Patent Publication Number: US-2022212777-A1

Title: Composite spars with integrated sacrificial surfaces

Description:
RELATED APPLICATIONS 
     This non-provisional patent application claims priority to U.S. Provisional Patent Application No. 63/134,071 filed on Jan. 5, 2021, which is incorporated by reference as if fully provided herein. 
    
    
     FIELD 
     This disclosure relates to the field of aircraft structures, and in particular, to aircraft structures that utilize spars as structural members. 
     BACKGROUND 
     During a wing assembly process for an aircraft, ribs and spars are assembled to form a skeleton for the wing. The spars typically form the main structural member of the wing, and run along a length of the wing. The ribs are attached to the spars (e.g., the ribs may be attached to a front spar at a leading edge of the wing and a rear spar at a trailing edge of the wing), and the ribs generally have a perimeter shape that defines the airfoil for the wing. The outer surface of the wing is formed by skin panels that are attached along the perimeter of the ribs and the perimeter of the spars to form a smooth surface for the wing. 
     In order to ensure a proper fit of the skin panels to the spars, shims may be installed at the perimeter of the spars where the inside surfaces of the skin panels contact the spars. The shims are used in order to preclude a possible machining process on the spars themselves, which may negatively impact the structural integrity of the spars. 
     Typically, installing the shims onto the spars is time consuming manual process, which entails additional man-hours of assembly time and disassembly time for a wing, or other aircraft structures that utilize spars, to shim gaps between skin panels and the spars. 
     Based on the forgoing discussion, it therefore remains desirable to improve the fabrication process for aircraft, and in particular, to improve processes for assembling aircraft structures that include spars. 
     SUMMARY 
     Composite assemblies are described that include composite spars that are co-cured with one or more sacrificial members on their flanges, forming an integrated sacrificial surface for the composite spars. Generally, a spar is an elongated structure that includes a web defining a major surface of the spar and flanges that project from sides of the web. Spars run spanwise in a wing, at right angles or nearly right angles to the fuselage, and form the main structural components of a wing. The web in the spar is vertically disposed in the wing, and the flanges define surfaces for attachment of the skin panels. During assembly, gaps may exist between the flanges and the skin panels. In the embodiments described herein, a sacrificial member co-cured with a flange is machined to bring the outer surface of the sacrificial member into conformance with surfaces of the skin panel(s), thereby providing a technical benefit of mitigating gaps between spars and skin panels in the wing or other aircraft structures that utilize spars. 
     One embodiment comprises a composite assembly. The composite assembly includes a composite spar having a web and flanges that project from sides of the web. The composite assembly further includes a sacrificial member of composite materials co-cured with the composite spar on an outer surface of at least one of the flanges. In addition, the sacrificial member has an outer surface that has been machined into conformance with an inner surface of at least one skin panel for an aircraft structure to form a contact surface with the at least one skin panel. 
     Another embodiment comprises a method of fabricating a composite assembly. The method comprises assembling a first composite layup that defines a web portion and flange portions for a composite spar of the composite assembly, assembling a second composite layup onto at least one of the flange portions to define a sacrificial member for the composite assembly, and co-curing the first composite layup and the second composite layup to harden the composite assembly. The method further comprises machining an outer surface of the sacrificial member into conformance with an inner surface of at least one skin panel for an aircraft structure to form a contact surface for the at least one skin panel. 
     Another embodiment comprises a method of fabricating a composite assembly. The method comprises performing a first composite layup onto a layup mandrel that defines a contour for a composite spar of the composite assembly, performing a second composite layup onto a flange portion of the first composite layup defined by contour, and co-curing the first composite layup and the second composite layup to harden the composite assembly, wherein the first composite layup forms the composite spar and the second composite layup forms a sacrificial member on a flange on the composite spar. The method further comprises calculating a machining depth for the sacrificial member based on estimated spacing tolerances between the flange of the composite spar and at least one skin panel for an aircraft structure, and machining an outer surface of the sacrificial member along at least a portion of a length of the composite spar based on the machining depth to form a contact surface for the at least one skin panel. 
