Patent Description:
In aircraft production, skin panels are used in a variety of structural assemblies that make up the aircraft. For example, the wings of an aircraft are comprised of an upper skin panel and a lower skin panel, coupled to a front spar, a rear spar, and a plurality of ribs (i.e., the mating structure). The upper and lower skin panels typically include spanwise stiffeners (i.e., stringers) coupled to the skin inner surfaces. The skin outer surface (i.e. the outer mold line) of the upper and lower skin panels defines the aerodynamic shape or profile of the wings. The spars are the primary structural members of the wing, and the ribs transfer aerodynamic loads on the skin panels to the spars. Various methods have been proposed for manufacturing structural assemblies such as aircraft wings. In one approach for manufacturing a wing made of composite material (e.g., carbon-fiber), the spars and ribs are assembled to form a ladder assembly, and the upper and lower skin panels are each laid up and cured on a dedicated layup mandrel. After curing, each skin panel is removed from its layup mandrel, and is machined one or more times at the interface locations where the skin panel is coupled to the ladder assembly. Each machining operation requires a time-consuming and costly machining setup. In addition, because each skin panel cannot be machined until the ladder assembly is completed, there are significant holding costs associated with the skin panels waiting to be machined, and the ladder assembly waiting for completion of the skin panels. The manufacturing approach may also result in a non-nominal aerodynamic profile of the wings after the skin panels are fastened to the ladder assembly.

Other manufacturing approaches have been proposed in attempts to increase part accuracy, and reduce the time and costs associated with producing structural assemblies. For example, one approach uses a specific holding fixture for supporting each cured skin panel during machining, trimming, and drilling. Unfortunately, the holding fixtures introduce tolerances and variations into the manufacturing process. In this regard, because the holding fixture can never hold the skin panel at the same surface profile as the layup mandrel that created the skin panel, slight variations are machined into each skin panel, which results in gaps between the skin panels and the ladder assembly. Gaps that exceed allowable tolerances require the time-consuming and costly process of manufacturing and installing individual shims. Furthermore, because the skin panel must be re-indexed to the supporting structure (e.g., the holding fixture, the ladder assembly) each time the skin panel is moved, additional tolerances are introduced.

As can be seen, there exists a need in the art for a method of manufacturing structure assemblies that avoids the drawbacks associated with the above-described manufacturing approaches.

<CIT>, in accordance with its abstract, describes a method and apparatus for manufacturing wings including a fixture that holds wing panels for drilling and edge trimming by accurate numerically controlled machine tools using original numerical part definition records, utilizing spatial relationships between key features of detail parts or subassemblies as represented by coordination features machined into the parts and subassemblies, thereby making the parts and subassemblies intrinsically determinant of the dimensions and contour of the wing. Spars are attached to the wing panel using the coordination holes to locate the spars accurately on the panel in accordance with the original engineering design, and in-spar ribs are attached to rib posts on the spar using accurately drilled coordination holes in the ends of the rib and in the rib post. The wing contour is determined by the configuration of the spars and ribs rather than by any conventional hard tooling which determines the wing contour in conventional processes.

The above-noted needs associated with manufacturing panel assemblies are addressed by the present disclosure, which provides a method of manufacturing a panel assembly. The method includes supporting the panel assembly in a free state using a holding fixture in which the panel assembly has a geometric shape that is different than the geometric shape of the panel assembly in an as-designed nominal state. The panel assembly has a skin panel, and sacrificial material coupled to a skin panel inner surface respectively at interface locations for coupling the panel assembly to mating structure. The method includes acquiring a free state outer surface contour of the panel assembly by scanning a skin panel outer surface while the panel assembly is supported by the holding fixture. The method also includes developing a numerically controlled (NC) machining program having cutter paths configured for machining the interface locations to an inner surface contour that reflects nominal thicknesses of the panel assembly based off of the free state outer surface contour. In addition, the method includes machining the sacrificial material at the interface locations by moving a cutter along the cutter paths while the panel assembly is supported by the holding fixture, and while the cutter is backed by a backing device applying backing pressure against the skin panel outer surface. The backing device moves in a coordinated manner in alignment with the cutter as the cutter moves along the cutter paths on the opposite side of the skin panel.

Also disclosed is a method of manufacturing a panel assembly, comprising supporting a panel assembly in a free state using a holding fixture, and acquiring a free state outer surface contour of the panel assembly by scanning a skin panel outer surface of the skin panel while the panel assembly is supported by the holding fixture. The method includes developing an NC machining program having cutter paths configured for machining the interface locations to an inner surface contour that is based on the free state outer surface contour, and machining the sacrificial material by moving a cutter along the cutter paths while the panel assembly is supported by the holding fixture. The method also includes removing the panel assembly from the holding fixture, and indexing the panel assembly and the mating structure to each other, and fastening the panel assembly and the mating structure together to result in a structural assembly.

In another example, disclosed is a method of manufacturing an aerostructure of an aircraft. The method includes supporting a panel assembly of an aerostructure in a free state using a holding fixture in which the panel assembly has a geometric shape that is different than the geometric shape of the panel assembly in an as-designed nominal state. The skin panel has a skin panel outer surface configured to define an outer mold line (OML) that forms an aerodynamic contour of the aerostructure. The method further includes acquiring a free state outer surface contour of the skin panel by scanning the skin panel outer surface while the panel assembly is supported by the holding fixture in the free state, and developing an NC machining program having cutter paths configured for machining the interface locations to an inner surface contour that reflects nominal thicknesses of the panel assembly based on the free state outer surface contour. The method also includes machining the sacrificial material by moving a cutter along the cutter paths while the panel assembly is supported by the holding fixture, and while the cutter is backed by a backing device applying backing pressure against the skin panel outer surface.

The features, functions, and advantages that have been discussed can be achieved independently in various versions of the disclosure or may be combined in yet other versions, further details of which can be seen with reference to the following description and drawings.

The disclosure can be better understood with reference to the following detailed description taken in conjunction with the accompanying drawings, which illustrate preferred and exemplary versions, but which are not necessarily drawn to scale. The drawings are examples and not meant as limitations on the description or the claims.

The figures shown in this disclosure represent various aspects of the versions presented, and only differences will be discussed in detail.

Disclosed versions will now be described more fully hereinafter with reference to the accompanying drawings, in which some, but not all of the disclosed versions are shown. Indeed, several different versions may be provided and should not be construed as limited to the versions set forth herein. Rather, these versions are provided so that this disclosure will be thorough and fully convey the scope of the disclosure to those skilled in the art.

This specification includes references to "one version" or "a version. " Instances of the phrases "one version" or "a version" do not necessarily refer to the same version. Similarly, this specification includes references to "one example" or "an example. " Instances of the phrases "one example" or "an example" do not necessarily refer to the same example. Particular features, structures, or characteristics may be combined in any suitable manner consistent with this disclosure.

