GB2477882A - Energy absorbing structure for an aircraft - Google Patents

Abstract

An energy absorbing structure disposed beneath the cargo area of an aircraft comprises an elongated body 40,64 having first and second opposing ends and sides. The body is fabricated from laminated plies of reinforced composite material with the plies having varying lengths forming steps along the length of the member such that the compression strength of the member is progressively less from the first end to the second end. The body also has at least one reinforcing flange 68 extending from the main body along a first side, the flange tapering from the first end of the body to the second. In use, the energy absorbing structure provides maximum structural strength and, in the event of a crash, energy absorption. The energy absorbing structure uses the ply drop-offs to produce stable, progressive crushing of the stanchions during a wheels-up crash landing. Integrally formed flanges on the stanchions allow direct mounting of the stanchions on the structural members of the aircraft, without the need for specialized fittings.

Description

ENERGY ABSORBING STRUCTURE FOR AIRCRAFT

This invention broadly relates to energy absorbing structures, and deals more particularly with energy absorbing elements such as stanchions and assembly fittings used in aircraft.

An important objective in the design of modern military and civilian aircraft is the protection of occupants in the event of a survivable crash. Crash injuries may be caused by high acceleration loads experienced by the occupants, or the loss of structural integrity of the aircraft. Crashworthy designs for aircraft require the fuselage to provide energy absorption in order to slow down the aircraft during a crash impact. Controlled deceleration of the aircraft during impact reduces inertial loads on the occupants, and assists in maintaining structural integrity of the aircraft. The area beneath the fuselage floor can form an important part of the energy absorption system in the event of a crash, since this area is normally the part of the aircraft that is initially crushed during ground impact.

The present invention may solve one or more of the following problems of earlier generations of crash energy absorption systems. Earlier generations of crash energy absorption systems relied heavily on metal components to absorb aircraft kinetic energy during a crash. However, the increasing use of composite materials in aircraft designs has resulted in newer crash energy absorption systems that rely on components formed of composite materials. Energy absorbing systems that use composite materials present a design challenge because these materials are often brittle and do not exhibit either plasticity or high elongation prior to failure. In order to solve these design problems, a number of solutions have been proposed, including varying the density of stitching used in stitched composite structures, and using so called ply drop-offs to sustain a crushing process after it is initiated. These design solutions may sometimes be difficult to incorporate into manufacturing processes without significant cost increases. In addition, previous design solutions were limited to protecting occupants from impact forces acting vertically on the aircraft, rather than from off-vertical forces which can be imposed on the fuselage if the aircraft rolls slightly before ground impact.

Accordingly, there is a need for energy absorbing structure for aircraft that both improves crashworthiness and is cost effective to implement. The invention is directed towards satisfying this need.

The invention is set out in the independent claims.

Preferred or optional features are set out in the dependent claims.

In accordance with one aspect of the invention, an energy absorbing structural member comprises an elongated body having first and second opposite ends, and first and second opposite sides. The body contains laminated plies of a reinforced composite material wherein the plies have varying lengths forming steps along the length of the member such that the compression strength of the member is progressively less from the first end to the second end. At least one reinforced flange on the body is tapered from the first to the second end, and functions to both reinforce the structural member and promote elastic stability during crushing. The laminated plies are preferably formed from reinforced composite materials such as a carbon fiber reinforced epoxy. In one embodiment, the structural member includes first and second reinforcing flanges extending along the length of the body, and third and fourth flanges integrally formed with the first and second flanges, for attaching the structural member to the floor of an aircraft.

These and other features, aspects and advantages of the invention will become better understood with reference to the following description and claims, together with the accompanying drawings, in which: Figure 1 is a transverse, cross sectional view of an aircraft showing the space beneath the fuselage floor both before and after crushing due to ground impact; Figure 2 is a graph of typical load-deflection test data for crushing composite structures, depicting both stable and unstable crushing; Figure 3 is a graph showing load-deflection test data for the energy absorbing structure of the invention; Figure 4 is a perspective view of a lower section of an aircraft having the energy absorbing structure of the invention; Figure 5 is a front view of the section of the aircraft shown in Figure 4; Figure 6 is a perspective view of a stanchion forming part of the energy absorbing structure; Figure 7 is a side elevational view of the stanchion shown in Figure 6; Figure 8 is a diagrammatic view showing one ply lay-up that may be used in fabricating the stanchion shown in Figure 6; Figure 9 is an enlarged, fragmentary view of the energy absorbing structure, and depicting each of the crush zones in the stanchions; Figure 10 is a perspective view of a metal fitting used for attaching the stanchion to a floor beam; Figure 11 is a perspective view of a composite material fitting used for attaching the stanchion to a floor beam; Figure 12 is a cross sectional view of a floor beam shown in Figures 5 and 9; Figure 13 is a fragmentary, cross sectional view showing the attachment of the stanchions to the floor beam and roller tray; Figure 14 is a front view of the stanchion attachment shown in Figure 13; Figure 15 is an enlarged, fragmentary view showing a single angularly oriented stanchion; Figure 16 is a sectional view taken along the line 16-16 in Figure 15; Figure 17 is a plan view of a lay-up blank for forming a one piece stanchion in accordance with an alternate embodiment of the invention; Figure 18 is a perspective view of a stanchion formed from the blank shown in Figure 17; Figure 19 is an enlarged, fragmentary front view showing the attachment of the stanchion of Figure 18 to a roller tray; Figure 20 is a side view of the stanchion attachment shown in Figure 19; and Figure 21 is a sectional view taken along the line 21-21 in Figure 19.