     The features, functions, and advantages that have been discussed can be achieved independently in various embodiments or may be combined in yet other embodiments, further details of which can be seen with reference to the following description and drawings. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Some embodiments are now described, by way of example only, and with reference to the accompanying drawings. The same reference number represents the same element or the same type of element on all drawings. 
         FIG. 1  depicts a wing of an aircraft in the prior art. 
         FIG. 2  is an isometric view of a composite assembly in an illustrative embodiment. 
         FIG. 3  is a cross-sectional view of the composite assembly along cut lines A-A of  FIG. 2  in an illustrative embodiment. 
         FIG. 4  depicts a portion of an aircraft structure in an illustrative embodiment. 
         FIG. 5  is a view of a region in  FIG. 4  in an illustrative embodiment. 
         FIG. 6  is a method of fabricating a composite structure in an illustrative embodiment. 
         FIGS. 7-12  depict additional details of the method of  FIG. 6  in illustrative embodiments. 
         FIGS. 13-24  are isometric views of various stages in the fabrication method of  FIG. 6  in illustrative embodiments. 
         FIG. 25  is a flow chart of another method of fabricating a composite structure in an illustrative embodiment. 
         FIG. 26  depicts additional details of the method of  FIG. 25  in an illustrative embodiment. 
         FIGS. 27-28  are cross-sectional views of the composite assembly along cut lines C-C of  FIG. 2  in an illustrative embodiment. 
         FIG. 29  is a flow chart illustrating an aircraft manufacturing and service method in an illustrative embodiment. 
         FIG. 30  is a schematic diagram of an aircraft in an illustrative embodiment. 
     
    
    
     DETAILED DESCRIPTION 
     The figures and the following description illustrate specific exemplary embodiments. It will be appreciated that those skilled in the art will be able to devise various arrangements that, although not explicitly described or shown herein, embody the principles described herein and are included within the contemplated scope of the claims that follow this description. Furthermore, any examples described herein are intended to aid in understanding the principles of the disclosure are to be construed as being without limitation. As a result, this disclosure is not limited to the specific embodiments or examples described below, but by the claims and their equivalents. 
       FIG. 1  depicts a wing  100  in the prior art. Wing  100  includes a plurality of ribs  102 , which extend between a front spar  104  and a rear spar  105 , with front spar  104  near a leading edge  106  of wing  100  and rear spar  105  towards a trailing edge  107  of wing  100 . In  FIG. 1 , ribs  102  are spaced apart along a length of wing  100 , and their perimeters form a shape of an airfoil for wing  100 . In  FIG. 1 , wing  100  includes nose ribs  108 , which extend from front spar  104  towards leading edge  106  of wing  100  to form the front shape of the airfoil at leading edge  106 . Wing  100  also includes an outer skin  110 , which is partially removed in this view in order to show the underlying structure of wing  100 . 
     As evident in  FIG. 1 , ribs  102  and spars  104 - 105  form the skeleton of wing  100 , with outer skin  110  adopting the shape of ribs  102  along their peripheral surfaces between spars  104 - 105 . 
       FIG. 2  is an isometric view of a composite assembly  200  in an illustrative embodiment. In this embodiment, composite assembly  200  includes a composite spar  202  and one or more sacrificial members  204  of composite materials that are co-cured to form a structural component for an aircraft, such as a wing of an aircraft, an empennage for an aircraft, a tail of an aircraft, etc. 
     Composite parts, such as Carbon Fiber Reinforced Polymer (CFRP) parts and/or Glass Fiber Reinforced Polymer (GFRP) parts, are initially laid-up in multiple layers of material, which may be referred to as reinforcement layers. Individual fibers within each layer are aligned parallel with each other, but different layers exhibit different fiber orientations in order to increase the strength of the resulting composite part along different dimensions. The layers may be pre-impregnated with a matrix material, such as an uncured thermoset resin or a thermoplastic resin, which is referred to as a “prepreg.” Alternatively, the layers may be laid up dry (i.e., “dry fiber”), and are subsequently infused with a matrix material prior to curing. 
