Patent Publication Number: US-2022212815-A1

Title: Aircraft assembly

Description:
FIELD OF THE INVENTION 
     The present invention relates to an aircraft assembly. The present invention also relates to an aircraft structural component, an aircraft, and a method of assembling an aircraft assembly. 
     BACKGROUND OF THE INVENTION 
     During a process of assembling an aircraft assembly, structural components are brought together and fastened to each other with fasteners. One such process involves holding the components relative to each other in one or more jigs, undertaking a machining operation to drill holes in the components to receive fasteners, and fastening fasteners through the holes to mount the components to each other. 
     Following formation of the holes by drilling, the components need to be disassembled to allow for a deburring process to be undertaken. This allows for the removal of chaff from around the holes generated by the machining operation. The components are then repositioned and the fasteners inserted through the aligned holes in a fastening operation. 
     One such aircraft assembly is a wing box assembly which includes upper and lower covers with other components received between the upper and lower covers, such as spars, ribs and landing gear ribs. Removing the covers to enable the deburring operation to take place is an inefficient and time consuming process. 
     It has been recognised that it is unnecessary to undertake a deburring operation for components formed from some materials used in aerospace applications, for example aluminium and carbon fibre reinforced plastic. However, it is also recognised that the use of such materials in some applications is unsuitable, for example when a high loading capability is required. 
     SUMMARY OF THE INVENTION 
     According to an aspect of the invention, there is provided an aircraft assembly comprising: a first structural component; a second structural component; a fastener fastening the first component to the second component; wherein the first structural component comprises a body and an insert in the body, the insert having a machined hole through which the fastener extends; and wherein the material hardness of the insert is less than the material hardness of the body. 
     With such an arrangement, a machining process to form a hole in the first component during an assembly process will act on a material having a lower hardness value than the remainder of the component. As such, the wear on the tool, for example a drill bit or grinding tool is minimised. Furthermore, the need to provide a deburring process may be reduced or eliminated. As such, the need to separate the components during the assembly process following a machining operation and prior to a fastening operation is removed. 
     The likelihood of swarf formed by the machining process acting on adjacent parts is minimized. 
     The machining process to form a hole may be a drilling process. The machining process to form a hole may be a grinding process. 
     The material hardness of the second structural component adjacent to the insert may at least substantially correspond to the material hardness of the insert. 
     With such an arrangement, the ease of forming a hole through both structural components is maximized. 
     The second structural component may abut the insert. 
     The body and the insert may form a one piece component. The insert may be mechanically fixed in the body. The insert may be formed from a cured resin. The insert may be cured in the body during manufacture. 
     The insert may be a plug of material. 
     The insert may have a maximum material hardness of 200 Vickers Hardness (HV), preferably a maximum material hardness of 175 HV, and more preferably a maximum material hardness of 125 HV. However, it will be understood that the material hardness of the insert is dependent on the material hardness of the body. The material hardness of the insert may also be dependent on the material hardness of one or more adjacent components in the stack of components. 
     The insert may be formed from at least one of aluminium and carbon fibre reinforced plastic. 
     The body may be formed from one of steel and titanium. 
     The insert may extend through the body. 
     The insert may be in an interference fit with the body. The insert may be, for example, welded in the body, bonded in the body, cured in the body, cold worked in the body, or press fit in the body. Accordingly, the load transfer path between the insert and the body is enhanced in a shear load direction. 
     The insert may comprise a lip, the lip being engaged in the body to retain the insert in an axial direction of the fastener. Accordingly, the load transfer path between the insert and the body is maximized in a pull through load direction. 
     A portion of the insert may be retained between the body and the second component. 
     The aircraft assembly may comprise a key configuration between the insert and the body which is configured to prevent rotation of the insert relative to the body about an axis of the fastener. 
     The insert may have a central axis. The machined hole may be offset from the central axis. 
