Patent Application
20090184204
This disclosure generally relates to composite panels, especially those having honeycomb cores, and deals more particularly with a panel construction that is effective in distributing point loads, as well as a method for making the panels.
Aerodynamic features on aerospace vehicles may be formed by lightweight composite panels that are attached to the vehicle airframe. For example, aerodynamic pressure fairings are often formed from large, reinforced honeycomb panels that may be connected together and attached by struts to an airframe. In order to cover large spans, thicker and/or higher density panels are required in order to react loads across the span, however these thicker panels increase the weight of the aircraft.
The honeycomb panels described above may be attached to the airframe by composite laminates that are ramped down along the edges of the panel in order to allow attachment of the edges to the airframe by fasteners. However, ramping down the edges of the honeycomb panel to a solid laminate along the panel edges adds weight to the panel. Moreover, since the entire load on the panel is transferred to the edges, the ramped down laminate section must be thicker in order to carry the load, thereby further adding to vehicle weight.
Accordingly, there is a need for composite panels capable of distributing point loads, as well as a method of making the panels.
Embodiments of the disclosure satisfy the need for lightweight, honeycomb panels capable of distributing point loads. Point loads applied to medial regions of the panel are distributed by a plank that is incorporated in the core and is bonded to or co-cured with a surrounding section of honeycomb. The increased strength and load distribution provided by the plank allows the panel to be at least partially supported by a strut connected between a supporting frame and medial regions of the panel.
According to one disclosed embodiment, a composite panel comprises a core, and laminated plies of composite material bonded to or co-cured with first and second sides of the core. The core includes a first section of honeycomb material and a second section of essentially solid, laminated material surrounded by and bonded to the first section. The solid material may include laminated plies of composite material, and the first and second sections may be bonded by a foam adhesive. The panel may further comprise means for attaching a support to the panel in the area of the section of solid material.
According to another disclosed embodiment, a composite panel for aerospace vehicles comprises: a first, essentially solid laminate core section for distributing point loads across the panel; a second core section bonded to the first core section and having a load bearing capacity less than the first core section; laminated plies of material covering the first and second core sections; and, a load transmitting link connected to the panel in the area of the first core section for supporting the panel. The first core section may include laminated plies of composite material, and the second section of the core may include honeycomb material. The first and second core sections include abutting sides adhesively bonded to each other. The laminated plies may include doubler plies overlying the first core section. The first core section may be located in a medial region of the panel and may be generally circular in shape. The composite panel may further comprise a third core section having a load bearing capacity less than the second core section.
In accordance with another disclosed embodiment, an airplane fairing comprises: at least one panel including inner and outer laminated plies of composite material and a core sandwiched between the inner and outer laminated plies, wherein the core includes a section of increased density located in a medial region of the panel. The section of increased density may include a solid laminate plank, and the core may further include a honeycomb section bonded to and surrounding the plank. The airplane fairing may further comprise a strut for supporting the panel and means for attaching the strut to the panel in the area of the plank.
According to a disclosed method embodiment, fabricating a composite panel for an aerospace vehicle comprises: making a plank; forming a core by positioning the plank within a layer of honeycomb and bonding the honeycomb layer to the plank; forming a layup by placing layers of composite material on opposite sides of the core; compacting the layup; and, curing the layup. The plank may be fabricated by laminating plies of fiber reinforced resin. Bonding the honeycomb layer to the plank may be performed by introducing an adhesive foam between the sides of the honeycomb layer and the sides of the plank. The method may further comprise installing a strut attachment on the panel in the area of the plank.
Other features, benefits and advantages of the disclosed embodiments will become apparent from the following description of embodiments, when viewed in accordance with the attached drawings and appended claims.
Referring first to
In one embodiment, shown in
The adjustable panel attachment fitting assembly 44 includes a nut plate 48 that is secured to the inboard face of the panel 22 by fastener bolts 51 which pass through the panel 22 and are held by threaded retainer nuts 50. A threaded eyebolt 46 is connected to the link 26a by the pin 38 and clevis 34, and is threadably received within a threaded, cylindrical body 48a forming part of the nut plate 48. As will be described in more detail below, removal of the bolts 51 allows the nut plate 48 to be rotated in either of two rotational directions, in order to adjust the axial position of the link 26a.
