Device for attaching a heat shield to an aircraft panel and associated method

Abstract

The invention relates to a device (1) for attaching a heat shield (2) to an aircraft panel (3), the device comprising a first attachment element (4) associated with the heat shield (2), a second attachment element (5) associated with the aircraft panel, and a metal wire (6) forming a loop (7). At least one of the attachment elements is provided with means (10, 11) for retaining the loop, the loop being closed by a braking member (12) of the ends (9) of the loop intended for bringing the attachment elements together, the loop abutting one of the attachment elements at its middle portion (8) and the other one of the attachment elements at the braking member (12), respectively. The loop is set apart from the heat shield (2) and the aircraft panel (3), and passes outside the heat shield (2) and the aircraft panel (3).

Description

[0001] DESCRIPTION

[0002] TITLE: Device for attaching thermal protection to an aircraft panel and associated method

[0003] Technical field

[0004] The present invention relates to a device for fixing thermal protection to an aircraft panel.

[0005] Previous techniques

[0006] An aircraft is powered by one or more turbojet engines, each housed in a nacelle. A nacelle generally has a substantially tubular structure comprising an air inlet upstream of the turbojet engine, an intermediate assembly intended to surround a fan of the turbojet engine and a rear assembly which may incorporate thrust reversal means and intended to surround the combustion chamber and all or part of the compressor and turbine stages of the turbojet engine. The nacelle also includes an exhaust nozzle whose outlet is located downstream of the turbojet engine.

[0007] In a turbojet engine, certain areas are likely to catch fire in the event of a fault. These are called "fire compartments." Other areas, called "non-fire compartments," are, on the contrary, free from any direct risk of ignition. However, they must be protected from a fire originating from a fire compartment. For example, the interior of a nacelle is a fire compartment, and the exterior of the nacelle is a non-fire compartment.

[0008] It is therefore necessary to fix thermal protections inside the nacelle to prevent or delay the passage of a fire from the inside to the outside of the nacelle in the event of damage.

[0009] An example of a known fastening system is described in document FR 2 829 81 1. Such a fastening system uses a lock wire system and a retaining button. The retaining button bears against an external surface of the thermal protection so as to form a retaining and tightening washer, the retaining button being provided with holes for the passage of the lock wire, the ends of which are twisted and tightened. However, the use of such a fastening system becomes difficult, or even impossible, in spaces of reduced dimensions, typically when the width of the thermal protection to be fixed is less than 30 mm.

[0010] Generally, known fixing systems do not allow easy installation in deep and / or narrow spaces.

[0011] Statement of the invention

[0012] The invention therefore aims to facilitate the mounting of thermal protections on aircraft panels.

[0013] The invention relates to a device for fixing thermal protection to an aircraft panel, comprising a first fixing element associated with the thermal protection, a second fixing element associated with the aircraft panel and a wire forming a loop.

[0014] At least one of said first and second fixing elements is provided with means for retaining said loop, said loop being closed by a braking member of the ends of said loop intended to bring said fixing elements together, said loop bearing respectively in its middle part on one of said fixing elements and at the level of the braking member on the other of said fixing elements.

[0015] Preferably, the first fixing element is fixed to the thermal protection.

[0016] Preferably, the second fastening element is secured to the aircraft panel.

[0017] Advantageously, the loop is spaced from the thermal protection and the aircraft panel, the loop passing outside the thermal protection and the aircraft panel. Thus, the metal wire does not pass through the thermal protection or the aircraft panel. Advantageously, the metal wire is spaced from a contact area between said aircraft panel and said thermal protection.

[0018] According to an advantageous characteristic, said retention means aim to prevent out-of-plane movement of said loop. Preferably, out-of-plane movement of the loop is prevented by the retention means when the loop is braked in the tightened position.

[0019] For example, said device comprises a third fixing element comprising a central part pierced with two holes and connecting first and second parallel lateral parts, substantially perpendicular to the central part, said first and second lateral parts being arranged respectively between the thermal protection and the first fixing element, and between the aircraft panel and the second fixing element, said two holes being crossed by said loop and said central part being arranged between the twist of said loop and the fixing element on which said twist rests.