As used herein, "comprising" is an open-ended term, and as used in the claims, this term does not foreclose additional structures or steps.

As used herein, "configured to" means various parts or components may be described or claimed as "configured to" perform a task or tasks. In such contexts, "configured to" is used to connote structure by indicating that the parts or components include structure that performs those task or tasks during operation. As such, the parts or components can be said to be configured to perform the task even when the specified part or component is not currently operational (e.g., is not on).

As used herein, the phrase "at least one of," when used with a list of items, means different combinations of one or more of the listed items may be used, and only one of each item in the list may be needed. In other words, "at least one of" means any combination of items and number of items may be used from the list, but not all of the items in the list are required. The item may be a particular object, a thing, or a category.

Referring now to the drawings which illustrate various examples of the disclosure, shown in <FIG> is a flowchart of a method <NUM> of manufacturing a panel assembly <NUM> (<FIG>). The panel assembly <NUM> includes a skin panel <NUM> (<FIG>), and sacrificial material <NUM> (<FIG>) applied to or integrated with the skin panel <NUM> at a plurality of discrete interface locations <NUM> where the panel assembly <NUM> is attached to mating structure <NUM> (<FIG>). As described in greater detail below, the sacrificial material <NUM> is machined in a manner resulting in the panel assembly <NUM> having nominal thicknesses <NUM> (<FIG>) at each interface location <NUM>, thereby reducing or eliminating the need for shimming gaps that may otherwise occur at the interface locations <NUM> when the panel assembly <NUM> is attached to mating structure <NUM>.

The skin panel <NUM> (<FIG>) has a skin panel outer surface <NUM> (<FIG>) and a skin panel inner surface <NUM> (<FIG>). In some examples, the panel assembly <NUM> includes skin stiffeners <NUM> (<FIG>) extending along a spanwise direction of the skin panel <NUM>. The skin stiffeners <NUM> each have one or more stiffener flanges <NUM> (<FIG>), and a stiffener web <NUM> (<FIG>) extending outwardly from the stiffener flanges <NUM>. The stiffener flanges <NUM> are coupled to the skin panel inner surface <NUM>. The panel assembly <NUM> has a panel assembly outer surface <NUM> (<FIG>) and a panel assembly inner surface <NUM> (<FIG>). The panel assembly outer surface <NUM> comprises (i.e., is defined by) the skin panel outer surface <NUM>. The panel assembly inner surface <NUM> is defined by the skin panel inner surface <NUM> and the exposed surfaces of the stiffener flanges <NUM>.

As mentioned above, the method <NUM> include mating the panel assembly <NUM> (<FIG>) to mating structure <NUM> (<FIG>), to thereby result in a structural assembly <NUM>. Shown in <FIG> are illustrations of the implementation of the method <NUM> in manufacturing a structural assembly <NUM> configured as a wing <NUM> of an aircraft <NUM> (<FIG>). However, the presently-disclosed method <NUM> may be implemented for manufacturing other types of structural assemblies <NUM>, and is not limited to manufacturing a wing <NUM>. For example, <FIG> illustrates an aircraft <NUM> comprised of various structural assemblies <NUM>, which are described herein as aerostructures <NUM>. In the present disclosure, an aerostructure <NUM> is one in which the skin panel outer surface <NUM> defines the aerodynamic contour of at least a portion of the aircraft <NUM>. In this regard, the skin panel outer surface <NUM> defines the outer mold line (OML) for air flowing over the aerostructure <NUM> when the aircraft <NUM> is in operation, such as during flight.

Referring to <FIG>, the aircraft <NUM> includes other types of aerostructures <NUM> that may be manufactured using the presently-disclosed method <NUM>. Such aerostructures <NUM> include ailerons <NUM>, flaps <NUM>, and/or wingtip devices <NUM>, such as winglets. Other aerostructures <NUM> include horizontal stabilizers <NUM>, elevators <NUM>, vertical stabilizers <NUM>, rudders <NUM>, fuselage panels <NUM>, engine nacelles <NUM>, engine cowlings <NUM>, and any one of a variety of other types of aerostructures <NUM> that are part of an aircraft <NUM>. However, the method <NUM> may be implemented for manufacturing any type of structure, substructure, assembly, or subassembly, without limitation. In addition, the method <NUM> may be implemented for manufacturing structural assemblies <NUM> for any type of application, and is not limited to aircraft production. In this regard, the method <NUM> may be implemented for manufacturing any type of movable or non-movable structure. Examples of movable structures include, but are not limited to, any type of land-based vehicle, any type of air vehicle including fixed-wing aircraft (e.g., <FIG>) and rotary wing aircraft, any type of space vehicle, and any type of marine vessel. Examples of non-movable structures include, but are not limited to, buildings, architectural objects, utility structures such as wind turbines (e.g., turbine blades), and other types of generally non-movable objects.

Shown in <FIG> is an example of a wing <NUM>, which has a wing root <NUM> extending to a wingtip <NUM>. The wing <NUM> includes leading edge devices <NUM>, such as slats or leading edge flaps. The trailing edge includes the above-mentioned trailing edge devices, such as ailerons <NUM> and trailing edge flaps <NUM>. Referring to <FIG>, the wing <NUM> has upper and lower skin panels <NUM>, each of which is coupled to internal structural components <NUM> (i.e., the mating structure <NUM>). In the present example, the internal structural components <NUM> of the wing <NUM> include a front spar <NUM>, a rear spar <NUM>, and a plurality of ribs <NUM>. In the example shown, the upper and lower skin panels <NUM> include spanwise skin stiffeners <NUM> (i.e., stringers). The outer surfaces of the upper and lower skin panels <NUM> serve as the outer mold line (OML) of the wings <NUM>, and define the aerodynamic shape of the wings <NUM>. During flight, the ribs <NUM> transfer aerodynamic loads on the skin panels <NUM> into the front spar <NUM> and the rear spar <NUM>, which are the primary load-carrying members of the wing <NUM>.

Referring to <FIG>, shown in <FIG> are cross-sectional views of the wing <NUM> of <FIG>, showing the skin stiffeners <NUM> coupled to the inner surfaces of the upper and lower skin panels <NUM>. As mentioned above, each skin stiffener <NUM> is comprised of stiffener flanges <NUM> and a stiffener web <NUM>. The stiffener flanges <NUM> are coupled to the skin panel inner surfaces <NUM> in a manner described below.