     Referring again to  FIG. 2 , composite spar  202  in this embodiment is an elongated member extending along a length  206  generally in the shape of an inverted “U”. Composite spar  202  includes a web  208  that is substantially flat and defines a major surface  210  of composite spar  202 . In this embodiment, flanges  214 - 215  project from sides  212 - 213  of web  208  in a downward direction in  FIG. 2 , defining a width  216  of composite spar  202  and a height  218  of composite spar  202 . 
     In this embodiment, a sacrificial member  204  is co-cured to flange  214  and/or flange  215 . After composite assembly  200  is cured, outer surface  220  of sacrificial member  204  may be machined as desired during a fabrication process to mitigate gaps that may arise between flange  214  and/or flange  215  of composite spar  202  and other aircraft structures, such as skin panels. The use of sacrificial member  204  mitigates the use of manual shims that would typically be introduced when spar flanges are mated to skin panels and gaps exist, thereby providing a technical benefit of reducing the amount of time and effort when assembling various aircraft structures that utilize spars and skin panels. 
     Generally, it is desirable that sacrificial member  204  has a stiffness that is substantially less than the stiffness of composite spar  202 , because removing material from sacrificial member  204  minimizes the changes to the overall stiffness of composite assembly  200 . To achieve that type of result, composite spar  202  may, for example be formed from BMS8-276 tape (a CFRP material) using a quasi-isotropic layup, which has a very high stiffness. Sacrificial member  204  may, for example, be formed from BMS8-276 fabric (another CFRP material), which has a moderate stiffness compared to BMS8-276 tape. The BMS8-276 fabric for sacrificial member  204  may use a varying +/−45 degree orientation during layup in order to minimize the stiffness of composite assembly  200  in the spanwise direction. When sacrificial member  204  is formed from GFRP, which has a low stiffness as compared to either BMS8-276 tape or fabric, various types of materials and orientation of those materials may be implemented. 
     Although sacrificial member  204  is depicted on both flanges  214 - 215  in this embodiment, sacrificial member  204  may be formed on one of flange  214  or flange  215  in some embodiments. 
       FIG. 3  is a cross-sectional view of composite assembly  200  along cut lines A-A of  FIG. 2  in an illustrative embodiment. In this embodiment, composite spar  202  includes a bend  302  between web  208  and flange  214  at side  212  of web  208 , and sacrificial member  204  is disposed on outer surface  304  of flange  214 . Sacrificial member  204  extends from an end  306  of flange  214  towards bend  302 , and terminates at an edge  308  proximate to bend  302 . In this embodiment, edge  308  is a square edge, although in other embodiments, edge  308  may have other shapes, such as a taper or ramp. The use of a taper or ramp for edge  308  of sacrificial member  204  may mitigate the generation of compression waves in composite spar  202  during fabrication, which will be discussed later. 
     Sacrificial member  204  has an initial thickness  310  defined by the distance between outer surface  304  of flange  214  and an outer surface  220  of sacrificial member  204 . Typically, initial thickness  310  is pre-selected based on the expected tolerances between flange  214  and other components on the aircraft, such as skin panels. After co-curing composite assembly  200 , outer surface  220  of sacrificial member  204  may be machined down (e.g., material is removed from outer surface  220  of sacrificial member  204  along length  206  or portions of length  206  of composite spar  202  (see  FIG. 2 )), which reduces initial thickness  310  of sacrificial member  204  to a final thickness  314  based on a machining profile or machining depth. Machining sacrificial member  204  also generates a contact surface  316  for the skin panels. Although layers are not shown in  FIG. 3 , both composite spar  202  and sacrificial member  204  are formed from composite materials, such as CFRP plies. 