Referring first to Figure 1, an aircraft generally indicated by the numeral 20 has a cylindrical fuselage 21 defining an interior space that is separated into upper and lower lobes by a floor 23. The upper lobe includes a passenger compartment 22 while the lower lobe contains a cargo area 24. A volume of space below the cargo area 24 is normally available to contain energy absorption structure and constitutes the part of the aircraft 20 that is first to impact the ground in the event of a wheels up landing. The left side of the aircraft 20 shown in Figure 1 depicts the condition of the cargo area 24 before impact, while the right side of Figure 1 shows the aircraft having impacted the ground during a crash, producing a crush area 25 where part of the fuselage 21 and internal support structure collapse. To improve crashworthiness, it is desirable that the structure within the crush area 25 crushes in a stable manner in order to provide the required energy absorption while maintaining the structural integrity of the aircraft 20, to the extent possible. Thus, in designing the energy absorption structure, the structure must meet both minimum crush energy requirements for the particular application, and the design load requirements for the aircraft.

Figure 2 is a graph illustrating the difference in energy absorption between stable and unstable crush events.

Plots 26 and 28 illustrate the displacement of energy absorbing structure as a function of load. As shown by plot 26, the energy absorbing structure experiences early, peak loading which typically results in rapid failure of structural support components, usually in the form of buckling, breaking or shattering. Following failure, the structural support components carry a relatively low level of load, and thus absorb a relatively low amount of energy. In contrast, an energy absorbing structural design that results in stable crushing, indicated by the plot 28, sustains a higher level of loading over a longer duration, thus absorbing a larger amount of energy compared to a design that produces unstable crushing. An energy absorbing structural design producing stable crushing, as exemplified by plot 28, typically exhibits relatively high plasticity and high elongation prior to failure.

Figure 3, which will be discussed later in more detail, is a graph showing the load-deflection test data for crushing of an energy absorbing member in accordance with the invention. The energy absorbing member comprising a later discussed stanchion initially experiences a peak load at 30, as its connection to an outer frame member fails. Next, the stanchion contacts the outer skin of the fuselage 20, resulting in gradual loading of the stanchion as the stanchion progressively crushes.

Referring now also to Figures 4-14, the cargo area 24 includes a plurality of transversely extending floor beams 32 connected at their outer ends to a curved fuselage frame 36, and supported by a plurality of stanchions 40. The opposite ends of stanchions are respectively connected to the floor beams 32 and fuselage frames 36. A plurality of longitudinal beam members in the form of roller trays 24 are supported by the floor beams 32 and extend transversely thereto. The roller trays 24 include rollers 58 (Figure 14) which support cargo containers (not shown) within the cargo area 24.

As best shown in Figures 5 and 9, the stanchions 40 in combination with the fuselage frame 36 and floor beams 32, form energy absorbing structure disposed within the crush zone 25, near the bottom of the fuselage 21.

Each of the stanchions (structural members) 40 is generally C-shaped in cross section and comprises a main, flat body portion 40a, and a pair of spaced apart, generally parallel flanges 40b. The flanges 40b are tapered such that their height at one end of the stanchion 40 that is attached to floor beams 32 narrows to a lesser height at the other end of the stanchion 40 that is attached to the fuselage frame 36.

