Thermally configurable structural element, in particular for aircraft components

Abstract

The invention relates to a thermally configurable structural element particularly suitable for aircraft components. Specifically, a thermally configurable structural element (e.g., an aircraft component such as an aircraft winglet) capable of assuming at least a first structural configuration and a second structural configuration is provided, whereby the structural element includes an integral actuation mechanism, the integrated actuation mechanism is provided by at least one thermally configurable region and at least one non-thermally configurable region integrally adjacent to the at least one thermally configurable region. The at least one thermally configurable region is capable of assuming at least a first positional orientation and a second positional orientation in response to whether a thermal input is present, thereby causing the structural element to assume the at least a first structural configuration and a second structural configuration, respectively. The thermally configurable regions and non-thermally configurable regions may be formed from sintered shape memory alloy (SMA) particles and sintered non-SMA particles formed by an additive layer manufacturing (ALM) process, such as 3D printing.

Description

[0001] Division Application Description

[0002] This application is a divisional application of a Chinese patent application with an application date of November 25, 2019, an application number of 201911163200.8, and an invention title of "Thermally Configurable Structural Elements Particularly Suitable for Aircraft Components". Technical Field

[0003] The embodiments disclosed herein generally relate to thermally configurable structural elements that are adapted to and capable of assuming at least a first structural configuration and a second structural configuration. In a particularly preferred form, the thermally configurable structural elements can be implemented in aircraft components such as wing structures (such as winglets). Background Art

[0004] Several types of structural members have traditionally been designed to withstand various operating loads during their service life, which may promote conflicting goals when obtaining better operating performance. For example, an aircraft wing with a large span, while presenting good range and fuel efficiency, lacks maneuverability and has a relatively low cruise speed. On the other hand, an aircraft wing with a low aspect ratio, while being able to provide a faster cruise speed and higher maneuverability, presents low aerodynamic efficiency. Therefore, modifying the aerodynamic configuration of such aircraft wings can allow the design of aerodynamic structures with a larger operating envelope and aerodynamic efficiency.

[0005] Conventional actuators that can be used to modify the aerodynamic characteristics of aircraft components typically use hydraulic, pneumatic, or electric motors to actuate configurable structural members. However, due to the pressure, stress, or the number of moving parts, such conventional actuator systems may be heavy and prone to failure over time. Some alternatives to conventional actuators use thermally actuated shape memory alloys (SMAs) to transition a structural member between at least two different stable configurations. Generally, an SMA has two preformed stable configurations based on a phase change of the material at a specific temperature. As demonstrated by U.S. Patent Nos. 7,744,038; 8,110,050; 8,348,201; and 8,726,652 (the entire content of each such prior patent is hereby expressly incorporated by reference), the preformed stable configurations of structural members formed by SMAs have thus been adopted as actuators to cause these members to assume different configurations in response to a thermal input.

[0006] Typically, traditional SMA actuator components are limited to relatively simple structural shapes that can be fabricated from traditional SMA billets, rods, fibers, and bar stock shapes, such as rods, bars, tubes, plates, cables, etc. As a result, there has not been provided a relatively complex structural member having a so-called "intelligent" thermally configurable orientation that allows the structural member to deform during use as required.

[0007] Accordingly, it would be highly desirable if complex structural members could be provided with SMA actuators to allow the structural members to assume various structural orientations in response to thermal input. The embodiments described herein are precisely for meeting such a need. SUMMARY OF THE INVENTION

[0008] The embodiments described herein generally relate to providing a thermally configurable structural element (e.g., an aircraft member such as an aircraft wing spar) capable of assuming at least a first structural configuration and a second structural configuration, wherein the structural element includes an integral actuation mechanism provided by at least one thermally configurable region and at least one non-thermally configurable region, the at least one non-thermally configurable region being integrally adjacent to the at least one thermally configurable region. The at least one thermally configurable region is capable of assuming at least a first position orientation and a second position orientation in response to the presence or absence of thermal input, thereby causing the structural element to assume the at least first structural configuration and the second structural configuration, respectively. The thermally configurable region and the non-thermally configurable region may be formed of sintered shape memory alloy (SMA) particles and sintered non-SMA particles, the sintered shape memory alloy (SMA) particles and sintered non-SMA particles being formed by an additive layer manufacturing (ALM) process such as 3D printing.