     The insert may be one of an array of inserts in the body. 
     The fastener may be one of a plurality of fasteners. At least one of a plurality of fasteners may extend through each of the inserts. 
     The fastener may comprise a blind fastener. 
     The aircraft assembly may be a landing gear assembly. 
     According to another aspect of the invention, there is provided an aircraft assembly comprising: a first structural component; a second structural component; a fastener fastening the first component to the second component; wherein the first structural component comprises a body and an insert in the body, the insert having a machined hole through which the fastener extends; and wherein at least one of the material hardness, the material toughness, the material abrasiveness and the material ductility of the insert is less than the corresponding material hardness, material toughness, material abrasiveness and material ductility of the body. 
     According to another aspect of the present invention, there is provided an aircraft structural component for assembly in an aircraft assembly, the structural component comprising: a body; an insert in the structural component; wherein the insert is arranged to be bored to form a fastener receiving hole during assembly of the component with another component; and wherein the material hardness of the insert is less than the material hardness of the body. 
     The insert may be a solid portion. 
     The insert may be a disc. 
     The insert may be fixed in the body. 
     The insert may be one of an array of inserts, wherein each of the array of inserts corresponds to a component mounting point. 
     According to another aspect of the present invention, there is provided an aircraft comprising at least one of the aircraft assembly as set out above and the aircraft component as set out above. 
     According to another aspect of the present invention, there is provided a method of assembling an aircraft assembly, the method comprising: providing first and second aircraft structural components, the first aircraft component comprising a body with an insert wherein the material hardness of the insert is less than the material hardness of the body; aligning the first component with the second component; forming a hole in the insert; and inserting a fastener through the hole in the insert to fasten the first component with the second component. 
     According to another aspect of the present invention, there is provided a method of assembling an aircraft assembly, the method comprising: providing first and second aircraft structural components, the first aircraft component comprising a body with an insert wherein at least one of the material hardness, the material toughness, the material abrasiveness and the material ductility of the insert is less than the corresponding material hardness, material toughness, material abrasiveness and material ductility of the body; aligning the first component with the second component; forming a hole in the insert; and inserting a fastener through the hole in the insert to fasten the first component with the second component. 
     The method may comprise, following forming the hole in the insert, without moving the first and second components apart, inserting the fastener to fasten the first and second components together. 
     The method may comprise providing the first aircraft component with the material hardness of the insert substantially corresponding with the material hardness of the first aircraft structural component. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Embodiments of the invention will now be described with reference to the accompanying drawings, in which: 
         FIG. 1  is a perspective view of an aircraft; 
         FIG. 2  is a perspective view of part of a wing box assembly including a rear spar, and a landing gear rib with upper and lower covers omitted from view; 
         FIG. 3  is a cross sectional schematic view of a first component of an aircraft assembly prior to assembly; 
         FIG. 4  is a cross sectional schematic view of aligned first and second components of the aircraft assembly at a first stage of the assembly process; 
         FIG. 5  is a cross sectional schematic view of the first and second components shown in  FIG. 4  at a second stage of an assembly process in which the first component is shown partially machined by a bit to form a machined hole; 
         FIG. 6  is a cross sectional schematic view of the first and second components shown in  FIG. 5  following the machining process; 
         FIG. 7  is a cross sectional schematic view of the first and second components of  FIG. 6  at a third stage of the assembly process in which a fastener is received through the fastening bore formed through both the first and second components; 
         FIG. 8  is a cross sectional schematic view of first and second components of an aircraft assembly during assembly following the machining of the fastening bore through the first and second components with another insert embodiment; and 
         FIG. 9  is a plan view of another embodiment of the first component of the aircraft assembly. 
     