The stanchions 40 are of unitary construction formed of laminated plies of a reinforced polymer-based material, which may comprise a fiber reinforcement held in a polymer matrix such as a thermosetting or thermoplastic

resin. In the present description, it is to be

understood that "reinforced polymer-based material" includes various non-homogeneous polymer-based materials, commonly referred to as "reinforced composites, carbon-fiber composites" or similar terms used in the art. In one embodiment, stanchions 40 are formed from a layup 44 (Figure 8) comprising multiple plies 46 of epoxy impregnated carbon fiber tape or fabric, fabricated using conventional layup techniques. The plies 46 are arranged so that their respective directions of orientation alternate at different angles in order to obtain the stiffness of the cured laminate. The plies 46 may be arranged in groups 48, 50 having differing lengths. In Figure 8, for example, it can be seen that a first group of plies 48 extends the entire length of the stanchion 40 and comprises a thickness of 7 plies arranged in a 0/+45/90/0/-45/90/+45 configuration. The first group 48 of plies is followed by successive groups 50 each comprising two plies, which are of successively shorter lengths.

The progressively shorter ply groups 48, 50 may be referred to as ply drop-offs, which means that the first end of the stanchion 40 that is connected to the floor beams 32 has maximum ply thickness, and the ply layers 46 progressively drop-off toward the opposite end of the stanchion 40 so that the minimum number of plies 46 occurs at the end of the stanchion 40 that is attached to the fuselage frame 36. These ply drop-off result in the cross sectional area of the stanchion becoming progressively greater along its length. In other words, the cross sectional area of the stanchion 40 is tailored along its length. The tapering of the flanges 40b also contributes to the tailoring of stanchion cross sectional area. As best seen in Figure 9, the piy drop-off arrangement used in stanchions 40 creates a plurality of crush zones 52. Due to the stepped cross sectional areas created by the ply drop-offs, the compressive strength of the stanchion 40 progressively decreases from the end of the stanchion 40 connected to the floor beam 32 to the other end of the stanchion. As will be discussed below, the ply drop-offs creating the successive crush zones 52 results in controlled, stable crushing, rather than buckling or shattering of the stanchion 40 during a crash.

As best seen in Figures 12 and 13, each of the floor beams 32 comprises a central web 32b connecting a flat cap 32a with a lower flange 32c. The base of the roller tray 54 rests on the flat cap 32a. The upper end of each stanchion 40 is secured to one face of a web 32b by means of a pair of fittings 42. One of the fittings 42 is disposed within the channel formed by the stanchion 40, and engages the flat body 40a, in face-to--face contact.

A second fitting 42 engages the opposite face of the web 32b, so that the combination of the stanchion 40 and floor beam 32 are sandwiched between the two fittings 42.

Suitable fasteners such as Hi-Lok® airframe fasteners 56 are used to draw the fittings 42 toward each other, thereby securely fastening the stanchion 40 to the floor beam 32. Each of the fittings 42 further includes an upper flange 42c reinforced by a pair of triangularly shaped side flanges 42b. Additional Hi-Lok® fasteners 56 pass through flange 42c in order to secure the fittings 42 to the roller tray 54. Hy-lock or similar fasteners may be used to attached the lower ends of the stanchions to the fuselage frame 36.

The fittings 42 may be constructed of metal such as aluminum or titanium as shown in Figure 10. Alternately, the fittings 42 may comprise a composite material, such as laminated plies of epoxy impregnated carbon fiber as shown in Figure 11. An area of increased ply thickness 42d may be provided at the upper end of the fitting 42 to provide additional strength.

As best shown in Figures 5, 9 and 15, the stanchions 40 may be arranged such that their central axes 57 form an angle relative to the longitudinal axis of the corresponding floor beam 32. Thus, stanchions 40 are effectively inclined from vertical when the aircraft 20 is level. As best shown in Figure 5, the stanchions 40 are arranged in two groups respectively on opposite sides of the center line of the fuselage 21. As a result of this angular configuration, each of the stanchions 40 is capable of absorbing energy resulting from partial side loads on the fuselage 21 during a crash.

Upon ground impact during a wheels up "belly" landing, the stanchions 40 will fail near the fasteners securing the stanchions 40 to the fuselage frame 36, following which the lower ends of the stanchions 40 are driven by the impact into contact with the fuselage skin (not shown), placing the stanchions 40 in compression. The resulting load deflection occurring during these two events is shown in Figure 3. As previously mentioned, initial loading of the stanchions 40 results in a peak load that rapidly diminishes when the stanchions 40 near the fasteners holding the stanchion 40 to the fuselage frame 36. The load briefly goes to zero until the stanchions 40 contact the fuselage skin, following which the stanchions 40 are then gradually loaded. The plies of the stanchions 40 having the longest length, i.e. ply group 48 (Figure 8) initially bears the full amount of the load, as this comprises the weakest crush zone 52 at the bottom of the stanchion 40.