[0009] According to certain exemplary embodiments, the at least one thermally configurable region and the non-thermally configurable region may be composed of 3D laser sintered shape memory alloy (SMA) particles and 3D laser sintered non-SMA particles, respectively. The sintered SMA particles are composed of a Ni-Ti based alloy and / or a Cu-based alloy, e.g., a shape memory alloy selected from the group consisting of Ni-Ti, Ni-Al, Cu-Zn, Cu-Zn-Al, Cu-Zn-Sn, Cu-Zn-Si, Cu-Zn-Ga, Cu-Cu-Sn, Au-Cd, Fe-Pt, Mg-Cu, Fe-Mn-Si-Cr-Ni. The sintered non-SMA particles may be formed of almost any non-SMA material commonly employed in the aerospace industry, such as aluminum alloy, magnesium alloy, and / or titanium alloy. By way of example, when implementing a thermally configurable structural member in an aircraft member, the sintered SMA particles may be formed of a shape memory Ni-Ti alloy, while the sintered non-SMA particles may be formed of a non-shape memory aluminum alloy.

[0010] These and other aspects and advantages of the present invention will become more apparent after a careful consideration of the following detailed description of the preferred exemplary embodiments of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0011] The disclosed embodiments of the present invention will be better and more fully understood by reference to the following detailed description of the exemplary non - limiting schematic embodiments in conjunction with the accompanying drawings:

[0012] Figure 1 is a perspective view of an exemplary aircraft in which a thermally configurable structural element can be implemented as a component part of an aircraft winglet;

[0013] Figure 2A and Figure 2B are enlarged views of the aircraft winglet depicted in a first structural configuration and a second structural configuration, respectively, in Figure 1 ;

[0014] Figure 3 is an enlarged perspective view of a portion of the aircraft winglet depicted with the wing skin removed;

[0015] Figure 4 is a detailed front view of a representative spar employed in the aircraft winglet; and

[0016] Figure 5 is Figure 4 an enlarged schematic view of a portion of the spar shown in DETAILED DESCRIPTION

[0017] In Figures 1 to 5 an exemplary and non - limiting embodiment of the present invention described herein is depicted in the form of an angle - adjustable winglet 10 associated with an aircraft wing 12 of an aircraft AC. Under certain flight conditions (e.g., cruise flight versus takeoff and landing), it may be desirable for the winglet to assume at least a first angular orientation and a second angular orientation relative to the wing 12, as shown in Figure 2A and Figure 2B . To enable the winglet 10 to move between different angular orientations (as indicated by the angles α and β in Figure 2B ) relative to the longitudinal and / or chordwise extent of the aircraft wing 12, a spar 20 is provided to allow the winglet to bend and / or twist relative to the oncoming airflow.

[0018] As in Figure 3 and Figure 4As shown in more detail below, each spar in spar 20 will preferably include at least one thermally configurable region 20-1. In the depicted embodiment, spar 20 will also respectively include an inner non-thermally configurable beam region 20-2 and an outer non-thermally configurable beam region 20-3 that are integrally contiguous with the thermally configurable region. Thus, spar 20 is implemented in a one-piece structural member that may include one or more corresponding SMA regions (e.g., corresponding to region 20-1), with one or more corresponding SMA regions implemented integrally contiguous with adjacent non-SMA regions (e.g., corresponding to region 20-2). Accordingly, in certain embodiments, thermally configurable region 20-1 is formed of SMA material, while non-thermally configurable beam regions 20-2, 20-3 are formed of non-SMA material. Of course, it will be understood that although only one thermally configurable region 20-1 is depicted in exemplary embodiments in conjunction with inner region 20-2 and outer region 20-3 respectively, different numbers and configurations of such regions may be provided depending on the particular structural member design and its intended function. Figure 3 and Figure 4 In the exemplary embodiments of, only one thermally configurable region 20-1 is respectively depicted in conjunction with inner region 20-2 and outer region 20-3, but different numbers and configurations of such regions may be provided depending on the particular structural member design and its intended function.