    
    
     DETAILED DESCRIPTION OF EMBODIMENT(S) 
       FIG. 1  shows an aircraft  10 . The aircraft  10  has a fuselage  12 , and starboard and port fixed wings  13 ,  14 . An engine  15  is mounted to each wing  13 ,  14 . The aircraft  10  is a typical jet passenger transport aircraft but the invention is applicable to a wide variety of fixed wing aircraft types, including commercial, military, passenger, cargo, jet, propeller, general aviation, etc. with any number of engines attached to the wings or fuselage. The invention is also applicable to other aircraft, such as helicopters. 
     Each wing has a cantilevered structure with a length extending in a span-wise direction from a root  18  to a tip  19 , with the root  18  being joined to the aircraft fuselage  12 . The wings  13 ,  14  are similar in construction and so only the starboard wing  13  will be described in detail. The wing  13  has a leading edge  16  and a trailing edge  17 . The leading edge  16  is at the forward end of the wing and the trailing edge  17  is at the rearward end of the wing. 
     The wing  13  comprises a wing box  20 . The wing box  20  forms a structural assembly including forward and rear spars (part of the rear spar shown in  FIG. 2 ), ribs extending between the forward and rear spars, upper and lower covers,  21 ,  22 , and other components. 
     The wing  13  has a span-wise axis which extends in a direction from the wing root  18  to the wing tip  19 , and a chord-wise axis which extends in the direction from the leading edge  16  to the trailing edge  17 . 
     The aircraft  10  has landing gear assemblies (not shown). A starboard landing gear is selectively extendable from the starboard wing  13 , a port landing gear is selectively extendable from the port wing  14 , and a nose landing gear is selectively extendable from the fuselage  12 . The starboard and port landing gears are mounted on the wing boxes  20  of the wings  13 ,  14 . 
     Referring to  FIG. 2 , a section of the wing box  20  is shown. The section of the wing box  20  shown includes part of a rear spar  23 . A landing gear rib  24 , also known as a gear rib, is mounted on the rear spar  23 . The gear rib  24  acts as part of the mount for the landing gear assembly. The gear rib  24  is fixedly mounted to the rear spar  23 . 
     The upper and lower covers,  21 ,  22  are omitted from view in  FIG. 2 . The upper cover is positioned on the upper side of the rear spar  23  and the lower cover is positioned on the lower side of the rear spar  23 . The gear rib  24  extends between the upper and lower covers. The gear rib  24  is mounted to the upper cover and the lower cover when the wing box  20  is assembled. 
     The gear rib  24  includes a body  25 . The body  25  includes an array of component mounting points  26 . The component mounting points  26  enable other components to be fastened with the gear rib  24 . The body  25  includes an upper cover mounting flange  27  and a lower cover mounting flange  28 . Component mounting points  26  are formed in each of the upper and lower cover mounting flanges  27 ,  28 . 
     As described herein, the gear rib  24  acts as a first component of an aircraft assembly. The present invention is described herein with reference to mounting the gear rib  24  with each of the upper and lower covers, each acting as a second component of the aircraft assembly, however it will be understood that each of the first and second components may be different components, and the arrangement of the aircraft assembly may differ. 
     As will become apparent hereinafter, the gear rib  24  is shown part way through an assembly process in which the upper and lower covers  21 ,  22  have already been positioned with respect to the gear rib  24  (although the upper and lower covers are omitted from view for clarity in  FIG. 2 ) and with fastening bores  29  formed through the upper and lower covers and the gear rib  24  but prior to fasteners being inserted. 
