RU2388659C1 - Device designed to secure gas turbine engine to aircraft - Google Patents

Abstract

FIELD: transport.

SUBSTANCE: invention relates to aircraft engineering, particularly to device designed to joint gas turbine engine to aircraft. Proposed device comprise load bearing part to transmit loads, crosspieces and brackets fixed on said parts, including standby parts. All said elements are interconnected with the help of fasteners, including rod-type and hinge-type fasteners, with a spacing relative to load bearing parts transmitting loads. Stand by parts are distributed in engine vertical lengthwise cross section in front and rear attachment rings.

EFFECT: higher reliability of attachment and reduced weight.

3 cl, 9 dwg

Description

The claimed technical solution relates to the fastening of a gas turbine engine to an aircraft using parts that duplicate the main parts in the device in case of non-calculated damage to the latter.

The analysis showed that when one of the parts of the engine mount to the aircraft in the front and rear planes of the suspension, or the rod for removing the thrust, is destroyed, the engine displaces relative to the aircraft pylon.

At the same time, the loads acting on the other engine mountings on aircraft can exceed the calculated ones, which aggravates the relevance of the search for necessary solutions in world aviation.

So it is known the device of the front engine mount with duplication of elements (see the description for the patent belonging to SNECMA company US No. 5871176, class MKI B64D 27/26, class US 244-54, published 02.16.1999), where in the front node An additional support unit has been introduced to the suspension, which, in the event of the destruction of the main support unit, can take over in 3 measurements of the load from a working gas turbine engine.

The duplicating part is installed in the power part of the engine casing with a gap and comes into operation only in case of destruction of the main suspension unit.

The disadvantage of this design is the fastening of the engine in the front belt at one point, both the main node and the backup, which requires the formation of two parallel nodes on the front mounting node and the motor housing - the main and the backup. In this case, the total mass of the nodes of the device for fastening the engine to the aircraft increases significantly.

A close technical solution for mounting a gas turbine engine to an aircraft, both to its wing and to the fuselage, with two parallel backup links, is known from the description of US patent No. 6,682,015, cl. IPC B64D 27/00, cl. US 244-54, owned by Airbus France and published June 27, 2004).

An intermediate assembly such as an upper traverse or strut for attaching the engine to the pylon of the aircraft and the engine power housing contain three oncoming and parallel eyelets on both sides of the vertical plane through the longitudinal axis of the engine.

Moreover, on one side in the attachment point between the eyes, flat links are sequentially fixed by means of two cylindrical rods parallel to each other and relative to the longitudinal axis.

On the other side of the engine axis between the two eyes, the flat links in the shape of a boomerang are fastened by three cylindrical rods parallel to each other and to the longitudinal axis. In the nodes of the installation of the said rods between the last and the indicated flat links spherical bearings are placed.

An inclined rod for transmitting engine thrust to the aircraft pylon is also provided.

According to the problem being solved, this technical solution is close to the claimed one, but tends to increase the weight of the structure due to additional nodes and parts in the form of intermediate flat links and their upper and lower eyes holding them.

From the description of US patent No. 6059227 of the French company Aerospatiale SNI, cl. IPC B64D 27/00, cl. US 244-54, published 09.05.2000, it is also known the device of the rear mounting unit of a gas turbine engine with backup links installed with a gap. The known technical solution, in particular, contains the lower platform of the pylon, to which is attached four bolts of the yoke.

An eye with a hole, the axis of which is parallel to the longitudinal axis of the engine, is made on the lower platform of the rack, also directed further down. The traverse in its lower part has a transverse groove into which the aforementioned rack eye is inserted.

A reserve rod was inserted into the eye with a gap for mounting the engine. Bottom in the traverse, on both sides of the vertical plane, there are eyes on which flat links with cylindrical rods are fixed, the axes of which are parallel to the longitudinal axis of the engine.

These rods are inserted into the eyes on the engine power housing.

In this case, on one side from the bottom of the traverse, a flat link is made in the form of a boomerang, in the eye of which there is one rod parallel to the longitudinal axis of the fixed engine.