[0019] As Figure 5 shown in, a system is provided to deform the thermally configurable beam region 20-1 of spar 20 in a controlled manner to allow its movement, as previously described. In the system depicted in Figure 5 includes one or more thermal actuators 22 that are operably associated (e.g., in thermal contact) with thermally configurable beam region 20-1 (e.g., in thermal contact with an internal beam component between beam flanges as depicted). Thermal actuator 22 can be any conventional actuator that transfers heat energy to the SMA material forming thermally configurable region 20-1, for example, using on-board electrical energy of the aircraft AC. In a particularly preferred embodiment, the actuator may be implemented in the actuator as more fully disclosed in the pending U.S. patent application Ser. No. 15 / 910,584, filed Mar. 2, 2018, the entire content of which is expressly incorporated herein by reference.

[0020] A plurality of sensors 24 are provided to be operably associated with thermally configurable beam region 20-1 to determine temperature and / or position parameters of thermally configurable region 20-1 and output these parameters to a multi-channel or multiplexed variable voltage controller MVVC. Depending on flight conditions and / or flight profile, manual input (e.g., pilot initiated) or automatic input (e.g., initiated based on an aircraft performance data computer) is provided to the MVVC. Accordingly, MVVC outputs a signal to thermal actuator 22 to cause the thermal actuator to rapidly and controllably heat the thermally configurable region to achieve the desired angular orientations α and β as described above.

[0021] The SMA material forming the thermally configurable region 20-1 is typically a metallic alloy that exhibits a structure deformation triggered by heating (i.e., one-way shape memory effect) while remembering its original shape. During heating and cooling cycles, the SMA can undergo large deformations without showing residual strain and can recover its original shape through thermal cycling (e.g., shape memory effect). This material behavior is attributed to a material microstructure having two different crystal structures (i.e., austenite and martensite). Austenite is the more crystallographically ordered phase and has a higher modulus, while martensite is the more disordered phase and has a lower modulus. Thus, the SMA material will exist as the long-range ordered austenite phase (parent phase or memory phase) at high temperatures.

[0022] The thermally configurable region 20-1 of the spar 20 can actually be formed from any metallic alloy having shape memory properties. While a relatively large variety of alloys exhibit the required shape memory effect, in the embodiments disclosed herein, ideally only those alloys that can recover large deformations or generate significant restoring forces during shape change are employed. For example, suitable SMAs include Ni-Ti based alloys and Cu-based alloys (such as Cu-Zn-Al, Cu-Al-Ni). More specifically, alloys of Ni-Ti, Ni-Al, Cu-Zn, Cu-Zn-Al, Cu-Zn-Sn, Cu-Zn-Si, Cu-Zn-Ga, Cu-Cu-Sn, Au-Cd, Fe-Pt, Mg-Cu, Fe-Mn-Si-Cr-Ni can be satisfactorily employed in the embodiments disclosed herein.

[0023] One of the most common shape memory alloys is NITINOL ® alloy, which is an alloy of nickel and titanium (Ni-Ti) that was discovered at the U.S. Naval Ordnance Laboratory (NOL) in the 1960s. The acronym NiTi-NOL (or NITINOL ® alloy) is used herein to refer to Ni-Ti based shape memory alloys in which the mixture of nickel and titanium alters the material response.