Also shown in <FIG> are the front spar <NUM>, the rear spar <NUM>, and the ribs <NUM>. Each rib <NUM> includes mouse holes to allow the stiffener webs <NUM> to pass through the rib <NUM>. The forward and aft end of each rib <NUM> includes rib flanges <NUM> for coupling the rib <NUM> respectively to the front spar <NUM> and the rear spar <NUM>, via mechanical fasteners <NUM>. In addition, the top and bottom side of each rib <NUM> has a series of rib flanges <NUM> (e.g., rib shear ties) for directly or indirectly coupling the rib <NUM> to the skin panels <NUM> on the upper and lower sides of the wing <NUM>. For example, as shown in <FIG>, the interface location <NUM> for some of the rib flanges <NUM> is on the skin panel inner surface <NUM> of the skin panels <NUM>. As shown in <FIG>, the interface location <NUM> for other rib flanges <NUM> of the same rib <NUM> is on the stiffener flanges <NUM>. At each interface location <NUM>, mechanical fasteners <NUM> are installed for fastening the skin panels <NUM> to the ribs <NUM>. Mechanical fasteners <NUM> are also used for fastening the skin panels <NUM> to the spar flanges <NUM>.

<FIG> are exploded views of the wing <NUM> showing an upper panel assembly <NUM>, a lower panel assembly <NUM>, and the internal structural components <NUM>, comprising the front spar <NUM>, the rear spar <NUM>, and the ribs <NUM>. In <FIG>, shown are the interface locations <NUM> on the lower panel assembly <NUM> where the skin panel <NUM> is attached to the front spar <NUM>, the rear spar <NUM>, and the ribs <NUM>. <FIG> shows the upper and lower panel assembly <NUM>, <NUM> in a nominal state <NUM> prior to attachment to the front spar <NUM>, rear spar <NUM> and ribs <NUM>. When the upper and lower panel assembly <NUM>, <NUM> are in the nominal state <NUM> and are attached to the internal structural components <NUM> at the interface locations <NUM> (e.g., <FIG>), the skin panels <NUM> have an as-designed geometric shape, and each skin panel outer surface <NUM> has an as-designed nominal outer surface contour. <FIG> show the lower panel assembly <NUM>, and the interface locations <NUM> where the lower panel assembly <NUM> is attached to the front spar <NUM>, the rear spar <NUM>, and the ribs <NUM>.

Referring now to <FIG>, shown in <FIG> is a partially exploded view of a portion of the lower panel assembly <NUM>, illustrating sacrificial material <NUM> for application or integration at each interface location <NUM> on the skin panel inner surface <NUM> and on the stiffener flanges <NUM>. <FIG> is a magnified view of a portion of the lower panel assembly <NUM> showing an example of the sacrificial material <NUM> applied to the interface locations <NUM> on the skin panel inner surface <NUM> and on the stiffener flanges <NUM>. <FIG> is a sectional view of the lower panel assembly <NUM> showing the sacrificial material <NUM> at the interface locations <NUM>.

As described in greater detail below, the sacrificial material <NUM> is applied to, or integrated with, the panel assembly <NUM> at each interface location <NUM>. The sacrificial material <NUM> provides a means for manufacturing a structural assembly <NUM> such that the skin panel <NUM> has highly accurate thicknesses (i.e., nominal thicknesses <NUM> - <FIG>) at the interface locations <NUM>. Advantageously, by manufacturing the structural assembly <NUM> such that the skin panel <NUM> has nominal thicknesses <NUM> at the interface locations <NUM>, the need to install shims between the skin panel <NUM> and the mating structure <NUM> is reduced or eliminated, as described in greater detail below.

Referring to <FIG>, the method <NUM> will now be described in the context of manufacturing a panel assembly <NUM> of a wing <NUM> formed of composite material. The panel assembly <NUM> is an upper panel assembly <NUM> (<FIG>) or a lower panel assembly <NUM> (<FIG>) of a wing <NUM>. However, as mentioned above, the method <NUM> is applicable for manufacturing structural assemblies <NUM> any one of a variety of different types of panel assemblies <NUM> formed of any type of material, including any type of metallic material and/or any type of non-metallic material.

In manufacturing a composite wing <NUM>, the method <NUM> includes laying up a composite skin panel <NUM> on a mandrel surface <NUM> of a layup mandrel <NUM>. The mandrel surface <NUM> is shaped to the as-designed contour of the skin panel outer surface <NUM>. Each skin panel <NUM> is laid up by sequentially laying up individual plies (not shown) of composite material on the layup mandrel <NUM>. The composite material may be a fiber-reinforced polymer matrix material. In one example, the composite material is a prepreg material comprised of unidirectional reinforcing fibers pre-impregnated with resin. The reinforcing fibers may be formed of any one of a variety of materials, such as plastic, glass, ceramic, carbon, metal, or any combination thereof. The resin is a thermosetting resin or a thermoplastic resin, and may be formed of any one of a variety of organic or inorganic materials. In one example, the composite material is carbon-fiber-reinforced plastic (CFRP) prepreg.

The skin stiffeners <NUM> (i.e., stringers) may be laid up separate from the laying up of the skin panel <NUM>, and may be formed of the same material or a different material than the skin panel <NUM>. In one example, the skin stiffeners <NUM> are laid up using plies of carbon-fiber-reinforced plastic (CFRP) prepreg, or other material that is compatible with the material of the skin panel <NUM>. The skin stiffeners <NUM> may be coupled to the skin panel <NUM> via co-curing or co-bonding to the skin panel inner surface <NUM>, or by secondarily bonding to the skin panel <NUM> (after curing). The method <NUM> includes applying sacrificial material <NUM> (<FIG>) to the interface locations <NUM> on the panel assembly inner surface <NUM>, as described in greater detail below. As mentioned above, the interface locations <NUM> comprise locations where the panel assembly <NUM> (after curing) is to be attached to mating structure <NUM>. The sacrificial material <NUM> at each interface location <NUM> preferably has a footprint (e.g., a length and a width) that is approximately (e.g., within <NUM> percent) the same size as the footprint of the interface location <NUM> of the mating structure <NUM> to be attached to the panel assembly <NUM> at that interface location <NUM>. For example, the footprint of the sacrificial material <NUM> at an interface location <NUM> where a rib flange <NUM> (<FIG>) mates to the skin panel <NUM> is preferably the same size as the footprint of the rib flange <NUM>. Similarly, the footprint of the sacrificial material <NUM> at an interface location <NUM> where a rib flange <NUM> attaches to a stiffener flange <NUM> (<FIG>) is preferably the same size as the footprint of the rib flange <NUM> at that location. Likewise, the footprint of the sacrificial material <NUM> at an interface location <NUM> where a spar flange <NUM> (<FIG>) mates to the skin panel <NUM> is preferably the same size as the footprint of the spar flange <NUM> at that location. Alternatively, the sacrificial material <NUM> at each interface location <NUM> preferably has a footprint that is no smaller than the footprint of the interface location <NUM> of the mating structure <NUM> at that interface location <NUM>.