On the other hand, an earring is attached to the traverse on a cylindrical hinge, movably holding the other fastening rod parallel to the longitudinal axis of the engine fixed to it. The solution according to the well-known US patent No. 6059227 may be the prototype of the claimed by the generality of the task and the proximity of the totality of essential features. In addition, it also allows the engine to be mounted to the pylon of the aircraft in different positions - top, bottom and side.

The authors were faced with the task of ensuring the reliable performance of the gas turbine engine mounting device by redundant parts that could collapse, regardless of the position of the engine relative to the wing or fuselage of the aircraft when the overall technical result was achieved, namely:

- in the attachment device of a gas turbine, in particular dual-circuit engine, to the aircraft in the event of non-design damage or destruction of one of the engine mounts and the displacement of the latter relative to the pylon, to prevent exceeding the calculated loads acting in the mountings remaining intact, and to ensure their operability, preventing their possible further destruction, while

- to prevent significant movement of the engine relative to its mounting on the aircraft:

- reduce the weight of the device while increasing its reliability.

The problem is solved in that in the known device for mounting a gas turbine engine to an aircraft with load transfer nodes including support nodes, traverses and brackets, including backup, interconnected by threaded and articulated joints, as well as rod parts, including backup, introduced with a gap in the holes of the support nodes, an improvement is made.

The improvement consists in the fact that the spare parts of the mount are distributed between the front and rear mounts along the power housing of the engine.

In the front mounting belt, its crosshead is placed in front of the backup bracket mounted relative to this crosshead with a gap along the axis of the engine.

In this traverse, and further in the backup bracket of the front attachment belt, holes are sequentially made with a common axis parallel to the axis of the engine.

An additional backup bolt is introduced into these holes by its rod part.

At the entrance to the traverse, the backup bolt is locked in its longitudinal position.

At the same time, in the hole of the backup bracket, the rod part of the bolt is installed with an annular gap. A thrust washer is fixed at the outlet of the hole of the backup bracket on the threaded part of the specified bolt. Between the specified washer and the backup bracket also provides a gap along the longitudinal axis of the engine.

The traverse of the front mount may be made of two adjacent in the transverse plane of the parts interconnected and sequentially placed along the longitudinal vertical section of the device.

The rear mounting belt can be divided into upper and lower interconnected mounting areas.

In the lower region of the rear belt, the lower crosshead in this solution is fixed below between two power ribs of the engine housing with two earrings on both sides of a longitudinal vertical section, and from above - by a bracket and strut with the upper mounting area proper to the aircraft, in particular to its pylon.

The backup bolt is located in the indicated longitudinal section of the engine.

With its core part, it is sequentially, in a common axis, tightly placed in the hole of one rib, then, with an annular gap, in the hole in the traverse and then in the hole of the second power rib, outside of which it is fixed, for example, by a threaded connection.

In the event of the destruction of one of the yokes of the traverse of the lower region of the rear attachment belt connecting the rear suspension assembly with the power ribs of the engine, the latter is still held in two places due to the transfer of load to the backup bolt.

The upper part of the rear attachment belt directly to the aircraft, in particular to its pylon, is provided with a bracket with a connecting pad.

The opposing ribs of the lower side of this bracket, opposite the connection pad, encompass the next lower bracket.

Both brackets at the point of coverage are fastened together by two longitudinal bolts with hinged liners. Between these bolts, a backup bolt is additionally introduced, which is placed with an annular gap in its hole in the indicated lower bracket with its rod part.

The claimed technical solution is illustrated by figures, where

figure 1 is a General view of the inventive device for mounting a gas turbine engine to an aircraft and the mutual placement of the front and rear zones of the device;

figure 2 is a section aa of figure 1 of the front node of the inventive fastening device;

figure 3 shows a section bB Figure 1 of the rear node of the inventive device between the power parts of the body of the gas turbine engine and the pylon of the aircraft;

figure 4 presents an enlarged view In figure 1 of the relative position of the bracket, the beam, its components and installed in them backup bolt of the inventive mounting device relative to the power parts of the engine housing;

figure 5 is a section DD DD figure 4 mounting the backup bracket to the engine;

in Fig.6 shows a view of G Fig.4 - fixing the backup bolt backup bracket;

in Fig.7 presents a view G of Fig.3 of the rear attachment belt and its parts, interconnected;

in Fig.8 shows a section LL of Fig.7 at the location of the backup bolt in the power ribs of the engine housing;

Fig.9 shows a section KK in the connection of the two brackets of the rear belt assembly and the placement of the backup bolt in this connection.