[0020] For example, said third fixing element further comprises a third and fourth parallel lateral part, substantially perpendicular to the central part, extending in the extension of the first and second lateral parts on an opposite side of said central part.

[0021] According to another aspect, the invention relates to a method of fixing thermal protection to an aircraft panel using a fixing device as described above.

[0022] The process includes successive steps of:

[0023] - positioning said thermal protection and said panel; engaging a loop of metal wire in said means for retaining said fixing elements; and, closing said loop by twisting the ends of said loop to bring together and tighten said fixing elements. Preferably, the fixing method further comprises an additional step of folding said twist carried out after the closing step.

[0024] According to an advantageous characteristic, the method comprises an additional step of installing a third fixing element comprising a central part pierced with two holes, connecting first and second parallel lateral parts, substantially perpendicular to the central part, said central part being threaded onto the loop through the holes, said first and second lateral parts being installed respectively between the thermal protection and the first fixing element, and between the aircraft panel and the second fixing element, said installation step being carried out after the engagement step.

[0025] According to another aspect, the invention relates to a turbomachine nacelle provided with a fixing device as defined above.

[0026] Brief description of the drawings

[0027] Other aims, characteristics and advantages of the invention will appear on reading the following description, given solely by way of non-limiting example, and made with reference to the appended drawings in which:

[0028] [Fig 1] and [Fig 2] are sectional views of a fixing device according to the invention;

[0029] [Fig 3] is a perspective view of the fixing device of Figure 1, without the wire;

[0030] [Fig 4] is a perspective view of the fixing device of Fig. 2; and

[0031] [Fig 5] and [Fig 6] illustrate a second embodiment according to the invention;

[0032] [Fig 7] is a sectional view of a variant of the embodiment of Figures 5 and 6;

[0033] [Fig 8] is a flowchart of a fixing method according to the invention; and [Fig 9] is a schematic sectional view of a turbomachine nacelle provided with a fixing device according to the invention.

[0034] Detailed description of at least one embodiment

[0035] Figures 1 and 2 are sectional views of a device 1 for attaching a thermal protection 2 to an aircraft panel 3 according to one embodiment of the invention. The aircraft panel 3 comprises an axial portion 3a and a radial portion 3b. The attachment device 1 is illustrated in the untightened position in Figure 1 and in the tightened position in Figure 2.

[0036] The fixing device 1 comprises a first fixing element 4 fixed to the thermal protection 2, a second fixing element 5 fixed to the aircraft panel 3 and a wire 6 forming a loop

[0037] 7 which passes through retention means 10, 11 provided on at least one of the fixing elements 4, 5. The loop 7 has a middle part

[0038] 8 and two ends 9. The retaining means 10, 11 of the loop 7 are configured to prevent out-of-plane movement of the loop 7. In the illustrated embodiment, the fixing elements 4, 5 are both provided with retaining means 10, 11 of the loop 7. Alternatively, only one of the fixing elements 4, 5 is provided with such retaining means.

[0039] Wire 6 may be metallic in particular.

[0040] The fixing elements 4, 5 may in particular be plates, angles or hooks.

[0041] The retention means 10, 11 may in particular be hooks, grooves, notches, holes or grooves. The retention means 11 provided on the fixing element 4 may be different from the retention means 10 provided on the fixing element 5.

[0042] The loop 7 is closed by a braking member 12. The braking member 12 is in particular a twist of the ends 9 of the wire 6.

[0043] When the loop 7 is closed by a twist 12 of the ends 9, the fixing elements 4, 5 move closer together and the loop 7 bears respectively in its middle part 8 on the second fixing element 5 and at the level of the twist 12 on the first element

[0044] 4 fixing (figure 2). The twist 12 is made so that the thermal protection 2 is in contact with the axial part 3a of the aircraft panel 3 on a contact zone 13.