The sacrificial material <NUM> is applied at each interface location <NUM> in a thickness such that, after machining (as described below), the skin panel <NUM> has nominal thicknesses <NUM> (<FIG>) at each interface location <NUM>. In addition, the sacrificial material <NUM> is applied at each interface location <NUM> in a thickness that, prior to machining, results in the combined thickness of the skin panel <NUM> and the sacrificial material <NUM> being greater than the maximum thickness tolerance at that interface location <NUM>. In one example, the sacrificial material <NUM> is applied in a pre-machined thickness of no less than <NUM> inch at each interface location <NUM>, in a manner described below. However, the sacrificial material <NUM> may be applied in any pre-machined thickness, and is not limited to a pre-machined thickness of no less than <NUM> inch.

The sacrificial material <NUM> may be formed of any material, and preferably sacrificial material <NUM> that is easily machinable (e.g., aluminum, fiberglass, etc.). In addition, the sacrificial material <NUM> is preferably mechanically and chemically stable during manufacturing, and when exposed to the service environment of the panel assembly <NUM>. For example, the sacrificial material <NUM> is preferably non-compressible or non-deformable by more than <NUM> percent (i.e., in the thickness direction) when in service. Furthermore, the sacrificial material <NUM> preferably has a melting temp that is below the service temperature of the panel assembly <NUM>. In addition, the sacrificial material <NUM> is preferably non-outgassing in the service environment of the panel assembly <NUM>, and/or is non-dissolvable when exposed to the elements. Other preferable mechanical properties include a coefficient of thermal expansion (CTE) that is approximately (e.g., ±<NUM> percent) of the CTE of the material of the panel assembly <NUM>, at least within the service temperature range of the panel assembly <NUM>.

Examples of the sacrificial material <NUM> include, but are not limited to, fiber-reinforced polymer matrix material (i.e., composite material) such as fiberglass or CFRP. Other examples include non-fibrous polymeric material, such as epoxy or moldable plastic. In another example, the sacrificial material <NUM> may comprise non-polymeric material, or metallic material (e.g., aluminum). Still other examples include fiber metal laminate, including GLARE™, described as glass-aluminum-reinforced epoxy. The sacrificial material <NUM> for each interface location <NUM> may be separately manufactured, and/or co-cured, co-bonded, or secondarily bonded to each interface location <NUM>.

In one example, the sacrificial material <NUM> may be applied by sequentially laying up a localized stack of plies (not shown) of fiber-reinforced polymer matrix material at each of the interface location <NUM>. The fiber-reinforced polymer matrix material may be a fiberglass material, a carbon-fiber-reinforced polymeric material, or other material. The method <NUM> may include laying up additional, localized plies of the same material (e.g., CFRP) as the skin panel <NUM>, or laying up localized plies of a different material than the skin panel <NUM>. In another example, the method <NUM> may include separately laying up the sacrificial material <NUM> for each interface location <NUM>, and then installing the uncured sacrificial material <NUM> at the interface locations <NUM>, followed by co-curing or co-bonding with the panel assembly <NUM> (i.e., the skin panel <NUM> and the skin stiffeners <NUM>), using the arrangement shown in <FIG>. Alternatively, the method <NUM> may include separately laying up and pre-curing the sacrificial material <NUM> for each interface location <NUM>, and then secondarily bonding the sacrificial material <NUM> to the cured panel assembly <NUM>.

Referring to <FIG>, after laying up the skin panel <NUM> and locating the skin stiffeners <NUM> on the skin panel inner surface <NUM>, a vacuum bag <NUM> and other processing layers (e.g., breather fabric, release film, etc.) are applied over the layup components, and the side edges of the vacuum bag <NUM> are sealed to the layup mandrel <NUM> using a bag sealant <NUM>. Vacuum pressure is applied to the interior of the vacuum bag <NUM> via a vacuum source <NUM>. The application of vacuum pressure results in compaction pressure <NUM> on the skin panel <NUM>. Heat <NUM> is applied to initiate and/or promote the curing of the composite material of the skin panel <NUM> and/or skin stiffeners <NUM>. During layup and curing, the skin panel outer surface <NUM> assumes the contour of the mandrel surface <NUM>.

Referring to <FIG>, shown is an example of the panel assembly <NUM> after curing is complete, and after the vacuum bag <NUM> and other layup components have been removed. As can be seen, the post-cured panel assembly <NUM> exhibits springback <NUM>, in which the panel assembly <NUM> assumes a geometric shape that is different than the geometric shape of the panel assembly <NUM> in the nominal state <NUM>, shown in phantom lines. In <FIG>, the springback <NUM> manifests as a decrease in the radius of curvature of the skin panel <NUM>, relative to the radius of curvature of the skin panel <NUM> in the nominal state <NUM>. Springback <NUM> occurs after curing, when the panel assembly <NUM> is released from forces (i.e., compaction pressure <NUM>) that hold the panel assembly <NUM> against the layup mandrel <NUM>. Springback <NUM> occurs primarily as a result of a mismatch in the CTE of the resin relative to the CTE of the reinforcing fibers of the composite material. In the case of a metallic panel assembly (not shown), springback <NUM> may occur when the metallic panel assembly is released from forming forces (e.g., brake-forming, hydroforming, etc.), causing the metallic panel assembly <NUM> to take on a geometric shape that is different than the geometric shape of the metallic panel assembly <NUM> in the nominal state.

It should be noted that, in addition to springback, other forces may cause the panel assembly <NUM> in the free state <NUM> to assume a geometric shape that is different than the geometric shape of the panel assembly <NUM> in the nominal state <NUM>. For example, changes in the orientation of the panel assembly <NUM> and/or the manner in which the panel assembly <NUM> is supported (e.g., fixturing) may cause the panel assembly <NUM> to assume a geometric shape that is different than the geometric shape of the panel assembly <NUM> in the nominal state <NUM> due to gravity and/or locally applied loads. It should also be noted that although the panel assembly <NUM> in the present example is configured such that the skin panel outer surface <NUM> has a convex shape, in other examples not shown, the presently-disclosed method <NUM> may be implemented for a panel assembly <NUM> in which the skin panel outer surface <NUM> has a concave shape, or any other shape, including any simply curved shape or any complexly curved shape.

Referring to <FIG>, after initially forming the panel assembly <NUM>, step <NUM> of the method <NUM> (<FIG>) includes supporting the panel assembly <NUM> in a free state <NUM> using a holding fixture <NUM> in which the panel assembly <NUM> is supported in a geometric shape that is different than the geometric shape of the panel assembly <NUM> in the nominal state <NUM> (e.g., shown in phantom in <FIG>). When the panel assembly <NUM> is in the nominal state <NUM>, the skin panel outer surface <NUM> has a nominal outer surface contour.

