecular weight polyethylene fiber) and other suitable high-strength plastic fibers (e.g., p-phenylene benzobisoxazole fibers, polyarylate fibers). Plastic fibers are used in the tension cap only because they have poor compression values. A leg  162  having plastic fibers in the compression cap has a lighter weight than a leg having a glass compression cap. However, a leg  162  having both caps made of glass is less expensive. 
       FIG. 7  is a representative graph  200  of stress versus strain for carbon (CB), a glass (GS 1 ), and a plastic (P 1 ) material. Similarly,  FIG. 8  is a representative table  300  of various material properties for various fiber materials, including carbon (CB), a glass (GS 1 ), and a plastic (P 1 ). The area under the stress versus strain curve for each material represents the amount of energy that may be absorbed by the material during an impact, such as during landing. The glass material (GS 1 ) can absorb more energy than carbon (CB), and the plastic material (P 1 ) can absorb more energy than both the glass material (GS 1 ) and carbon (CB). Furthermore, as shown in  FIG. 8 , the glass material is more dense (i.e. heavier) than carbon, while the plastic material (P 1 ) is less dense (i.e. lighter) than both carbon (CB) and the glass material (GS 1 ). By proper selection of materials for the cap fibers  148  and  152 , the compression cap may be designed to exhibit desired compression strength characteristics (e.g. a desired ε c ), and the tension cap may be designed to exhibit desired tensile strength characteristics (e.g. a desired ε t ). 
     In one particular embodiment, fiberglass unidirectional tape (or glass planks) (e.g. E-glass) with a modulus of approximately 5 MSI may be used to form the compression and tension caps of the composite leg  162  to achieve very high deflections required for energy absorption. Similarly, graphite (GR) bidirectional weave (e.g. T-300 at ±45 degrees) may be placed in an interleaved fashion between the glass plies. The fibers of the graphite weave transfer shear, provide torsional rigidity (for improved ground handling stability of the aircraft  100 ), and prevent transverse cracking of the glass and to improve transition of the loads into the attachment members  170 . 
     The resulting embodiment of the composite leg  162  may flex a substantial amount. Referring to the equation above for determining radius of curvature p, strains are driven up to make the radius of curvature smaller and deflections larger so more energy can be absorbed. In one particular embodiment, strain levels exceeding 20,000 micro-inch may be achieved. Consequently, a landing gear assembly herein can absorb large impacts upon landing via large leg deflections as compared to traditional landing assemblies, which use articulated structures. 
     Embodiments herein may provide other significant advantages over traditional landing gear assemblies. Traditional landing gear assemblies typically use steel forgings, which require relatively expensive tooling, and which make the assemblies undesirably heavy. A traditional landing gear assembly typically weighs between about 5 to 10 percent of max takeoff weight. Embodiments herein, however, may weigh between about 2 to 5 percent of max takeoff weight, a weight savings of approximately 50% and avoid the cost of expensive forging tools. 
     Moreover, the use of high elongation fibers in the tension cap, high compression fibers in the compression cap, and relatively-stiff carbon fibers in the torque box, the composite legs  162  may provide significantly enhanced structural performance in comparison with traditional landing assemblies. The composite leg  162  described herein may flex a substantial amount to absorb energy during abnormal landing conditions, yet may still maintain highly desirable strength and weight characteristics. Furthermore, because the receiving member  168  includes the structurally-weakened portions  171 , the receiving member  168  is further adapted to absorb additional energy during landing of the aircraft  100 . In this way, embodiments of the present invention may advantageously prevent damage to other portions of the aircraft  100  during landing. 
     In the event that an abnormal landing condition causes damage to a receiving member  168  (e.g. by causing the attachment members  170  to deform the “plow fields”  171 ), the receiving member  168  may be easily replaced. The closed cell (or box beam) construction of the composite leg  162  suppresses free edge effects (e.g. delamination) that might otherwise occur in alternate designs due to abnormal loading conditions. Also, because the composite leg  162  is received within and attached to the retaining member  168  using a relatively simple design, repair and replacement of composite legs  162  that have been damaged is simplified. Therefore, a landing gear assembly herein may improve the maintainability and overall cost of operating the aircraft  100  in comparison with traditional landing gear assemblies. 
     Further improvement of the composite leg  162  can be obtained by additional plies of carbon fibers that form pads around the attachment members  170 . The pads protect the composite leg  162  from damage during deformation of the attachment members  170 . The pads also reduce damage during fabrication, by providing a thicker structure through which to drill holes for the attachment member  170 .