The claimed technical solution can be equivalently implemented when attaching the engine to the aircraft from above, side and bottom.

According to the presented figures, we consider the inventive device for mounting the engine to the aircraft, in particular, in its embodiment under the pylon of the aircraft, as shown in Figure 1.

The engine 1 on the power pylon 2 of the aircraft is fixed in the longitudinal direction sequentially in two zones.

In the front belt to the pylon 2 by known means and a traverse 3 is rigidly connected (see Figure 2). In this case, through the said crosshead 3, the engine1 is mounted on the pylon 2 with its upper part.

In this case, to the engine 1, the front crosspiece 3 is pivotally fixed with its lower part in two design places on both sides of the vertical longitudinal plane of the cross section of the engine 1. Fastening was carried out on one side by an arm 4 with a hinge, and on the other hand by an earring 5 with two hinges.

At the same time, the said front crosshead 3 can be made of two transversely adjacent and arranged one after another in the longitudinal direction of the individual parts 3a and 3b (see Fig. 4, view B), interconnected, as will be described later.

In the rear belt, engine 1 is attached to the pylon 2 of the aircraft using interconnected brackets 6 and 7 (Fig. 7, view D), strut 8 and crosshead 9, successively placed downward vertically to engine 1.

The engine 1 is attached to the lower traverse 9 of the rear belt (see also Fig. 7) using two earrings 10 on both sides of the longitudinal vertical plane of the cross section of the engine 1.

Moreover, the above-described series of sequentially interconnected parts connects the platform 11 of the aircraft pylon 2 with the power ribs 12 on the outer power housing of the engine 1 (Fig. 8, section L-L).

The above-mentioned bracket 6 by its connecting pad in a known manner, for example, using threaded connections, is attached through the pad 11 to the pylon 2 of the aircraft.

The bracket 6 in the vertical direction to the engine 1 is divided into at least two power parts between which a subsequent bracket 7 is inserted.

In this case, the brackets 6 and 7 are fastened, for example, by two bolts 13.

Around the bolts 13 in the places of their installation in the inner bracket 7, to relieve additional stresses, spherical joints 14 are introduced (Fig. 9).

To transfer the thrust of the engine 1 to the pylon 2 of the aircraft, an inclined rod 15 is also provided, which is located in the case under consideration under the pylon 2 of the aircraft (Figure 1).

With its front end, the rod 15 is fixed on the bracket 16, located mainly on the dividing body 17 at the front mounting belt. The other, its rear end, the rod is attached to the bracket 6 of the aforementioned rear mounting belt (see view D, Fig.7).

In the front mounting belt on the power, in particular dividing, housing 17 of the engine 1 (type B, FIG. 4), a backup bracket 18 with a longitudinal clearance C ₁ relative to the end face of the said beam 3 (its part 3b) is additionally rigidly fixed.

In the front traverse 3 (3a, 3b), an opening was made for the design unit without a gap of the rod cylindrical part of the bolt 19. Parts 3a and 3b, united by a bolt 19 and fixed with a lock plate 20 and mounting bolts 21, are adjacent to each other,

The bolt 19 is held by the locking plate 20 and the bolts 21 with an emphasis on the open end of part 3a of the beam 3. At the end of the bolt 19 - its threaded part behind the backup bracket 18, a thrust washer 22 is installed, and the whole set is fixed with a nut 23.

Between the backup bracket 18 and the thrust washer 22, a design clearance C ₂ is provided in the longitudinal direction.

Moreover, outside the crosshead 3, namely its part 3b, the rod part of the bolt 19 penetrates the bracket 18 with an annular gap C ₃ (see also Figure 5, section DD).

Due to the gaps (longitudinal - C ₁ , C ₂ and ring - C ₃ ), the backup bracket 18 (see Figs. 4 and 5) only accepts load if one of the engine mounts in the front belt is destroyed, or the rod 15 for pulling the pull force .

When the bracket 4 or earring 5 (see FIG. 2) is destroyed, providing a joint working relationship between the engine 1 through the yoke 3 (3a, 3b) to the pylon 2, a certain displacement of the engine 1 relative to the specified pylon 2 occurs. In this case, an annular gap C ₃ is selected between the bolt 19 and the backup bracket 18 (Fig.6).

