Patent Application
20150090771
This application claims the benefit of the European Patent Application No. 13186578.4 filed on Sep 30, 2013, the entire disclosures of which are incorporated herein by way of reference.
The present invention relates to a method and a system for fabricating a module, such as a stiffened panel, for an airframe or fuselage structure of an aircraft or spacecraft. The module fabricated according to the method and system of the invention comprises parts or members, typically of metal or a metal alloy, which are or may be welded together.
The method and system of the invention are especially suitable for use in fabrication of an airframe or fuselage structure of an aircraft or spacecraft, and it is convenient to describe the invention herein in this exemplary context. It will be appreciated, however, that the method and system of the invention are not limited to this application, but may be used to provide modules, such as panels, for the body structure of other vehicles, such as trains, automobiles, trucks or ships. Accordingly, the method and system of the invention may be suitable for a whole range of nautical, aeronautical, automotive and aerospace applications.
Current airframe and fuselage structures, including wing and tail structures of commercial aircraft, are typically built from panel modules which are joined together with fasteners, such as rivets. In recently developed commercial aircraft, welding has been used to join stiffener members, such as stringers, to skin panels. With current manufacturing techniques, however, the welding of stringers to skin panels often generates distortions (sometimes called the “Zeppelin effect”) which need to be corrected in separate manufacturing procedures. In other words, the welded and stiffened panels need to be transferred or transported to other processing stations for further treatment and this in turn significantly increases the workload and the lead times for the overall manufacturing process of integrally stiffened panel modules in aircraft manufacture.
It is therefore an idea of the present invention to provide a new and improved method and system for overcoming one or more of the problems discussed above. In particular, it would be useful to provide a new method of fabricating a module for an airframe or fuselage structure of an aircraft which may enable a faster and/or more automated production procedure.
According to one aspect, the invention provides a method of fabricating a module for an airframe or fuselage structure of an aircraft or spacecraft, including the steps of positioning a first member adjacent to a second member at a work station or assembly station, welding the first member to the second member at the work station or assembly station to produce the module having a welded joint between the first and second members, and treating the first member, the second member, and/or the welded joint at the work station or assembly station to compensate for, or correct, distortion in the module caused by the welding.
In another embodiment, the step of treating the first member, the second member, and/or the welded joint at the work station or assembly station includes laser peening (also called “laser shock peening” (LSP) or “laser shot peening”). That is, the first member, the second member, or the welded joint may be laser peened. Laser peening is a process of hardening or peening metal using a powerful laser and typically involves generating high amplitude shock waves which are applied to a selected region of the module. The peak pressure of the shock waves typically exceeds the dynamic yield strength of a material of the module, i.e. the material of at least one of the first member, the second member, and/or the welded joint to which the shock waves are applied, and this in turn produces the laser peening effect. In a preferred embodiment, therefore, each of the first member and the second member is comprised of metal or metal alloy, such as an aluminium alloy. The laser peening may preferably also be employed to induce compressive stresses in the module, which may inhibit fatigue-induced weaknesses in the material. Laser peening can impart a layer of residual compressive stresses in a surface that is four times deeper than that which is attainable from conventional shot peening treatments.
In another embodiment, the step of treating the first member, the second member, and/or the welded joint at the work station or assembly station to compensate for, or to correct, distortion caused by welding includes generating residual compressive stresses in a surface region of the module; i.e. a surface region of the first member, the second member, and/or the welded joint. To this end, the technique employed to generate the residual compressive stresses in a surface region may comprise peening, including any one or more of laser peening, shot peening, and sand-blasting. While laser peening provides for a highly precise treatment of a surface region of the module, the other techniques, such as shot peening or sand-blasting, can be performed substantially faster.
In another embodiment, the first member of the module is a stiffening member. For example, the first member may be an elongate member having a transverse cross-section or profile that may be e.g. I-shaped, C-shaped, Z-shaped, T-shaped, or L-shaped. The second member of the module, on the other hand, preferably comprises an area member, such as a sheet member or a panel, which may define a multi-dimensional surface area, such as a two- or three-dimensional surface area. As will be appreciated by persons skilled in the art, the module may include a number of first members (e.g. stiffening members) welded to the second member (e.g. an area member, sheet member, or panel).
The conventional riveting procedures used to construct and also to interconnect panel modules or shell sections in airframe and fuselage structures carries a significant weight penalty due to the need to overlap material along the riveted seams, and due to the inclusion of doublers, crack stoppers, and sealing material, not to mention the rivets themselves. Furthermore, the use of rivets is an expensive procedure despite the fact that the riveting procedures today are highly automated. As an alternative to riveting, welding has been contemplated. Conventional welding methods, however, generate distortions in the welded parts, especially in sheet members or panels, due to residual stresses induced in the material by thermal stresses during formation of the welded joint and/or due to a change in the microstructure of the material at the welded joint. These distortions in the welded parts thus require additional manufacturing operations which, in turn, result in greater lead times in airframe and fuselage structure assembly.