[0024] To provide the desired deflection response characteristics (actuation range) of the beam 20 when heated, the thermally constructible region 20-1 can consist only of SMA (i.e., formed from 100 wt% SMA material (0 wt% of other alloys)), or can include SMA material in addition to at least one other alloy material. Such other alloy can be a non-SMA material (i.e., a material that does not exhibit shape memory characteristics) as long as the thermally actuated shape memory characteristics of the region 20-1 are not adversely affected. Alternatively, the other alloy can be an alloy material that exhibits a superelastic (SE) effect. Although such SE alloys are known and can be formed from the same alloys as those described above in this document that provide a shape memory effect (e.g., NITINOL ® alloy), different heat treatments are performed according to techniques known to those skilled in the art so that the alloy can exhibit the SE effect (i.e., the two-way shape memory effect). Thus, while the SMA material is thermally actuated, the SE alloy material exhibits mechanical superelasticity in response to strain. At low temperatures, the SE alloy material will exist in the martensite form, while at high temperatures, substantially at high temperatures, the SE alloy material exists in the austenite form. At a temperature just above the transformation temperature of the SE alloy material to austenite, the applied stress can transform the austenite into martensite, causing the material to exhibit increased strain under a constant applied stress (i.e., a relatively large deformation for a relatively small applied stress). When this stress is removed, the martensite of the SE alloy material will revert to austenite, and the material returns to its original shape.

[0025] Therefore, the SE alloy material can recover from large structural displacements without being thermally actuated. Moreover, such SE alloy material may have a memorized shape in the bending position such that when the temperature of the SMA material is changed (thereby changing its modulus of elasticity), the balance between these SMA materials and the SE alloy material will determine the actuation range of the region 20-1.

[0026] Depending on the particular structural member and the mechanical deflection characteristics that may be desired, SMA materials and other alloy materials (e.g., SE materials) can be provided in region 20-1 as a homogeneous mixture with one another or in the form of discrete regions (e.g., islands of SE alloy material in a sea of SMA material). By way of example, based on the total weight of the SMA material in region 20-1, e.g., from about 5 wt% to about 50 wt%, about 10 wt% to about 40 wt% or about 20 wt% to about 30 wt%, the SE alloy material can be present in region 20-1 in an amount from 0 wt% to about 60 wt%. Thus, in order to achieve the desired structural deflection of a structural member including at least one region of region 20-1, the SE alloy can be present in any amount deemed necessary based on the total weight of the SMA material in region 20-1, e.g., less than about 60 wt%, less than about 50 wt%, less than about 40 wt%, less than about 30 wt%, less than about 20 wt% or less than about 10 wt%.

[0027] The non-SMA materials forming the non-thermally actuated beam regions 20-2, 20-3 can be any alloy that does not exhibit a shape memory effect. Thus, in practice, any alloy conventionally employed in the aerospace industry can be used for this purpose, such as alloys of aluminum, titanium, magnesium, etc.

[0028] To achieve a complex one-piece structure, most preferably, the spar 20 is formed by an additive layer manufacturing (ALM) process (e.g., 3D printing). In this regard, for example, as more fully described in U.S. Patent Nos. 9,388,078 and 10,065,240 (the entire contents of each such previously issued patent are hereby expressly incorporated by reference), the metal powders of the SMA material and the non-SMA material can be "printed" using laser sintering according to the computer-aided design of the structural member being manufactured. As is known, a complete structural member can be designed using a 3D computer model, which then assists the ALM process in a layer-by-layer additive manner. That is, a thin layer of alloy powder can be spread onto a support tray for subsequent laser sintering based on the first slice of the computer-aided 3D model. Then, subsequent layers corresponding to subsequent slices of the 3D model can be laser sintered in a similar manner until the complete structural member is fabricated. Thus, as can be appreciated, depending on the design of the spar 20 during the ALM process, regions 20-1, 20-2, and 20-3 can be fabricated by providing different powders of the SMA material and the non-SMA material.