     Referring now to  FIGS. 3 to 7 , a process for assembling an aircraft assembly  30 , for example the wing box  20 , will now be described. The assembly process will be described with reference to first and second components  40 ,  50 . The first and second components  40 ,  50  are described above as a gear rib and a cover respectively, however it will be understood that the first and second components and the assembly process may relate to alternative components of an aircraft. Furthermore, it will be understood that the assembly process may be applied to more than two components, for example three components having parts in a stacked configuration. 
     A body  41  of the first component  40  is shown schematically in  FIG. 3 . The body  41  may include a flange. The first component  40  is formed as a one piece component. The body  41  has a first side  42  and a second side  43 . Although the first and second sides  42 ,  43  are shown parallel to each other, it will be understood that they may be formed at an incline to each other. 
     An insert  44  is in the body  41 . The insert  44  is accommodated extending across the body  41 . The insert  44  forms an interference fit with the body  41 . The insert  44  may be in the flange. It will be understood the insert may be accommodated in the body  41  in different configurations. The fit between the insert  44  and the body  41  is sufficient to allow for a seamless load transfer between the insert  44  and the body  41  in a shear load direction. The insert  44  and the body  41  are pre-assembled. The insert  44  is pre-formed with the body  41 . 
     The insert  44  is a solid part. That is, the insert  44  is formed without one or more holes extending through the insert through which a fastener may be received. The insert  44  is a disc in an aperture  45  in the body  41 . The insert  44  is cylindrical, however it will be understood that the insert  44  may have alternative configurations. For example, the insert  44  may have a non-circular cross-section and may have one or more protrusions and/or recesses formed in the insert  44 . 
     The body  41  of the first component  40  is formed from a titanium alloy. Titanium alloys typically have a material hardness of at least 300 HV, although some alloys, for example dependent on treatment, may have a lower hardness. Alternative materials may be used. For example, the body  41  of the first component  40  may be formed from steel. The material hardness of the material forming the body  41  of the first component  40  has a material hardness value of at least 200 HV. Such materials typically require deburring following the machining of a hole through the material, for example through use of a drill bit or grinding tool. 
     The insert  44  is formed from a different material to the body  41 . The insert  44  is formed from aluminium. The insert  44  may be formed from an alternative material such as carbon fibre reinforced plastic (CFRP). The material forming the insert  44  is a softer material than the material forming the body  41 . That is, the material hardness of the insert  44  is lower than the material hardness of the body  41 . The material forming the insert  44  has a material hardness of less than 200 HV. However, it will be understood that this is dependent on the relative material hardness of the body  41 . That is, the material hardness of the insert  44  is less than the material hardness of the body  41 . The insert  44  has sufficient outer dimensions to accommodate a hole for receiving a fastener therethrough. The size of the hole required to be formed through the insert should be sufficient to accommodate the required fastener for fastening the components  40 ,  50  at the component mounting point  26 . The insert  44  is at a predetermined one of the component mounting points  26 . The insert  44  is configured to be sized to accommodate any tolerance build up at the component mounting point  26  as predetermined for the assembly of the aircraft assembly  30 . 
     Hardness is described herein by reference to Vickers hardness (HV) as a measure of material hardness, although it will be understood that other methods are used to determine material hardness. Examples of Vickers hardness values are provided below: 
     