The present invention replaces the conventional manufacturing techniques with a method that is able to save weight and cost at same time. In particular, the present invention improves conventional assembly procedures in terms of cost reduction and time efficiency by correcting the distortion created by welding during the same manufacturing operation without generating additional lead time. That is, both the welding step and the peening step take place at the same work station or assembly station, and no separate manufacturing operations are required with the method of the invention.
In another embodiment, therefore, the invention provides a method of fabricating a panel module for an airframe or fuselage structure of an aircraft or spacecraft, including the steps of positioning a stiffening member on one side of a panel so that the stiffening member extends over said one side of the panel at a work station, welding the stiffening member to the panel at the work station to produce a panel module including a welded joint between the stiffening member and the panel, and laser peening the welded joint and/or the panel at the work station to compensate for, or correct, distortion in the stiffened panel module caused by the welding.
In another embodiment, the welding of the first member to the second member includes at least one of laser beam welding (LBW) and friction stir welding (FSW). Furthermore, the step of welding preferably involves butt welding or T-joint welding the first member to the second member. It will be understood that the method may include the step of clamping the first and second members together during the welding step to ensure proper positioning of the first member and the second member is maintained.
In another embodiment of the invention, the laser peening includes imparting laser beam pulses to one or more of the first member, the second member, and the welded joint through an overlay or coating. A coating, usually black tape or paint, may be applied to absorb the laser energy. Short energy pulses are then focused to explode the ablative coating, producing a shock wave with the laser. The beam is then repositioned and the process is repeated, creating an array of slight indentations of compression and depth with about 5-7% cold work. Preferably, a transparent overlay may also be provided. The transparent overlay may, for example, comprise a layer of liquid, such as water. The liquid may optionally flow over a surface of the module in the region of the laser shock peening. This translucent layer is provided over the coating and acts as a tamp, directing the shock wave into the treated material. In this way, the liquid may effect or perform cooling of the module in the region of the laser shock peening to prevent or limit thermal effects by the laser beam pulses. The laser peening process will typically be computer-controlled and may be repeated, e.g. as many as three times, until the desired compression level is reached, producing a compressive layer as deep as 1-2 mm average.
In another embodiment, the step of laser shock peening is carried out before, during and/or after the welding step. In this regard, the laser shock peening is preferably performed or carried out during the welding step (e.g. after material consolidation) to correct or counter-act the distortion in the first and/or second member caused by the welding. That is, the welding and laser peening (LSP) steps or operations may take place or be carried out substantially simultaneously or contemporaneously. It is to be noted, however, that the laser peening may also be performed or carried out before the welding step to compensate for, or counteract, subsequent deformation which occurs during or after the welding step. Alternatively, or in addition, the laser shock peening may be performed or carried out after the welding step.
In another embodiment of the invention, the method comprises: measuring or sensing a distortion in the module (e.g. in either or both of the first member and second member) during and/or after the welding step; and controlling the laser peening based on the distortion of the module that is measured or sensed. For example, the step of measuring or sensing the distortion in the module may comprise measuring or sensing reaction forces, or changes therein, in a clamping mechanism or device for positioning and holding the first member with respect to the second member during welding. Alternatively, or in addition, the measuring or sensing step may include detecting other property changes, such as surface strain, due to distortion of the first or second member. The size and location of the region to be laser peened may also be predicted based on welding process parameters and geometry details, thereby reducing a need for measurements during the manufacturing process.
According to another aspect, the present invention provides a system for fabricating a module, such as a stiffened panel module, for an airframe or fuselage structure of an aircraft or spacecraft. The system includes a frame for positioning and holding a first member with respect to and adjacent a second member, a welding head for producing a welded joint between the first member and the second member held by the frame, and a treatment head, especially a laser shock peening head, to compensate for or correct distortion in the module caused by welding.
In another embodiment, the welding head includes a laser for laser beam welding (LBW). As the laser shock peening head will typically also include a laser, the laser shock peening head is preferably included in or combined with the welding head. In this regard, the welding laser may be separate or distinct from the shock peening laser.
In another embodiment, the frame includes a clamping device or clamping mechanism for holding the first member clamped or fixed in position with respect to the second member. Thus, where, for example, the first member is a stiffening member and the second member is a sheet member or panel, the clamping device may be configured to hold or fix an elongate stiffening member adjacent one side or facing surface of the sheet member or panel.
In another embodiment, the system further comprises a sensor device or measuring device for sensing or measuring a distortion of either or both of the first and second members during welding. As noted above, for example, the sensor device or measuring device may sense or measure a distortion in the module based on reaction forces, or changes therein, in the clamping mechanism or clamping device with which the first member (e.g. stiffening member) is held with respect to the second member (e.g. panel) during welding. Thus, the sensor or measuring device may be incorporated in the clamping mechanism. Alternatively, or in addition, the sensor or measuring device may sense or measure other property changes, such as surface strain, due to distortion of the panel or stiffener member.