[0029] When NITINOL is employed ®When using SMA, a desired shape can be imparted to the thermally constructable region 20-1 by annealing and / or isostatic pressing at a relatively high temperature (e.g., above 700 °C). The region 20-1 formed of NITINOL ® SMA can thus be trained to the desired shape at about 500 °C for at least 25 minutes. For example, the completed 3D printed beam 20 can be manufactured as a one-piece (integral) structural member using the laser sintered powders of the SMA material and non-SMA material as described above, and then thermally trained by annealing and / or isostatic pressing at a high temperature. If an SE alloy material is combined with the SMA material for the thermally constructable region, the same alloy can be employed, but the 3D laser sintering parameters are to be appropriately manipulated so as to impart SE characteristics to certain sintered layers, elements, or regions forming the thermally actuated region 20-1. For example, different temperatures and / or heat treatment times can be employed so as to trigger different crystal structures within the region 20-1, thereby allowing the shape memory or superelastic effect to be achieved.

[0030] Various modifications within the capabilities of those skilled in the art can be envisioned. Accordingly, although the invention has been described in connection with the presently considered most practical and preferred embodiments, it is to be understood that the invention is not limited to the disclosed embodiments, but rather, the invention is intended to cover various modifications and equivalent arrangements included within its spirit and scope.

Claims

1. A thermoformable winglet that is capable of assuming at least a first angular orientation and a second angular orientation relative to an aircraft wing, wherein, The winglet includes a winglet spar having an integral actuation mechanism, the integral actuation mechanism including: at least one thermally constructable region formed of a sintered shape memory alloy; and at least one non-thermally constructable region formed of a non-shape memory material, the at least one thermally constructable region being integrally adjacent to the at least one non-thermally constructable region, wherein the sintered shape memory alloy is an alloy of nickel and titanium in addition to sintered superelastic alloy particles, wherein the amount of the superelastic alloy particles is less than 60% by weight, the winglet spar is formed by an additive manufacturing process and is fabricated as a one-piece structural member using sintered powder of a shape memory alloy in combination with superelastic alloy particles and a non-shape memory material, the winglet spar has one or more thermal actuators and a plurality of sensors, the thermal actuators and sensors being operably associated with the at least one thermally constructable region to transfer thermal energy to the shape memory alloy and respectively determine temperature and / or position parameters of the at least one thermally constructable region, such that the winglet can assume at least a first angular orientation and a second angular orientation in response to the presence or absence of a thermal input.

2. The thermally configurable winglet according to claim 1, wherein, The sintered shape memory alloy particles are shape memory alloys selected from the group consisting of Ni-Ti, Ni-Al, Cu-Zn, Cu-Zn-Al, Cu-Zn-Sn, Cu-Zn-Si, Cu-Zn-Ga, Cu-Cu-Sn, Au-Cd, Fe-Pt, Mg-Cu, Fe-Mn-Si-Cr-Ni.

3. The thermally configurable winglet according to claim 1, wherein, The sintered non-shape memory alloy particles are non-shape memory alloys selected from the group consisting of aluminum alloys, magnesium alloys, and titanium alloys.

4. The thermally configurable winglet according to claim 1, wherein, Based on the total weight of the shape memory alloy particles, the amount of the superelastic alloy particles is from 5% to 50% by weight, or from 10% to 40% by weight, or from 20% to 30% by weight.

5. The thermally constructible winglet according to claim 1, wherein, The sensor outputs the position parameter of the at least one thermally constructable region to a multiplexed variable voltage controller, which outputs a signal to the one or more thermal actuators.

6. An aircraft, the aircraft including an aircraft winglet according to claims 1 to 5.

7. The aircraft according to claim 6, further including: an aircraft wing, wherein the winglet is at the end of the wing; and a thermal control system operably connected to the winglet spar to provide a thermal input to the winglet spar, thereby enabling the winglet to move controllably between a first angular orientation and a second angular orientation of the winglet.

8. A method of manufacturing a thermo-constructible fin according to any one of claims 1 to 5, wherein, The method includes: (i)sintering a layer of shape memory alloy particles in an additive manner in addition to sintered superelastic alloy particles to form at least one thermally constructable region of the structural element; and (ii)sintering a layer of non-shape memory alloy particles in an additive manner to form at least one non-thermally constructable region of the structural element, wherein steps (i) and (ii) can be performed in any order.

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