Referring now to
In order to protectively enclose the through-hole 60 as well as the fastener bolts 86, a circular cover 90 is provided, which may be formed of a flexible, but durable material such as nylon. The cover 90 includes a tubular portion 92 that is received within sleeve 88. A retainer bolt 92, which also may comprise a nylon material, passes through the center of the cover 90 into the key-like opening 48c in order to hold the cover 90 against the outer surface of the panel 22.
In accordance with the disclosed embodiments, point loads imposed on the panel 22 resulting from the attachment of the link 26a to medial regions 28 are laterally distributed through at least a portion of the panel 22. The medial regions 28 of the panel 22 are effectively reinforced by the plank 56 which forms part of the core 62 of the panel 22. The plank 56 may comprise, for example, a solid, rigid material formed by laminated plies of a composite material such as fiber reinforced resin, i.e., a solid laminate. The use of a plank 56 formed from a solid laminate allows the panel 22 to carry higher out-of-plane shear loads and bending loads compared to other types of core constructions. In the illustrated example, the plank 56 is circular in shape and is slightly larger in diameter than the diameter of the nut plate 48. Other geometries, however, are possible, although a circular geometry aids in uniformly transferring moments in any radial direction through the plane of the panel 22.
The solid plank 56 is effectively embedded in, and forms an integral part of the honeycomb panel core 62, enabling the link 26a or similar strut/tie-rod to be directly attached to the middle of the panel 22 while remaining capable of reacting large out-of-plane loads. The nut plate 48 can be directly attached to the plank 56 and then adjusted to fit against the panel 22 from outside of the fairing 20 using the adjustable panel attachment fitting 44. The plank 56 may be either pre-cured or uncured (green) when it is inserted into the honeycomb panel core 62. The plank 56 is then co-cured with or co-bonded to the honeycomb core assembly 62 during fabrication of the panel 22.
The core 62 further includes a ring shaped honeycomb section 58 surrounding and attached to the sides of the plank 56. The ring shaped honeycomb section 58 is surrounded by another honeycomb section 52. In one embodiment, the density of the core section 58 is less than that of the plank 56 but greater than the density of the core section 52. For example, in one application providing satisfactory results, the plank 56 comprises 60 laminated plies of fiberglass, core section 58 comprises a heat resistant phenolic honeycomb having a density of eight pounds per cubic foot, and core section 52 is also a heat resistant phenolic honeycomb having a density of three pounds per cubic foot. Where the plank 52 is green (uncured) during the assembly phase, the honeycomb core section 58 may be attached by co-curing plank 56 and core section 52 using a suitable foam adhesive which forms a splice 84 (see
Referring now simultaneously to
Assembly of the core 62 in step 98 may be carried out by forming a first cut-out in honeycomb section 52 having the size and shape of honeycomb section 58, placing section 58 in the cut-out and then bonding the sides of section 52 to the sides of section 58. Similarly, a second cut-out is formed in the honeycomb section 58 that matches the size and shape of the plank 56, following which the plank 56 is placed within the cut-out in section 58. Then, the sides of the plank 56 are attached to the sides of the honeycomb section 58 using a foaming adhesive to complete assembly of the core 62. At step 100, the assembled core 62 is placed on top of the first group of plies 64.
At step 102, the plies in the second group 66 are successively laid up over the assembled core 62. Then, at 104, a suitable release film 80, such as TEDLAR® is placed over the ply group 66. Next, at 106, a caul plate 82 is placed on top of the layup. At 108, the layup is vacuum bagged and compacted, following which the compacted layup and core are co-cured cured at step 110. Then, the through-holes 60, 86 are formed in the panel as shown at 112, following which at step 114, the adjustable panel attachment fitting assembly 44 is secured to the panel 22.
Although the embodiments of this disclosure have been described with respect to certain exemplary embodiments, it is to be understood that the specific embodiments are for purposes of illustration and not limitation, as other variations will occur to those of skill in the art.