       
         
           
               
               
               
             
               
                   
                   
               
               
                   
                 Material 
                 Vickers Hardness 
               
               
                   
                   
               
             
            
               
                   
                 Ti-6Al-2Sn-4Zr-2Mo (Ti-6-2-4-2), Sheet 
                 333 
               
               
                   
                 Titanium Ti-6Al-4V (Grade 5), Annealed 
                 349 
               
               
                   
                 Titanium Ti-6Al-4V (Grade 5), STA 
                 396 
               
               
                   
                 Ti-15V-3Cr-3Al-3Sn Solution Treated 
                 222 
               
               
                   
                 Steel S99 Forging 
                 286 
               
               
                   
                 Aluminium 2014-T451 
                 118 
               
               
                   
                 Aluminium 2014-T651 
                 155 
               
               
                   
                 Aluminium 7050-T7651 
                 171 
               
               
                   
                 Glass Fibre Reinforced Plastic 
                 62-74  
               
               
                   
                 CFRP 
                 80-100 
               
               
                   
                   
               
            
           
         
       
     
     Referring to  FIG. 4 , the first component  40  is aligned with the second component  50 . The second component  50  is moved into abutment with the first side  42  of the body  41 . The first and second components  40 ,  50  are aligned to be fastener together in a predetermined stacked arrangement. 
     The second component  50  includes a body  51 . The body  51  may form the whole or part of the second component  50 . The body  51  may include a flange. The second component  50  is formed from carbon fibre reinforced plastic. It will be understood that the second component  50  may be formed from an alternative material such as aluminium, titanium, or steel. 
     In the present configuration, the second component  50  is shown with a pre-formed hole  52 . The pre-formed hole  52  extends through the body  51 . The hole  52  may be preformed prior to bringing the first and second components  40 ,  50  together. The hole  52  may be formed during the assembly process. It will be recognised that in an embodiment in which the second component is formed from a material having a material hardness substantially corresponding to that of the insert then any hole formed during the assembly process can be formed without a requirement for a subsequent deburring operation. 
     The hole  52  is aligned with the insert  44 . That is, the hole  52  fully overlaps the insert  44 . The hole  52  does not overlap the body  41 . In an arrangement in which the hole  52  is formed during the assembly process, then the position of the hole is pre-defined as a component mounting point  26 . The insert  44  is comparatively sized with the preformed hole  52  to accommodate any pre-determined tolerance build ups during assembly of the components  40 ,  50 . 
     The preformed hole  52  has a second component hole axis  53 . It will be noted that the second component hole axis  53  is offset from a central axis  46  of the insert  44 . In the event of no misalignment or tolerance build-up, then the second component hole axis  53  and central axis  46  of the insert  44  may be coaxial. 
     Upon alignment of the first and second components  40 ,  50  in an arrangement for assembly, a machine operation is performed. The machine operation bores a hole. A drill bit  60  is used to bore a through hole  47  in the insert  44 . The drill bit  60  is a boring tool. A grinding tool may be used to bore the through hole  47 . The drill bit  60  is aligned at the component mounting point  26 . In an embodiment in which the hole  52  in the second component  50  is preformed, then the drill bit  60  may be aligned with the axis  53  of the preformed hole  52 . Alternatively, the component mounting point  26  is determined and the drill bit  60  is used to form the hole through both the first and second components  40 ,  50 . In  FIG. 5 , the drill bit  60  is shown during the machining operation partially engaged with the insert  44 . The drill bit  60  is acting in a direction through the second component  50  and into the first component  40 . In embodiments, the opposite direction may be used. 
     The machining operation forming the machined hole ensures alignment of the holes  52 ,  47  through both the first and second components  40 ,  50 . The holes  47 ,  52  form a fastening bore  48 . The axis  53  of the hole  52  in the second component is therefore coaxial with the axis of the through hole  47  in the first component  40 . The through hole  47  is formed fully through the insert  44 . The insert  44  forms a collar around the through hole  47 . 
     Once the machining operation is complete, a fastening operation is performed. A fastener  70  is inserted through the fastening bore  48 . The fastener  70  is fastened in an engaged position to mount the first and second components  40 ,  50  with each other. It will be recognised that following the machining operation there is no need to deburr either of the first or second components  40 ,  50 , in particular as the machining process acts on a softer material. The material hardness of the insert is less than the corresponding material hardness of the material surrounding the insert. 