In <FIG>, the holding fixture <NUM> is configured as an orthogonally-shaped picture frame tool <NUM>, comprised of a pair of horizontally-oriented beams interconnected on opposite ends by a pair of vertically-oriented beams. In this example, step <NUM> comprises supporting the panel assembly <NUM> at attachment locations along the perimeter edges <NUM>. The picture frame tool <NUM> has a plurality of support arms <NUM> located at spaced intervals along the beams. The panel assembly <NUM> is supported by at least two support arms <NUM> at each attachment location. Each support arm <NUM> is telescopically adjustable in length to allow the holding fixture <NUM> to adapt to panel assemblies <NUM> of different sizes and/or shapes. Once adjusted, the length of each telescopically adjustable support arm <NUM> is locked.

After the panel assembly <NUM> is loaded into the holding fixture <NUM>, the holding fixture <NUM> may be rotated from a horizontal orientation (e.g., <FIG>) to the vertical orientation shown in <FIG>, to facilitate further processing (e.g., scanning, machining, drilling, trimming, etc.) of the panel assembly <NUM> in the manner described below. Although shown and described as a picture frame tool <NUM> having support arms <NUM>, the holding fixture <NUM> may be provided in any one of a variety of sizes, shapes, and configurations, and is not limited to a picture frame tool <NUM> for supporting a panel assembly <NUM> of a wing <NUM>.

Referring to <FIG>, step <NUM> of the method <NUM> (<FIG>) includes acquiring a free state outer surface contour <NUM> of the panel assembly <NUM> by scanning the skin panel outer surface <NUM> of the skin panel <NUM> while the panel assembly <NUM> is supported in the free state <NUM> by the holding fixture <NUM>. In the example shown, the skin panel outer surface <NUM> is scanned using a scanning device <NUM> that is supported by a robotic arm <NUM> of a robotic device <NUM>. The robotic device <NUM> is movable along a track <NUM> located on one side of the holding fixture <NUM>. Although shown supported by a robotic device <NUM>, the scanning device <NUM> may be supported by alternative means, such as a gantry system (not shown), or other automated and/or programmable controlling device.

In the example of <FIG>, the scanning device <NUM> is a laser line scanner. However, the skin panel outer surface <NUM> may be scanned using any type of three-dimensional (3D) metrology system <NUM>, and is not limited to scanning via a laser line scanner. For example, the skin panel outer surface <NUM> may be scanned using a laser radar device, a surface profiler, a photogrammetry system, or any one of a variety of other types of metrology systems <NUM> for acquiring a digital representation of the three-dimensional shape of an object. During or after scanning, a processor <NUM> (<FIG>) generates a digital representation of the free state outer surface contour <NUM> of the panel assembly <NUM> based on the scanning data received from the scanning device <NUM>.

Referring to <FIG>, step <NUM> of the method <NUM> (<FIG>) includes developing, using a processor <NUM>, a numerically controlled (NC) machining program (i.e., a free state NC machining program <NUM>) having cutter paths <NUM> (<FIG>) configured for machining the interface locations <NUM> to an inner surface contour (i.e., a free state inner surface contour <NUM> - <FIG>) that reflects the nominal thicknesses <NUM> (<FIG>) based off of the free state outer surface contour <NUM>. The cutter paths <NUM> of the free state NC machining program <NUM> are configured for machining the interface locations <NUM> on the skin panel <NUM>, and machining the interface locations <NUM> on the stiffener flanges <NUM>. As mentioned above, the free state outer surface contour <NUM> is acquired by scanning the skin panel outer surface <NUM> while the panel assembly <NUM> is supported in the free state <NUM> by the holding fixture <NUM>. As shown in <FIG> and mentioned above, the geometric shape of the panel assembly <NUM> in the free state <NUM> is different than the geometric shape of the panel assembly <NUM> in the nominal state <NUM>. In the example shown, the radius of curvature of the skin panel outer surface <NUM> in the free state <NUM> is smaller than the radius of curvature of the skin panel outer surface <NUM> in the nominal state <NUM>.

One process for performing step <NUM> of developing the NC machining program comprises: creating cutter paths <NUM> (<FIG>) of a new free state NC machining program <NUM> mapped to a digital representation of the free state inner surface contour <NUM>. The processor <NUM> generates the digital representation of the free state inner surface contour <NUM> by offsetting nominal thicknesses <NUM> (i.e., the as-designed thicknesses) of the panel assembly <NUM> from the digital representation of the free state outer surface contour <NUM> acquired during scanning. More specifically, the processor <NUM> offsets the nominal thicknesses <NUM> of the panel assembly <NUM> respectively from each of a plurality of points in the digital representation of the free state outer surface contour <NUM>. The nominal thicknesses <NUM> of the panel assembly <NUM> are extracted from a computer-aided-design (CAD) model <NUM> of the panel assembly <NUM>.

<FIG> shows a portion of the panel assembly <NUM> in the nominal state <NUM>, and illustrates the nominal thicknesses <NUM> at different interface locations <NUM> on the panel assembly <NUM>. <FIG> shows the same portion of the panel assembly <NUM> in the free state <NUM>, and illustrates the free state inner surface contour <NUM>, which is generated by offsetting the nominal thicknesses <NUM> (<FIG>) of the panel assembly <NUM> from the free state outer surface contour <NUM>.

In above-describe example of performing step <NUM>, the nominal thicknesses <NUM> are used as a proxy for generating the inner surface contour of the panel assembly <NUM>. The resulting cutter paths <NUM> of the new free state NC machining program <NUM> are configured to machine the sacrificial material <NUM> at the interface locations <NUM> in a manner such that the panel assembly <NUM> has nominal thicknesses <NUM> at each interface location <NUM>, thereby reducing or eliminating the need for shimming of gaps that may otherwise occur at the interface locations <NUM> when the panel assembly <NUM> is attached to mating structure <NUM>. In addition, the cutter paths <NUM> of the free state NC machining program <NUM> are configured to machine the sacrificial material <NUM> at the interface locations <NUM> in a manner such that when the panel assembly <NUM> moves into the nominal state <NUM> during attachment to the mating structure <NUM>, the effects of springback <NUM> (<FIG>) are reversed, and the skin panel outer surface <NUM> assumes the nominal state (i.e., the as-designed contour) of the skin panel outer surface <NUM>.