     In a similar manner to flange  214 , composite spar  202  includes a bend  303  between web  208  and flange  215  at side  213  of web  208 , and sacrificial member  204  is disposed on an outer surface  305  of flange  215 . Sacrificial member  204  in this embodiment extends from an end  307  of flange  215  towards bend  303 , and terminates at edge  308  proximate to bend  303 . In this embodiment, edge  308  is a square edge, although in other embodiments, edge  308  may have other shapes, such as a taper or ramp, for reasons similar to edge  308  of sacrificial member  204  on flange  214 . Sacrificial member  204  on flange  215  has an initial thickness  310  defined by the distance between outer surface  305  of flange  215  and an outer surface  220  of sacrificial member  204 , which may be the same or different than initial thickness  310  of sacrificial member  204  on flange  214 . Typically, initial thickness  310  is selected based on the expected tolerances between flange  215  and other components on the aircraft, such as skin panels, and then outer surface  220  of sacrificial member  204  may be machined down (e.g., material is removed from outer surface  220  of sacrificial member  204  at flange  215  along length  206  of composite spar  202  (see  FIG. 2 )), which reduces initial thickness  310  of sacrificial member  204  to a final thickness  314  based on a machining profile. In addition, the amount of material removed from sacrificial member  204  may vary along length  206  or a portion of length  206  of composite spar  202  and/or may vary depending upon which of flanges  214 - 215  that sacrificial member  204  is located. 
       FIG. 4  depicts a portion of an aircraft structure  400  in an illustrative embodiment. In this embodiment, aircraft structure  400  includes composite assembly  200  and skin panels  402 - 403 . Sacrificial member  204  for flange  214  is located proximate to an inner surface  404  of skin panel  402  in this embodiment, and sacrificial member  204  of flange  215  is located proximate to an inner surface  405  of skin panel  403  in this embodiment. 
     As discussed previously, shims were often manually installed between spar flanges and skin panels during an assembly process in order to compensate or mitigate gaps that formed between these two components. Generally, it is undesirable to machine directly on composite spar  202 , because the machining process removes fiber layers and may compromise the structural integrity of composite spar  202 . While the use of shims precludes machining directly on spars, the use of shims is a time-consuming process that includes temporarily assembling the components together, measuring any gaps that may exists between spar flanges and their skin panels, disassembling the components, and bonding shims to the spar flanges. The shims may then be machined to achieve the final fit between the spar flanges and the skin panels. A final assembly of the spars and skin panels may then be performed. 
     With sacrificial member  204  co-cured to flange  214  and/or flange  215  of composite spar  202 , a machining process can be performed on outer surface  220  of sacrificial member  204  prior to assembling aircraft structure  400 , which saves time and effort over the prior manual shim process. For example, the machining profile or machining depth for sacrificial member  204  may be generated based on a number of different factors, including an expected tolerance between flanges  214 - 215  of composite spar  202  and inner surfaces  404 - 405  of skin panels  402 - 403 . Once a machining profile or machining depth is selected and composite assembly  200  is tested for fit with respect to skin panels  402 - 403 , subsequent spars with integrated sacrificial members may be machined to the same profile or depth, ensuring that each spar is interchangeable between different builds of the same aircraft structure. If gaps are found between skin panels  402 - 403  and their corresponding sacrificial members  204  after machining and during assembly, the machining profile or machining depth can be adjusted for subsequent fabrications of composite assembly  200  to mitigate the gaps in future builds of aircraft structure  400 . 
     In another example, a 3-Dimensional (3D) scan may be performed on composite assembly  200  and skin panels  402 - 403 , which may then be used to determine the machining profile to apply to sacrificial member(s)  204 . In this case, composite assembly  200  and skin panels  402 - 403  may be serialized or marked for use as a group, which makes it less likely that composite assembly  200  is interchangeable between different build instances of the same aircraft structure  400 . Both of these different types of profile generating processes will be discussed in more detail later. 
       FIG. 5  is a view of region  406  of  FIG. 4  in an illustrative embodiment. In this embodiment, edge  308  of sacrificial member  204  is in the shape of a ramp or taper, which is proximate to bend  303  in composite spar  202 . The use of a taper or ramp for edge  308  may be used to mitigate the deformation in composite spar  202  that may be generated during the fabrication process due to vacuum compression during cure, which may generate waves in the fiber layers that make up composite spar  202 . Edge  308  may be cut into a ramp after layup or laid-up as a ramp during fabrication. When edge  308  is laid-up as ramp, layers may decrease in width as the distance from outer surface  305  of flange  215  increases (e.g., by having layer  502 , which is proximate to outer surface  305  of flange  215 , wider than layer  504 , which is distal to outer surface  305  of flange  215 ).  FIG. 5  also illustrates bonding surface  316  in contact with inner surface  405  of skin panel  403 , with contact surface  316  generated after machining outer surface  220  (see  FIG. 3 ) of sacrificial member  204  until final thickness  314  is achieved. 