As the loading on the ply group 48 increases, the individual plies 46 in the ply group 48 develop interlaminer cracks that allow subsequent bending of the individual plies and ultimately, delamination and breaking of those plies. Thus, when the initial group 48 of plies 46 fail, the next group 50 of plies 46 is subjected to the load, resulting in crushing of those plies. This orderly crushing of the stanchion 40, beginning at the bottom of the stanchion 40 and propagating upwardly through successive crush zones 52, results in the stanchion 40 failing in an elastic, controlled manner, rather than buckling or shattering.

As a result, the stanchions 40 are capable of absorbing a higher level of energy from the impact while aiding in maintaining the structural integrity of the aircrafts frame. As the stanchions 40 are crushing, the fuselage frame 36 is also bending, absorbing additional energy.

In some cases, the floor beams 32 may likewise bend to absorb some of the crash energy. Consequently, it can be appreciated that the stanchions 40, fuselage frames 36 and floor beams 32 cooperate and function as an integrated energy absorption system.

The exact number and orientation of the stanchions 40 will vary with the particular application and aircraft design. However, as previously mentioned, in accordance with the invention, the central axes 57 of at least some of the stanchions 40 may be disposed at an angle relative to the floor beam 32 and fuselage frame 36 so as to absorb energy developed by both vertical and horizontal force components resulting from the ground impact. This is particularly important where the aircraft 20 rolls slightly before ground impact.

Attention is now directed to Figures 17-21 wherein an alternate stanchion 62 is depicted which is of one piece, unitary construction, and eliminates the need for the use of the previously described fittings 42. Stanchions 62 include a main body 64 provided with a pair of spaced apart, flared reinforcing flanges 68 at the upper end thereof. The upper ends of the reinforcing flanges 68 are provided with inwardly turned, oppositely facing mounting flanges 66 that extend perpendicular to flanges 68. Alternatively the mounting flanges 66 could be turned outward and attached to the roller tray 24 or other reinforcement member (not shown) in a similar manner.

Stanchion 62 may be formed from material similar to the previous discussed stanchion 40, including ply drop-offs to create crush zones exhibiting progressively less compressive strength near the outer end of the stanchions 62. The upper end of the stanchion 62 may be provided with additional ply layers and a flared flange 68 that increases the structural strength of the upper end of the stanchion 62. The stanchions 62 are provided with through holes to allow direct attachment to a floor beam 32 using fasteners 56 such as Fli_Lok® fasteners Additionally, the inwardly (or outwardly) turned, oppositely facing mounting flanges 66 provide a means of directly attaching the stanchions 62 to the roller tray 34 using fasteners 56. The lower end of the stanchion 62 may be connected to the fuselage frame 36, also using Hi-Lok® fasteners.

Stanchion 62 may be economically produced using ordinary lay-up techniques in which a tailored ply buildup with suitable ply drop-offs is created to provide a master lay-up (not shown). A water jet cutter or diamond saw (not shown) may then be used to cutout multiple individual blanks 60 from the master layup. The blanks, which are either uncured or partially cured, have an outline geometry forming various structural features of the completed stanchion 62 shown in Figure 18. The blank is then placed over a tool (not shown) and subjected to pressure and elevated temperature, using conventional techniques, to compact and fully cured the stanchion 62.

Although this invention has 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.

Claims (6)

1. CLAIMS1. 1. An energy absorbing structural member, comprising: an elongated body (40a, 64) having first and second opposite ends, and first and second opposite sides, the body containing laminated plies (46) of reinforced composite material, the plies having varying lengths forming steps along the length of the member such that the compression strength of the member is progressively less from the first end to the second end; and, at least a first reinforcing flange (40b, 68) extending from the main body along the first side, the first flange tapering in width from the first end of the body to the second end of the body.
2. 2. The energy absorbing structural member of claim 1, wherein the body contains laminated plies (46) of reinforced composite material, the plies in the body having varying lengths forming steps along the length of the body such that the compression strength of the body is progressively less from the first end of the body to the second end of the body.
3. 3. The energy absorbing structural member of claim 2, further comprising a second reinforcing flange (40b, 68) extending from the body along the second side, the second flange tapering in width from the first end of the body to the second end of the body.
4. 4. The energy absorbing structural member of claim 1, further comprising: a second reinforcing flange (68) extending from the main body along the second side, the second flange tapering in width from the first end of the body to the second end of the body; and third and fourth flanges (66) for mounting the structural member, the third and fourth flanges being respectively connected to the first and second flanges.
5. 5. The energy absorbing structural member of claim 4, wherein the third and fourth flanges (66) extend toward each other and perpendicular to the first and second flanges.
6. 6. An energy absorbing structural member, substantially as described herein with reference to the accompanying drawings.