An alternative process for performing step <NUM> of developing the free state NC machining program <NUM> comprises: adjusting, using the processor <NUM>, the cutter paths <NUM> of an existing nominal state NC machining program <NUM> in a manner reflecting differences between the free state outer surface contour <NUM> (<FIG>) and the nominal state outer surface contour <NUM> (<FIG>) of the panel assembly <NUM> in the nominal state <NUM>. The cutter paths <NUM> of the nominal state NC machining program <NUM> are originally configured for machining the interface locations <NUM> of the panel inner surface of the skin panel <NUM> to the nominal state <NUM> inner surface contour when the skin panel <NUM> is in the nominal state <NUM>. In this alternative process, for each point on the skin panel outer surface <NUM>, the processor <NUM> calculates the differences between the digital representation of the free state outer surface contour <NUM> and the digital representation of the nominal state outer surface contour <NUM>.

As mentioned above, the digital representation of the free state outer surface contour <NUM> is a result of scanning the panel assembly <NUM> while supported in the free state <NUM> by the holding fixture <NUM>. The digital representation of the nominal state outer surface contour <NUM> is extracted from the CAD model <NUM> of the panel assembly <NUM>. The processor <NUM> adjusts the cutter paths <NUM> of the nominal state NC machining program <NUM> to account for differences between the free state outer surface contour <NUM> and the nominal state outer surface contour <NUM>. For example, for each point along the cutter paths <NUM> of the nominal state NC machining program <NUM>, the processor <NUM> adjusts the spatial location (i.e., the three-dimensional location) of the cutter <NUM> at each point along the cutter paths <NUM>. In addition, for each point along the cutter paths <NUM> of the nominal state NC machining program <NUM>, the processor <NUM> may also adjust the spatial orientation (i.e., the three-dimensional orientation) of the cutter <NUM> to be locally perpendicular to the surface being machined.

Referring to <FIG>, step <NUM> of the method <NUM> (<FIG>) includes machining the sacrificial material <NUM> at the interface locations <NUM> by moving a cutter <NUM> along the cutter paths <NUM> (<FIG>) of the free state NC machining program <NUM> while the panel assembly <NUM> is supported in the free state <NUM> by the holding fixture <NUM>, and while the cutter <NUM> is backed by a backing device <NUM> applying backing pressure against the skin panel outer surface <NUM>. In the example shown, the cutter <NUM> is supported by a robotic arm <NUM> of a robotic device <NUM> on one side of the holding fixture <NUM>, and the backing device <NUM> is supported by a robotic arm <NUM> of a robotic device <NUM> on an opposite side of the holding fixture <NUM>. Each robotic device <NUM> is independently movable along a track <NUM>. However, the cutter <NUM> and the backing device <NUM> may be supported using any one of a variety of means, and are not limited to being supported by robotic devices <NUM>. For example, the cutter <NUM> and the backing device <NUM> may each be independently movable by a gantry system (not shown), or other automated and/or programmable controlling device.

In the example of <FIG>, the cutter <NUM> is a high-speed rotary cutter <NUM>, such as an end mill. However, the cutter <NUM> may be provided in any one of a variety of alternative devices for machining the sacrificial material <NUM>. The backing device <NUM> is configured to apply backing pressure at the skin panel outer surface <NUM> opposite the cutter <NUM> on the skin panel inner surface <NUM>. The backing device <NUM> is configured to move in a coordinated manner with the cutter <NUM> as the cutter <NUM> moves along the cutter paths <NUM> on the opposite side of the skin panel <NUM>.

The backing device <NUM> is configured to remain in alignment with the cutter <NUM> as the cutter <NUM> moves along the cutter paths <NUM>. For example, the centerline or axis of the backing device <NUM> remains parallel to and/or generally aligned with the centerline or axis of the cutter <NUM> during the machining process. The application of backing pressure by the backing device <NUM> prevents the panel assembly <NUM> from moving in response to pressure applied by the cutter <NUM> against the skin panel <NUM>. In the example shown, the backing device <NUM> is a sphere <NUM> configured to roll along the skin panel outer surface <NUM> and apply backing pressure equal in magnitude to the pressure applied by the cutter <NUM> on the opposite side of the skin panel <NUM>. However, the backing device <NUM> may be provided in alternative configurations for applying backing pressure to counteract the cutter <NUM>. For example, the backing device <NUM> may be configured to direct a stream of fluid (not shown) against the skin panel outer surface <NUM> to counteract the pressure applied by the cutter <NUM> on the opposite side of the skin panel <NUM>. Referring to <FIG>, step <NUM> of machining the sacrificial material <NUM> may include machining into the skin panel <NUM> at one or more of the interface locations <NUM>. In this regard, in order to achieve the nominal thickness <NUM> at a given interface location <NUM>, the cutter path <NUM> may be such that the cutter <NUM> machines off the entire thickness of the sacrificial material <NUM> at that interface location <NUM>, and then machines into the skin panel <NUM>, until achieving the nominal thickness.

In <FIG>, shown are several examples of the different types of surfaces that may be machined on the sacrificial material <NUM> at the interface locations <NUM>. For example, <FIG> illustrates machining the panel assembly <NUM> to achieve a nominal thickness <NUM> that is a constant thickness <NUM> at all points of the interface location <NUM>. <FIG> illustrates an example of an interface location <NUM> in which the sacrificial material <NUM> has been machined to several different nominal thicknesses <NUM>, to result in a linear tapered thickness <NUM>. <FIG> illustrates an example in which the sacrificial material <NUM> has been machined to result in a planar surface <NUM>. <FIG> illustrates an example of a ruled surface <NUM> machined into the sacrificial material <NUM>. Although the ruled surface is shown as a cylindrical surface, other types of ruled surfaces (e.g., a conical surface) may be machined into the sacrificial material <NUM>. In still other examples, a complex surface with different radii of curvature may be machined into the sacrificial material <NUM>. Still other examples of surfaces machined by the cutter <NUM> include non-uniform rational B-spline (NURBS) surfaces, smooth and continuous surfaces, or any one of a variety of other types of surfaces that achieve nominal thicknesses <NUM> at the interface locations <NUM>. The type of surface that is machined at each interface location <NUM> may be dictated in part by the surface of the mating structure <NUM> at that location.

After machining the inner surface of the panel assembly <NUM>, the panel assembly <NUM> is removed from the holding fixture <NUM>, and assembled to the mating structure <NUM> (e.g., <FIG>) via a drill-on-assembly process. The panel assembly <NUM> may include one or more datum features (not shown) to facilitate the indexing or aligning of the panel assembly <NUM> with the mating structure <NUM> (<FIG>). Once the panel assembly <NUM> is indexed to the mating structure <NUM>, fastener holes <NUM> (<FIG>) are installed at the interface locations <NUM>. Pin elements <NUM> (e.g., tooling pins, temporary fasteners, reusable fasteners such as Clecos™, undersize fasteners, full-size fasteners, etc.) are installed in the fastener holes <NUM> (<FIG>) at the interface locations <NUM> between the panel assembly <NUM> and the mating structure <NUM>, in a manner causing the geometric shape of the panel assembly <NUM> to transition into the nominal state <NUM> (<FIG>).