     It will be understood that other material properties may contribute to aid the machining operation. For example, in embodiments at least one of the material toughness, the material abrasiveness and the material ductility of the insert is less than the corresponding material toughness, material abrasiveness and material ductility of the body. 
     The fastener  70  is shown as a bolt  71  and a nut  72  arrangement. However, it will be appreciated that the fastener  70  may be a blind fastener. That is a fastener that is inserted through the fastening bore  48  and engaged with both of the first and second components  40 ,  50  from one side of the assembly only. An advantage of this arrangement is that the machining operation and the fastening operation may be performed from the second component side of the assembly  30  only. 
     The interference fit between the insert  44  and the body  41  provides for shear loads to be sufficiently transferred between the first component  40  and the fastener  70  to the second component  50 . In  FIG. 7 , the first component side of the fastener is shown in contact with the insert  44  only, however it will be understood that the end  73  of the fastener  70  may be configured to extend over at least part of the body  41 . Such a configuration would aid the transfer of a pull through load on the first component  40 . 
     Another embodiment is shown in  FIG. 8 . The embodiment in  FIG. 8  is generally the same as described above and the assembly process is generally the same and so a detailed description will be omitted herein. However, in this embodiment the configuration of the insert differs.  FIG. 8  shows a partially assembled aircraft assembly  30  with first and second components  40 ,  50 . The partially formed aircraft assembly  30  is shown following the machining operation and prior to the fastening operation. As such, a through hole  87  is formed through an insert  80 . The insert  80  is received in the body  51  of the first component  40 . The insert  80  is generally the same as the insert  44  described above, however in this embodiment the insert  80  includes a lip  88 . The lip  88  is a circumferentially extending flange. The lip  88  may have a different configuration, and extend only partially around the insert  80 . The lip  88  protrudes outwardly. The lip  88  is a protrusion. The lip  88  is received on a shoulder  49  of the body  41 . The lip  88  is received between the shoulder  49  and the first side  42  of the body  41 . The lip  88  aids retention of the insert  44  in the body  41 . When assembled, the lip  88  is received between the shoulder  49  of the body  41  of the first component  40  and the second component  50 . As such, the insert  44  is able to handle greater pull through loads acting on the aircraft assembly  30 . In embodiments the lip  88  is a countersink. 
     Referring to  FIG. 9 , another embodiment is shown. The arrangement of this embodiment is generally the same as the embodiments shown above.  FIG. 9  shows the first component  40  prior to assembly with the second component  50 , and prior to the machining operation. As such, no hole is formed through the insert. An insert  90  is shown in the body  41  of the first component  40 . The insert  90  has a key configuration  91 . The key configuration includes a key  93  and a key slot  92 . The key  93  protrudes from a main part of the insert. The key  93  protrudes radially outwardly in the present embodiment. The key  93  is received in a corresponding key slot  92  in the body  41 . The key  93  may have differing configurations and may comprise two or more key features. The key configuration  91  aids prevention of any relative rotation of the insert  90  and the body  41 , for example, such as may be applied during the machining process. 
     In each of the embodiments described above, it will be appreciated that the insert and the body together form the first component  40  as a one piece component. The first component  40  includes a plurality of inserts preassembled with the body  41 . The location of the insert  90  corresponds to the position of predetermined component mounting points  26 . The inserts are preformed without any through holes formed therein through which fasteners may be engaged, and therefore the fastener receiving holes are formed during the assembly process. It has been recognised that by using a relatively softer material than that of the body of the component, that it is possibly to remove the need for subsequent machining processes following the forming of the hole in the insert and therefore reducing the assembly time. It will be recognised that in some embodiments two or more through holes arranged to receive fasteners may be formed in a single insert. 
     In the embodiment shown in  FIG. 2  in which the first component is a landing gear rib  24  and the second component is one of the covers  21 ,  22 , it will be appreciated that a component that is required to carry a significant load transfer may lead to the cover having to be removed in order to deburr holes machined in the component. However, with the arrangements described above it has been recognised that inserts may be used to allow the holes to be formed in a relatively softer material to remove the further machining requirement and so remove the need to remove the cover. As such, the assembly time and complexity of the assembly process may be reduced. Furthermore, as the tools, for example the drill bits used during the assembly process are required to act on a softer material hardness only, then the wear on these tools is minimised. 
     Where the word ‘or’ appears this is to be construed to mean ‘and/or’ such that items referred to are not necessarily mutually exclusive and may be used in any appropriate combination. 
     Although the invention has been described above with reference to one or more preferred embodiments, it will be appreciated that various changes or modifications may be made without departing from the scope of the invention as defined in the appended claims.