[0045] Figure 3 is a perspective view of the first 4 and second

[0046] 5 fixing elements without the 6 metal wire.

[0047] Figure 4 is a perspective view of the fixing device 1 in the clamped position.

[0048] It should be noted that thanks to the fixing device 1, the metal wire 6 is moved away from the contact zone 13 (figure 2). In other words, the fixing elements 4 and 5 extend from the panel 3 and the thermal protection 2 in the opposite direction to the contact zone 13, creating an interaction zone between the wire 6 and the elements 4 and 5 which is at a distance from the contact zone 13.

[0049] The installation of such a fixing is thus facilitated, as it benefits from better accessibility.

[0050] Furthermore, thanks to the fixing device 1, the loop 7 is at a distance from the thermal protection and the aircraft panel, without the metal wire 6 passing through either said thermal protection or said aircraft panel. Thus, the ease of assembly and the effectiveness of the thermal protections are improved because it is not necessary to provide any reservations or holes in the parts to be fixed, while avoiding thermal bridges.

[0051] Figures 5 and 6 illustrate a second embodiment according to the invention.

[0052] In this embodiment, the fixing device 1 comprises a first fixing element 4 fixed to the thermal protection 2, a second fixing element 5 fixed to the aircraft panel 3 and a wire

[0053] 6 metal forming a loop 7 closed by a twist 12.

[0054] The fixing device 1 further comprises a third fixing element 14 comprising a central part 15 pierced with two holes, connecting first 16 and second 17 lateral parts parallel and substantially perpendicular to the central part 15. In the example illustrated, the first 16 and second 17 lateral parts have the same length. Alternatively, it is possible for the first 16 and second 17 lateral parts to have different lengths.

[0055] The first 16 and second 17 lateral parts are arranged respectively between the thermal protection 2 and the first fixing element 4, and between the aircraft panel 3 and the second fixing element 5.

[0056] The two holes are crossed by the loop 7 and the central part 15 is arranged between the twist 12 of the loop 7 and the fixing element 4 on which the twist 12 rests.

[0057] Such a fastening device 1 provided with the third fastening element 14 makes it possible to reduce the risk of breakage of the metal wire 6 when closing the loop by twisting the ends. Indeed, in the absence of the third fastening element 14, the loop bears at the level of the twist on a first 4 or second 5 fastening element. In this zone, the metal wire 6 undergoes stresses due to the crushing generated by the support of the twist 12 in addition to the significant tangential torsional stresses. Given that the crushing stresses are inversely proportional to the size of the support, there is a significant risk of breakage of the wire 6 when the twist bears on a zone of relatively small size such as the thickness of a fastening element 4, 5. The use of the third fastening element 15 makes it possible to enlarge the size of the support so as to reduce the risk of breakage of the wire 6.

[0058] Figure 7 is a sectional view of a variant of the second embodiment according to the invention.

[0059] In this variant, the third fixing element 14 further comprises third 18 and fourth 19 parallel lateral parts, substantially perpendicular to the central part 15 and extending in the extension of the first 16 and second 17 lateral parts on an opposite side of the central part 15.

[0060] The third and fourth side portions protect the thermal protection and the aircraft panel from possible damage caused by the twist 12. Indeed, when bending the twist 12, it is possible that the twist 12 is inadvertently inclined towards the thermal protection or towards the aircraft panel with a risk of damaging the latter. The use of a third fixing element 15 with third 18 and fourth 19 side portions introduces a physical protective barrier between the twist 12 and the parts to be protected so as to reduce the risk of damage.

[0061] Figure 8 is a flowchart of a method 20 of attaching thermal protection to an aircraft panel according to the invention.

[0062] The method begins with a step 21 of positioning the thermal protection 2 and the aircraft panel 3.

[0063] The method continues with a step 22 of engaging a loop 7 of metal wire 6 in retention means 10, 11 provided on at least one of the first and second fixing elements 4, 5.

[0064] In the next step 24, the loop is closed and the ends of the loop are twisted to bring the fastening elements together and tighten them.