       FIG. 6  is a method  600  of fabricating a composite structure in an illustrative embodiment,  FIGS. 7-12  depict additional details of method  600  in illustrative embodiments, and  FIGS. 13-22  are isometric views of various stages in the fabrication process in illustrative embodiments. The methods described herein will be discussed with regard to various embodiments of composite assembly  200  and aircraft structure  400 , although the methods may apply to other configurations of composite assembly  200  and aircraft structure  400 , not shown or described. The steps of the methods described herein may include other steps, not shown. Also, the steps may be performed in an alternate order. 
     Step  602  comprises assembling a first composite layup that defines portions of composite spar  202 . In one example, the first composite layup may be formed flat, and shaped over a layup mandrel  1302  that defines a contour for composite spar  202  (see  FIG. 13 ). Layup mandrel  1302  includes a flat region  1304  that defines the shape of web  208  for composite spar  202 , a first side region  1306  that defines the shape and orientation of flange  214  relative to web  208  for composite spar  202 , and a second side region  1308  that defines the shape and orientation of flange  215  relative to web  208  of composite spar  202 . 
     In another example of assembling the first composite layup, an Automated Fiber Placement (AFP) machine may perform a layup directly on layup mandrel  1302 .  FIG. 14  depicts first composite layup  1402  on layup mandrel  1302 , which defines a web portion  1404 , flange portions  1406 - 1407 , and bends  1408 - 1409  for composite spar  202  defined by the underlying shape or contour of layup mandrel  1302 . Generally, first composite layup  1402  comprises continuous fiber layers that may vary in orientation from one layer to the next, formed as a prepreg or a dry layup. Further, the layers may continue from end  306  of flange  214  to end  307  of flange  215 . As discussed previously, first composite layup  1402  may comprise BMS8-276 tape with a quasi-isotropic layup, which has a very high stiffness after curing. 
     Step  604  comprises assembling a second composite layup  1502  (see  FIG. 15 ) onto flange portion  1406  and/or flange portion  1407  of first composite layup  1402 . For example, a flat layup may be performed for second composite layup  1502 , and transferred to, and directly contacting, flange portion  1406  and/or flange portion  1407  of first composite layup  1402 . In another example, an AFP machine may perform a layup directly on flange portion  1406  and/or flange portion  1407  of first composite layer  1402 . As discussed previously, second composite layup  1502  may comprise BMS8-276 fabric layered in a varying +/−45-degree orientation, which has moderate stiffness after curing as compared to BMS8-276 tape, or a GFRP material, which has a low stiffness after curing as compared to both BMS8-276 tape or BMS8-276 fabric. 
     Step  606  comprises co-curing first composite layup  1402  and second composite layup  1502  to harden composite assembly  200 . For example, first composite layup  1402  and second composite layup  1502  may be bagged, placed under vacuum to apply pressure to first composite layup  1402  and second composite layup  1502 , and heated in order to harden composite assembly  200 . The resulting structure after cure is composite assembly  200  as illustrated in  FIG. 2 , with composite spar  202  having web  208  formed from web portion  1404 , and flanges  214 - 215  formed from flange portions  1406 - 1407 , respectively, of the first composite layup  1402 , and with sacrificial member  204  formed from second composite layup  1502  on one or both of flanges  214 - 215 . 
     Step  608  comprises machining outer surface  220  of sacrificial member  204  into conformance with surfaces of one or more skin panels. For example, machining outer surface  220  of sacrificial member  204  is performed to remove material from sacrificial member  204  and form contact surface  316  for skin panel  402  and/or skin panel  403 . Prior to the machining process, initial thickness  310  of sacrificial member  204  may be between about 0.08 inches and 0.12 inches. After the machining process, final thickness  314  of sacrificial member  204  may be between about 0.03 inches and 0.08 inches. 