As a result, the engine 1 in the front attachment belt is held by the backup bracket 18 and the remaining undamaged either the bracket 4 or the earring 5.

When the rod 15 for removing the thrust or individual parts of the front attachment is destroyed, a longitudinal clearance C1 is selected between the backup bracket 18 and part of the yoke 3 (its part 3b), through which the thrust is transmitted to the pylon 2 of the aircraft.

When one of the parts of the traverse 3a is destroyed, all of the above clearances C ₁ , C ₂ and C ₃ provide their safety role, and at the same time, the traverse 3, as a common node, remains operational.

Consider further, in the claimed technical solution, the rear belt of the gas turbine engine 1 to the aircraft, which provides engine protection in case of non-calculated damage or destruction of one of the parts of this engine belt and also the rod 15.

So, in the section LL of the rear attachment belt (see Figs. 7, 8), a backup fastening with a bolt 24 of the beam 9 to the engine 1 is provided through the power ribs 12. An annular gap M is provided between the beam 9 and the shaft part of the bolt 24 (see Fig. 8). The backup bolt 24, shifting in the cavity of the gap M, comes into operation and assumes the load in the event of the destruction of one of the earrings 10 or parts of their fastening, connecting the crosshead 9 of the rear mounting belt with the power ribs 12 of the engine housing 1. At the same time, the engine 1 continues to be held in the back fixing belt described above for two points - a backup bolt 24 and one of the earrings 10.

In the same back zone, in the immediate vicinity of the pylon 2, in the connection of the ribs of the bracket 6 and the bracket 7 located between them, with an annular gap H relative to the latter, a backup bolt 25 is inserted (see section KK, Figs. 7 and 9).

The bolt 25 is placed in the fastening of the brackets 6 - covering and 7 - covered, between and parallel to the described two bolts 13 with hinged inserts 14. In the event of the destruction of at least one of the bolts 13, or its eyes, in the bracket 6 or 7 takes the load a backup bolt 25, displaced in the cavity of the gap H and thereby also ensuring the safety of the aircraft in critical situations.

A certain compensation for such a displacement will be taken over by the hinged insert of the remaining one of the bolts 13.

This technical solution of fastening the engine to the aircraft allows you to duplicate the main components of the device with optimal load distribution along with minimizing their weight.

In the event of non-designated destruction of one of the attachment points, the reservation details that come into operation prevent significant movement of the engine relative to the pylon of the aircraft. In this case, the loads acting in the remaining attachment points will not exceed the calculated ones.

The solution has been developed as applied to D-436-148 engines installed on AN-148 aircraft, but can be widely used for the D-436T1 family of engines and their modifications, in particular, for example, mounted with side suspension on Tu-334 aircraft.

Claims (3)

1. A device for attaching a gas turbine engine to an aircraft, comprising power parts for transferring loads and traverses and brackets fixed to power parts, including backup, interconnected by means of fasteners, including rod, and articulated joints, including backup bolts interconnected with a gap relative to the power parts of the load transmission, characterized in that the spare parts of the mount are distributed in a longitudinal vertical section of the engine in the front and rear mounting belts. 2. The attachment device of a gas turbine engine to an aircraft according to claim 1, characterized in that in the front attachment belt the backup bolt with its rod part in longitudinal and circumferential directions is rigidly mounted in its traverse, and further, relative to the additional backup bracket with gaps, annular and transverse, and in the rear attachment belt, its lower crosshead is connected to the engine housing by earrings on both sides of the backup connection and, subsequently upwards with the aircraft, through the lower and upper crowns matte by means of fasteners on both sides of the backup bolt with an annular gap around the latter. 3. The device for attaching a gas turbine engine to an aircraft according to claim 2, characterized in that the crosshead of the front mount is made of two adjacent in the transverse plane parts interconnected and sequentially placed along the longitudinal vertical section of the device, and the backup bolt is rigidly fixed in coaxial the holes of both indicated parts of the traverse, and in the rear attachment belt on the opposing power parts of the backup bolt input are made connections in the form of interconnected paired power ribs one power unit, between which a power rib of the oncoming power unit is introduced in each of the connections.

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