Referring to <FIG>, as an alternative to the drill-on-assembly process described above, the method <NUM> includes drilling a pattern of fastener holes <NUM>, datum features, index holes <NUM>, and/or pilot holes in the panel assembly <NUM> while supporting the panel assembly <NUM> in the free state <NUM> using the holding fixture <NUM>. For example, the method <NUM> includes drilling a pattern of undersized fastener holes or full-size fastener holes at the interface locations <NUM> of the panel assembly <NUM>. In the example of <FIG>, the panel assembly <NUM> is drilled using a drilling device <NUM> supported by a robotic arm <NUM> of a robotic device <NUM> that is movable along a track <NUM>, similar to the arrangement shown in <FIG>. However, the drilling device <NUM> may be movable via a gantry system (not shown), or other automated and/or programmable controlling device. The drilling device <NUM> is configured to drill fastener holes <NUM>, index holes <NUM>, pilot holes, and/or other datum features in the panel assembly <NUM>.

Examples of datum features include keyholes, slots, grooves, or any other type of indexing feature for aligning the panel assembly <NUM> with the mating structure <NUM> (<FIG>). The fastener holes <NUM> drilled into the panel assembly <NUM> are configured to align with fastener holes (not shown) pre-installed in the mating structure <NUM>. When the panel assembly <NUM> is in the nominal state <NUM>, the fastener holes <NUM> in the panel assembly <NUM> are configured to align with fastener holes <NUM> in the mating structure <NUM>, as described in greater detail below. Mechanical fasteners <NUM> are installed in the fastener holes <NUM> at the interface locations <NUM> to thereby attach the panel assembly <NUM> to the mating structure <NUM>.

The method <NUM> optionally includes developing a free state NC hole-drilling program <NUM> (<FIG>) for drilling a pattern of fastener holes <NUM> into the panel assembly <NUM> while supported in the free state <NUM> via the holding fixture <NUM>. The free state NC hole-drilling program <NUM> may be generated by adjusting an existing nominal state NC hole-drilling program (not shown) originally configured for drilling a pattern of fastener holes <NUM> in the panel assembly <NUM> when in the nominal state <NUM>. Similar to the above-described process for generating the free state NC machining program <NUM>, the nominal state NC hole-drilling program is adjusted by an amount reflecting differences between the free state outer surface contour <NUM> and the nominal state outer surface contour <NUM> of the skin panel <NUM> in the nominal state <NUM>. For example, the adjustment of the nominal state NC hole drilling program comprises adjusting the three-dimensional location and three-dimensional orientation of the hole centerline of each fastener hole. The drilling device <NUM> is configured to drill the pattern according to the free state NC hole-drilling program <NUM> while the panel assembly <NUM> is supported in the free state <NUM> by the holding fixture <NUM>. The pattern is drilled in a manner such that when the panel assembly <NUM> is moved into the nominal state <NUM>, the fastener holes <NUM> in the panel assembly <NUM> will align with the fastener holes <NUM> in the mating structure <NUM>.

Referring to <FIG>, in addition to machining and optionally drilling the panel assembly <NUM>, the method <NUM> may further include trimming the skin panel <NUM> by moving a trimming device <NUM> along final trim lines <NUM> of the skin panel <NUM> while the panel assembly <NUM> is supported in the free state <NUM> via the holding fixture <NUM>. In the example shown, the final trim lines <NUM> are located inboard of the perimeter edges <NUM>. The panel assembly <NUM> may be trimmed using a trimming device <NUM> supported by a robotic arm <NUM> of a robotic device <NUM> movable along a track <NUM>. However, the trimming device <NUM> may be moved by a gantry system (not shown) or other suitable means, such as an automated and/or programmable controlling device. Trimming of the panel assembly <NUM> may include forming one or more openings (e.g., access holes, inspection holes - not shown) in the skin panel <NUM>. During trimming, the panel assembly <NUM> remains supported by the holding fixture <NUM> by narrow tabs <NUM> located at spaced intervals along the final trim lines <NUM>. After the panel assembly <NUM> is removed from the holding fixture <NUM>, the tabs <NUM> are severed to thereby separate the trimmed-off portions from the trimmed skin panel <NUM>.

The method <NUM> may optionally include developing a free state NC trimming program <NUM> (<FIG>) for trimming the panel assembly <NUM> while supported in the free state <NUM> using the holding fixture <NUM>. Similar to the above-described process for developing the free state NC hole-drilling program <NUM>, the free state NC trimming program <NUM> may be developed by adjusting an existing nominal state NC trimming program (not shown) originally configured for trimming the panel assembly <NUM> when in the nominal state <NUM>. The nominal state NC trimming program is adjusted by an amount reflecting differences between the free state outer surface contour <NUM> and the nominal state outer surface contour <NUM> of the skin panel <NUM>.

After the panel assembly <NUM> is removed from the holding fixture <NUM> after machining, drilling, and (optionally) trimming, the method <NUM> further includes indexing the mating structure <NUM> and the panel assembly <NUM> to each other, using one or more datum features, such as index holes <NUM> formed in the panel assembly <NUM> and/or in the mating structure <NUM>. Indexing the mating structure <NUM> and the panel assembly <NUM> to each other comprises either: indexing the mating structure <NUM> to the panel assembly <NUM>, or indexing the panel assembly <NUM> to the mating structure <NUM>. The mating structure <NUM> is provided with fastener holes <NUM> (<FIG>) at interface locations <NUM> on the mating structure <NUM>. Once the panel assembly <NUM> and the mating structure <NUM> are indexed to each other, the method <NUM> includes installing pin elements <NUM> (e.g., tooling pins, temporary fasteners, full-size fasteners, etc.) in the fastener holes <NUM> at the interface locations <NUM> to couple the panel assembly <NUM> to the mating structure <NUM>. The mating structure <NUM> is preferably built to nominal dimensions. As a result, the process of fastening the panel assembly <NUM> to the mating structure <NUM> at the interface locations <NUM> causes the geometric shape of the panel assembly <NUM> to transition into the nominal state <NUM> (e.g., <FIG>).

Referring to <FIG>, shown is a process for assembling a wing <NUM> of an aircraft <NUM>. As mentioned above, the wing <NUM> is an aerostructure <NUM> having an upper panel assembly <NUM> and a lower panel assembly <NUM>. The process of manufacturing the wing <NUM> includes separately machining, drilling, and (optionally) trimming each of the upper and lower panel assemblies <NUM>, <NUM> while supported in the free state <NUM> by the holding fixture <NUM>. After removal from their holding fixtures <NUM>, the method <NUM> includes indexing the upper and lower panel assemblies <NUM>, <NUM> and the mating structure <NUM> to each other.