     In some cases, it may be desirable that edge  308  of sacrificial member  204  that is proximate to bend  302  and/or bend  303  in composite spar  202  has a specific shape in order to prevent structural changes in composite spar  202  during the fabrication process for composite assembly  200 . In these cases, edges  1504  (see  FIG. 15 ) in second composite layup  1502  may be formed or cut into a shape of a taper or ramp (see step  702  of  FIG. 7 ), and aligned with their respective bends  1408 - 1409  in first composite layup  1402  (see step  704  of  FIG. 7 ). The use of a taper or ramp shape for edges  1504  of second composite layup  1502  prevents waves in first composite layup  1402  when compression is applied during the cure. The ramp or taper may have a particular ratio, such as 1:1 (see step  802  of  FIG. 8 ), where the ramp or taper generally forms a 45-degree angle as depicted in  FIG. 5  after cure. However, other types of shapes for edges  1504  of second composite layup  1502  may be used in other embodiments in order to generate any type of edge  308  that is desired in sacrificial member  204  after cure. 
     When a hand placement of second composite layup  1502  onto first composite layup  1402  is performed prior to cure, second composite layup  1502  may be assembled onto an adhesive film  1602  placed onto a work surface  1604  as depicted in  FIG. 16  (see step  902  of  FIG. 9 ). As depicted in  FIG. 16 , second composite layup  1502  includes a number of layers  1606  that are stacked vertically to achieve a desired thickness  1608  for second composite layup  1502 , which is used to define initial thickness  310  after cure (see  FIG. 3 ). An outer surface  1610  of second composite layup  1502  forms outer surface  220  of sacrificial member  204  after composite assembly  200  is cured. 
     After assembling second composite layup  1502  onto adhesive film  1602 , second composite layup  1502  and adhesive film  1602  are transferred to flange portion  1406  and/or flange portion  1407  of first composite layup  1402  during the assembly process (see step  904 ), as depicted in  FIG. 17 , with adhesive film  1602  contacting outer surface  304  of flange  214  and/or outer surface  305  of flange  215 . Vacuum bagging, compression, and cure may be performed in order to generate composite assembly  200  of  FIG. 18 , with adhesive film  1602  disposed between sacrificial member  204  and flange  214  and/or flange  215 .  FIG. 19  depicts a cross-sectional view of  FIG. 18  along cut lines B-B, which illustrates the placement of adhesive film  1602  between sacrificial member  204  and flange  214  and/or flange  215 . 
     When edge  308  is tapered in sacrificial member  204 , second composite layup  1502  may be assembled on adhesive film  1602  as previously described with respect to  FIG. 9  and  FIG. 16 , and a taper may be cut in edge  1504  of second composite layup  1502  (see step  1002  and  FIG. 20 ), and second composite layup  1502  may then be transferred to flange portion  1406  and/or flange portion  1407  of first composite layup  1402  (see step  904 ), with edge  1504  aligned to bend  1408  and/or bend  1409  (see step  1004  and  FIG. 17 ). After cure, edge  1504  of second composite layup  1502  forms edge  308  of sacrificial member  204  for composite assembly  200  as depicted in a cross-section of composite assembly  200  in  FIG. 21 . 