According to a further aspect, the present invention provides a method of fabricating a module for an airframe or fuselage structure of an aircraft or spacecraft, including the steps of positioning a first member adjacent to a second member at a work station or assembly station, welding the first member to the second member at the work station or assembly station to produce the module having a welded joint between the first and second members, and treating the first member, the second member, and/or the welded joint at the work station or assembly station to compensate for, or correct, distortion in the module caused by the welding. The method further provides generating residual compressive stresses in a surface region of the module, and generating a shock wave in a region of at least one of the first member, the second member, and the welded joint.
For a more complete understanding of the present invention and the advantages thereof, exemplary embodiments of the invention are explained in more detail in the following description with reference to the accompanying drawings, in which like reference characters designate like parts and in which:
The accompanying drawings are included to provide a further understanding of the present invention and are incorporated in and constitute a part of this specification. The drawings illustrate particular embodiments of the invention and together with the description serve to explain the principles of the invention. Other embodiments of the invention and many of the attendant advantages of the invention will be readily appreciated as they become better understood with reference to the following detailed description.
It will be appreciated that common and/or well understood elements that may be useful or necessary in a commercially feasible embodiment are not necessarily depicted in order to facilitate a more abstracted view of the embodiments. The elements of the drawings are not necessarily illustrated to scale relative to each other. It will further be appreciated that certain actions and/or steps in an embodiment of a method may be described or depicted in a particular order of occurrences while those skilled in the art will understand that such specificity with respect to sequence is not necessarily required. It will also be understood that the terms and expressions used in the present specification have the ordinary meaning as is accorded to such terms and expressions with respect to their corresponding respective areas of inquiry and study, except where specific meanings have otherwise been set forth herein.
With reference firstly to
Referring to
Referring now to
Undesirable thermal effects caused by the laser shock peening beam 5 in this embodiment can be avoided by providing a fluid overlay 8 which is transparent to the laser beam but confines the plasma generated. For example, the fluid overlay 8 may be in the form of a layer of water flowing over a surface of the panel 3. This liquid layer 8 provided over the coating also acts as a tamp, directing the shock wave 6 into the treated material. The non-uniform residual stress distribution that is generated through a thickness t of the aluminium material in the panel module 1 itself leads to deformation which can be precisely controlled for shape correction of the initially distorted parts. Thus, in this way, the production method of this embodiment creates substantial cost reductions and time efficiencies in that the distortions D created by the welding can be corrected in the same manufacturing operation, i.e. during welded assembly of the panel module 1 (i.e. at the panel module assembly station S or with the fabrication equipment or system) without creating additional lead time and correspondingly increased cost by separate manufacturing operations required to further process the panel modules.
Because both laser beam welding (LBW) and laser shock peening (LSP) are repeatable processes, it is conceivable that the laser shock peening via laser beam 5 would be performed or carried out during and/or directly after the welding on the still clamped structure. Alternatively, or in addition, it is conceivable that the laser shock peening via the laser beam 5 may occur or take place during welding to achieve near-zero distortion after unclamping. In such a case, the laser shock peening could be based upon expected distortion derived from previous experience with the particular welding operation. The control of the laser shock peening could also be based on a model of the welding process, and/or on measurements of the distortion generated in the first or second members 2, 3 of the module 1 during welding; for example, via reaction force measurements in the clamping mechanism or via surface strain in the panel module 1.
Referring now to
Although the embodiments of the method shown schematically in
Although specific embodiments of the invention have been illustrated and described herein, it will be appreciated by those of ordinary skill in the art that a variety of alternate and/or equivalent implementations exist. It should be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration in any way. Rather, the foregoing summary and detailed description will provide those skilled in the art with a convenient road map for implementing at least one exemplary embodiment, it being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope as set forth in the appended claims and their legal equivalents. Generally, this application is intended to cover any adaptations or variations of the specific embodiments discussed herein.
In this document, the terms “comprise”, “comprising”, “include”, “including”, “contain”, “containing”, “have”, “having”, and any variations thereof, are intended to be understood in an inclusive (i.e. non-exclusive) sense, such that the process, method, device, apparatus or system described herein is not limited to those features or parts or elements or steps recited but may include other elements, features, parts or steps not expressly listed or inherent to such process, method, article, or apparatus. Furthermore, the terms “a” and “an” used herein are intended to be understood as meaning one or more unless explicitly stated otherwise. Moreover, the terms “first”, “second”, “third”, etc. are used merely as labels, and are not intended to impose numerical requirements on or to establish a certain ranking of importance of their objects.
As is apparent from the foregoing specification, the invention is susceptible of being embodied with various alterations and modifications which may differ particularly from those that have been described in the preceding specification and description. It should be understood that we wish to embody within the scope of the patent warranted hereon all such modifications as reasonably and properly come within the scope of my contribution to the art.
| Number | Date | Country | Kind |
|---|---|---|---|
| 13186578.4 | Sep 2013 | EP | regional |