In one example of assembling the wing <NUM>, the lower panel assembly <NUM> is supported in a free state <NUM> (<FIG>) by a panel assembly fixture (not shown), and the front spar <NUM>, rear spar <NUM>, and ribs <NUM> are indexed to the lower panel assembly <NUM>. The ribs <NUM> are fastened to the front spar <NUM> and the rear spar <NUM> to thereby form a ladder assembly <NUM> (<FIG>), and pin elements <NUM> (e.g., temporary fasteners, full-size fasteners, etc.) are installed in the fastener holes <NUM> at the interface locations <NUM> between the lower panel assembly <NUM> and the front spar <NUM>, the rear spar <NUM>, and the ribs <NUM>. Although not shown, fastener holes are pre-installed in the spar flanges <NUM> of the front and rear spars <NUM>, <NUM>, and in the rib flanges <NUM> of the ribs <NUM>. The spars <NUM>, <NUM> and ribs <NUM> of the ladder assembly <NUM> are preferably built to nominal dimensions, which results in the lower panel assembly <NUM> gradually conforming to its nominal state <NUM> (<FIG>) as the pin elements <NUM> (e.g., temporary fasteners, full-size fasteners, etc.) are installed at the interface locations <NUM>, as shown in <FIG>. After the spars <NUM>, <NUM> and ribs <NUM> are attached to the lower panel assembly <NUM>, the upper panel assembly <NUM> in the free state <NUM> (<FIG>) is indexed to the ladder assembly <NUM>, and pin elements <NUM> are installed in the fastener holes <NUM> at the interface locations <NUM>, thereby conforming the upper panel assembly <NUM> into its nominal state <NUM> (e.g., <FIG>).

In an alternative example of manufacturing the wing <NUM>, the spars <NUM>, <NUM> and ribs <NUM> are sequentially indexed and attached to the upper panel assembly <NUM> to thereby form the ladder assembly <NUM>, after which the lower panel assembly <NUM> is indexed and attached to the ladder assembly <NUM>. In still another example of manufacturing the wing <NUM>, the front spar <NUM>, the rear spar <NUM>, and the ribs <NUM> are interconnected to form the ladder assembly <NUM>, after which the upper or lower panel assembly <NUM>, <NUM> is indexed and fastened to the ladder assembly <NUM>, followed by indexing and fastening the remaining upper or lower panel assembly <NUM>, <NUM> to the ladder assembly <NUM>. <FIG> shows a portion of the wing <NUM> structural assembly <NUM> after installation of the fasteners <NUM> into the fastener holes <NUM> at the interface locations <NUM>. Also shown is the post-machined sacrificial material <NUM>, resulting in nominal thicknesses <NUM> at each interface location <NUM>.

As mentioned above, because the front spar <NUM>, the rear spar <NUM>, and the ribs <NUM> are built to nominal dimensions, and because the panel assembly <NUM> (i.e., the sacrificial material <NUM>) is machined to nominal thicknesses <NUM> at each interface location <NUM>, the occurrence of gaps between the panel assembly <NUM> and the rib <NUM> and spar flanges <NUM> is reduced or eliminated, which reduces or eliminates the need for shimming, as is typically required in conventional manufacturing and assembly methods. Any gaps that do occur are preferably within design allowances, such that shimming is unnecessary. A further advantage of machining each panel assembly <NUM> to its nominal thicknesses <NUM> is that all panel assemblies <NUM> can be interchangeably used with any mating structure <NUM> of the same configuration. For example, in the case of a wing <NUM>, the ability to machine each upper panel assembly <NUM> and lower panel assembly <NUM> to nominal thicknesses <NUM> allows for the interchangeability of the upper and lower panel assembly <NUM>, <NUM> with any ladder assembly <NUM> of the same configuration, such that no panel assembly is limited to use on a single production unit.

In addition to the interchangeability of panel assemblies <NUM> and reducing the need for shimming, the presently-disclosed method <NUM> results in structural assemblies <NUM> (e.g., wings <NUM>, horizontal stabilizers <NUM>, etc.) that have highly accurate (i.e., nominal) surface contours. In aircraft production, the ability to produce highly accurate surface contours translates into improved aerodynamic performance of the aircraft <NUM>. For example, the ability to manufacture the wings <NUM> to an as-designed aerodynamic contour reduces or eliminates the occurrence of drag-generating discontinuities (e.g., steps) that may otherwise occur in the outer mold line (OML). In addition to aerodynamic performance benefits, the use of sacrificial material <NUM> in the presently-disclosed method <NUM> results in significant savings in manufacturing costs and production flow time. For example, the ability to perform the steps of machining, trimming, and drilling in one tool setup, without unloading the panel assembly <NUM> from the holding fixture <NUM> and without changing the orientation of the panel assembly <NUM>, results in significant savings in manufacturing costs and production flow time.

Claim 1:
A method (<NUM>) of manufacturing a panel assembly (<NUM>), comprising:
supporting a panel assembly (<NUM>) in a free state (<NUM>) using a holding fixture (<NUM>) in which the panel assembly (<NUM>) has a geometric shape that is different than the geometric shape of the panel assembly (<NUM>) in an as-designed nominal state (<NUM>), the panel assembly (<NUM>) comprising a skin panel (<NUM>) and sacrificial material (<NUM>) coupled to a skin panel inner surface (<NUM>) respectively at interface locations (<NUM>) for coupling the panel assembly (<NUM>) to a mating structure (<NUM>);
acquiring a free state outer surface contour (<NUM>) of the panel assembly (<NUM>) by scanning a skin panel outer surface (<NUM>) of the skin panel (<NUM>) while the panel assembly (<NUM>) is supported by the holding fixture (<NUM>);
developing a numerically controlled 'NC' machining program (<NUM>) having cutter paths (<NUM>) configured for machining the interface locations (<NUM>) to an inner surface contour (<NUM>) that reflects nominal thicknesses (<NUM>) of the panel assembly (<NUM>) based off of the free state outer surface contour (<NUM>); and
machining the sacrificial material (<NUM>) at the interface locations (<NUM>) by moving a cutter (<NUM>) along the cutter paths (<NUM>) while the panel assembly (<NUM>) is supported by the holding fixture (<NUM>), and while the cutter (<NUM>) is backed by a backing device (<NUM>) applying backing pressure against the skin panel outer surface (<NUM>),
wherein the backing device moves in a coordinated manner in alignment with the cutter as the cutter moves along the cutter paths on the opposite side of the skin panel.