     As discussed previously, there are a number of processes that may be used to determine how much material to remove from outer surface  220  of sacrificial member  204  after composite assembly  200  is cured and hardened. One process comprises linking or associating composite assembly  200  with specific skin panels  402 - 403  in aircraft structure  400  (see  FIG. 4 ). First, one or more skin panels may be identified that are to be bonded or mated to composite spar  202  (see step  1102  of  FIG. 11 ). Depending on the design of the structure that utilizes composite spar  202 , multiple skin panels may be bonded or mated to composite spar  202  along its length  206 . For example, skin panel  403  may be identified as being assigned for installation at flange  215  of composite spar  202  (see  FIG. 5 ). An interference may be calculated between skin panel  403  and outer surface  220  of sacrificial member  204  (see step  1104 ). For instance, a 3D scan of skin panel  403  and composite assembly  200  may be performed, and used to calculate the interference. Outer surface  220  of sacrificial member  204  may then be machined based on the interference (see step  1106 ). For instance, outer surface  220  of sacrificial member  204  may have initial thickness  310  prior to machining as depicted in  FIG. 22 , which is based on thickness  1608  of second composite layup  1502  prior to cure (see  FIG. 16 ). Outer surface  220  may then be machined down to final thickness  314  as depicted in  FIG. 22  prior to assembling composite spar  202  and skin panel  403  together. A machining depth  2202  that cuts into sacrificial member  204  from outer surface  220  towards flange  215  removes material from sacrificial member  204  until a desired final thickness  314  is achieved. Machining depth  2202  may vary along length  206  or portions of length  206  of composite spar  202 , even though  FIG. 22  merely depicts a cross-section of a particular location in aircraft structure  400  for purposes of discussion. Further, these same processes may be performed on sacrificial member  204  at flange  214 , although machining depth  2202  and/or initial thickness  310  may be different between flanges  214 - 215 . 
     Another process that may be used to determine how much material to remove from sacrificial member  204  after composite assembly  200  is cured and hardened may statically define machining depth  2202  based on the expected tolerances between composite spar  202  and skin panel  402  and/or skin panel  403 , and then make adjustments to machining depth  2202  if gaps are found between composite spar  202  and skin panel  402  and/or skin panel  403 . 
     First, a machining depth  2202  may be initially calculated for sacrificial member  204  (see step  1202  of  FIG. 12 ). For example, the components of aircraft structure  400 , when fabricated, have manufacturing variations that can be estimated and used to generate a worst-case scenario for gaps between flange  214  and skin panel  402  and/or flange  215  and skin panel  403  for aircraft structure  400 , which may be used to initially determine machining depth  2202 .  FIG. 22  illustrates one example of machining depth  2202  calculated for sacrificial member  204 , which indicates how much material will be removed from an outer surface  220  of sacrificial member  204 . 
     A machining process is performed on outer surface  220  of sacrificial member  204  to remove material to machining depth  2202  (see step  1204 ), which is depicted in  FIG. 23 , and composite assembly  200  and skin panel  403  are re-assembled as depicted in  FIG. 5  (see step  1206  of  FIG. 12 ). After re-assembly, spacing between contact surface  316  of sacrificial member  204  and inner surface  405  of skin panel  403  are checked to determine if gaps are present (see step  1208 ). If no gaps are present as depicted in  FIG. 5 , then machining depth  2202  may be used again when machining subsequent composite assemblies (see step  1212 ). However, if a gap  2402  is present between contact surface  316  and inner surface  405  of skin panel  403  as depicted in  FIG. 24 , which may be at any point along length  206  of composite spar  202 , then machining depth  2202  may be recalculated (e.g., reduced) for machining subsequent composite spars that are fabricated (see step  1210 ) in order to mitigate the gap  2402 . 
       FIG. 25  is a flow chart of another method  2500  of fabricating a composite assembly  200  in an illustrative embodiment and  FIG. 26  is a flow chart depicting additional details of method  2500 .  FIGS. 27-28  are cross-sectional views of flange  215  along cut lines C-C of  FIG. 2  in illustrative embodiments. 
     Step  2502  of method  2500  comprises performing first composite layup  1402  onto layup mandrel  1302  (see  FIG. 14 ), which defines the contour of composite spar  202 . This step may be similar to step  602  of method  600  previously described. Step  2504  comprises performing second composite layup  1502  onto flange portion  1406  and/or flange portion  1407  of first composite layup  1402 , as depicted in  FIG. 15 . This step may be similar to step  604  of method  600 , previously described. 
     Step  2506  comprises co-curing first composite layup  1402  and second composite layup  1502 . This step may be similar to step  606  of method  600 , previously described. The result of this process is composite assembly  200  of  FIG. 2 . 