[0065] In the next step 25, the twist is folded to reduce its size.

[0066] The method comprises an optional step 23 of installing a third fastening element 14 as described previously, the step 23 being carried out after the engagement step 22 and before the closing step 24. During step 23, the central part 15 of the third fastening element 14 is threaded onto the loop 7 through the holes until it comes into contact with a fastening element, the first 16 and second 17 lateral parts being installed respectively between the thermal protection 2 and the first fastening element 4, and between the aircraft panel 3 and the second fastening element 5. It is understood that the twist 12 of the loop 7 is only carried out in the case of installation of the third fastening element 14 after the loop 7 of metal wire 6 has passed through the holes of the central part 15.

[0067] The fixing device 1 described above is advantageously intended to equip a nacelle 30 of a turbomachine or aircraft engine visible in FIG. 9.

[0068] Figure 9 shows very schematically an axial section of a turbomachine 31, with a general longitudinal axis X, for example of the double-flow, double-body turbojet type.

[0069] Access to the engine 31 is provided by a movable hood 32.

[0070] Numerous thermal protections 2 are attached using devices 1 on panels 3 of the nacelle 30 to prevent or delay the passage of a flame from the engine 31 to the outside of the nacelle 30, and to thermally insulate certain parts which may re-ignite gases outside the nacelle 30.

[0071] It should be noted that the fixing device 1 of the invention can equip any type of nacelle and fix any type of thermal protection.

Claims

CLAIMS 1. Device (1) for fixing a thermal protection (2) to an aircraft panel (3), comprising a first fixing element (4) associated with the thermal protection (2), a second fixing element (5) associated with the aircraft panel and a wire (6) forming a loop (7), characterized in that at least one of said first and second fixing elements is provided with means (10, 11) for retaining said loop, said loop being closed by a braking member (12) of the ends (9) of said loop intended to bring said fixing elements together, said loop bearing respectively in its middle part (8) on one of said fixing elements and at the level of the braking member (12) on the other of said fixing elements, said loop being at a distance from said thermal protection (2) and said aircraft panel (3), said loop passing outside said thermal protection (2) and the aircraft panel (3) aircraft.

2. Device according to claim 1, wherein said wire (6) is remote from a contact zone (13) between said aircraft panel and said thermal protection.

3. Device according to one of claims 1 and 2, wherein said retaining means (10, 11) are configured to prevent out-of-plane movement of said loop.

4. Device according to any one of claims 1 to 3, in which said device comprises a third fixing element comprising a central part (15) pierced with two holes and connecting first and second parallel lateral parts, substantially perpendicular to the central part, said first and second lateral parts being arranged respectively between the thermal protection and the first fixing element, and between the aircraft panel and the second fixing element, said two holes being crossed by said loop and said central part being arranged between the twist of said loop and the fixing element on which said twist rests.

5. Device according to claim 4, wherein said third fixing element further comprises a third (18) and fourth (19) parallel lateral parts, substantially perpendicular to the central part (15), extending in the extension of the first (16) and second (17) lateral parts on an opposite side of said central part (15).

6. Method for fixing thermal protection to an aircraft panel using a fixing device according to any one of claims 1 to 5, characterized in that it comprises successive steps of: - positioning said thermal protection and said panel; engaging a loop of metal wire in said means for retaining said fixing elements; and, closing said loop by twisting the ends of said loop in order to bring said fixing elements together.

7. A method of fixing according to claim 6, further comprising an additional step of folding said twist made after the closing step.

8. A fixing method according to one of claims 6 and 7, comprising an additional step of installing a third fixing element (14) comprising a central part (15) pierced with two holes, connecting first (16) and second (17) parallel lateral parts, substantially perpendicular to the central part, said central part (15) being threaded onto the loop through the holes, said first and second lateral parts being installed respectively between the thermal protection (2) and the first fixing element (4), and between the aircraft panel (3) and the second fixing element (5), said installation step being carried out after the engagement step.

9. Turbomachine nacelle provided with at least one fixing device according to any one of claims 1 to 5.

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