     Step  2508  comprises calculating a machining depth for sacrificial member  204  based on estimated spacing tolerances between flange  215  of composite spar  202  and skin panel  403  for aircraft structure  400 , and step  2510  comprises machining outer surface  220  of sacrificial member  204  based on the machining depth. 
     In some cases, a machining depth  2702  is constant along length  206  of composite spar, as depicted in  FIG. 27 . In other cases, the machining depth may vary along a portion of length  206  of composite spar  202  (see step  2602  of  FIG. 26 ), as depicted in  FIG. 28 . In  FIG. 28 , machining on outer surface  220  of sacrificial member  204  is performed to a first machining depth  2802  along portion  2804  of length  206  of composite spar  202 , and machining on outer surface  220  of sacrificial member  204  is performed to a second machining depth  2806  along portion  2808  of length  206  of composite spar  202 , where the depths are different. The result is that contact surface  316  of sacrificial member  204  may vary in shape along length  206  of composite spar  202 . Although method  2500  has been described with respect to flange  215  of composite spar  202  and skin panel  403  of aircraft structure  400 , method  2500  applies equally to flange  214  of composite spar  202  and skin panel  402  of aircraft structure  400 , or any other types of structures for aircraft that utilize spars. 
     The use of co-cured sacrificial member  204  on composite spar  202  integrates sacrificial surfaces into flange  214  and/or flange  215 , thereby eliminating the manual step of shimming spars during an assembly process. In some cases, machining of the sacrificial plies is performed based on the estimated tolerances in the stack-up of components that form aircraft structures that use spars, thereby allowing the spars to be re-purposed between different builds of the aircraft structure on the factory floor. In other cases, 3D scanning may be performed on the components of a particular assembly, which may link those specific components together for that build on the factory floor. 
     The embodiments of the disclosure may be described in the context of an aircraft manufacturing and service method  2900  as shown in  FIG. 29  and aircraft  3004  as shown in  FIG. 30 . During pre-production, exemplary method  2900  may include a specification and design  2902  of aircraft  3004 , and material procurement  2904 . During production, component and subassembly manufacturing  2906  and system integration  2908  of aircraft  3004  takes place. Thereafter, aircraft  3004  may go through certification and delivery  2910  in order to be placed in service  2912 . While in service by a customer, aircraft  3004  is scheduled for routine maintenance and service  2914  (which may also include modification, reconfiguration, refurbishment, and so on). 
     Each of the processes of method  2900  may be performed or carried out by a system integrator, a third party, and/or an operator (e.g., a customer). For the purposes of this description, a system integrator may include without limitation any number of aircraft manufacturers and major-system subcontractors; a third party may include without limitation any number of venders, subcontractors, and suppliers; and an operator may be an airline, leasing company, military entity, service organization, and so on. 
     As shown in  FIG. 30 , aircraft  3004  produced by exemplary method  2900  may include an airframe  3002  with a plurality of systems  3016  and an interior  3006 . Examples of high-level systems  3016  include one or more of propulsion systems  3008 , an electrical system  3010 , a hydraulic system  3012 , and an environmental system  3014 . Any number of other systems may be included. Although an aerospace example is shown, the principles described in this specification may be applied to other industries, such as the automotive industry. 
     Apparatus and methods embodied herein may be employed during any one or more of the stages of the production and service method  2900 . For example, components or subassemblies corresponding to process  2906  may be fabricated or manufactured in a manner similar to components or subassemblies produced while aircraft  3004  is in service. Also, one or more apparatus embodiments, method embodiments, or a combination thereof may be utilized during the component subassembly and manufacturing  2906  and system integration  2908 , for example, by substantially expediting assembly of or reducing the cost of aircraft  3004 . Similarly, one or more of apparatus embodiments, method embodiments, or a combination thereof may be utilized while aircraft  3004  is in service, for example and without limitation, to maintenance and service  2914 . 
     Although specific embodiments were described herein, the scope is not limited to those specific embodiments. Rather, the scope is defined by the following claims and any equivalents thereof.