Blowing conduit for a turbine engine

CA 02602172 2007-09-18 1 Fan duct for a turbomachine The present invention relates to a fan duct for a turbomachine such as an airplane turbojet.

The fan duct (often called the fan duct) of a turbomachine is part of the nacelle and extends around the engine of the turbomachine, between the fan and the exhaust nozzle.

It comprises two substantially cylindrical walls which extend one inside the other and define between them an annular space for the flow of part of the air sucked in by the fan, called cold flow or secondary flow. . The other part of the air sucked in by the fan enters the engine of the turbomachine which comprises from upstream to downstream a compressor, a combustion chamber and a turbine, and forms the hot flow or primary flow.

The internal wall of the pipe is fixed at its upstream end to an intermediate casing of the turbomachine, and is connected at its downstream end to an exhaust casing.

It is shaped to wrap a short radial distance around the engine of the turbomachine and may be barrel shaped when a centrifugal compressor is used in the engine.

The outer wall of the pipe is fixed at its upstream end on the intermediate casing and at its downstream end at the upstream end of a thrust reverser which is mounted around the exhaust casing.

The downstream end of the outer wall is also connected to the exhaust casing by support and centering means.

In the prior art, the internal wall of the pipe comprises openings for access to equipment, such as in particular fuel injectors and control cylinders for variable-pitch vanes, which are mounted on the body of the turbomachine. inside the inner wall.

The outer wall for its part comprises orifices for the passage of easements which are housed in arms which extend CA 02602172 2007-09-18 2 substantially radially between the inner and outer walls of the pipe.

The openings in the inner wall of the pipe are closed by removably attached panels.

The outer wall may include hinged cowls which open for inspection or maintenance operations, or hatches closed by bolted panels, in particular when the engine is fixed to the fuselage of the airplane.

The operations for mounting the fan duct on the engine are generally long and complex, as are the maintenance operations which may require more or less complete dismantling of the fan duct or the engine.

The object of the invention is in particular to provide a simple, effective and economical solution to these problems.

For this purpose, it proposes a fan duct for a turbomachine, comprising two coaxial cylindrical walls, internal and external respectively, intended to be fixed at one end to an intermediate casing of the turbomachine and to be connected at their other end to a casing of 'turbomachine exhaust, characterized in that, to facilitate the assembly of the pipe and allow access to equipment mounted on the body of the turbomachine, the internal cylindrical wall consists of removable panels fixed to a framework comprising an upstream annular structure, a downstream annular structure and longitudinal connecting elements of these structures, and in that the external cylindrical wall comprises openings closed by panels fixed from removable way on a unitary support structure, these openings having dimensions allowing the passage of the panels of the internal wall and the assembly and disassembly of these panels on the framework of the internal wall.

According to the invention, the parts that make up the framework of the internal wall can be placed and assembled to each other and on the body of the turbomachine so as to best match the shape CA 02602172 2007-09-18 3 of this body.

The support structure of the outer wall is advantageously rigid and unitary, which allows the transmission of forces between the turbomachine and the aircraft equipped with this turbomachine while facilitating the mounting of this wall.

The removable panels of the inner wall of the duct can pass through the openings in the outer wall, so that they can be removed without dismantling the support structure from the outer wall of the fan duct.

The maintenance operations of the equipment located inside the internal wall are simpler and faster because they are carried out directly through the openings of the external wall.

They therefore result in a relatively short immobilization time of the aircraft equipped with the turbomachine.

In a preferred embodiment of the invention, the support structure of the outer wall comprises upstream and downstream rings rigidly connected by longitudinal uprights defining between them the aforementioned openings and which are for example four in number and distributed at 900 from each other around the longitudinal axis of the outer wall.

At least some of these uprights may include orifices for mounting easement passage arms such as fluid circulation conduits and electrical cables.

Advantageously, the outer wall comprises another downstream ring fixed to the support structure and comprising connecting rods with the exhaust casing of the turbomachine.

The rods provide support and centering of the engine, and allow axial and radial expansion of the exhaust casing in operation.

The internal wall of the fan duct is not a force transmission part and its frame and its removable panels are for example made of sheet metal.

The removable panels of the internal wall are substantially aligned with the openings of the external wall in order to facilitate maintenance operations on the equipment mounted on the body of the turbomachine.

The longitudinal connecting elements of CA 02602172 2007-09-18 4 the framework of the internal wall comprise mounting holes for the auxiliary passage arms, these holes being radially aligned with the corresponding holes in the uprights of the external wall.

The invention also relates to a turbomachine, such as an airplane turbojet, characterized in that it comprises a fan duct as described above.

Preferably, the upstream ends of the inner and outer walls of the pipe are fixed to annular flanges of the intermediate casing of the turbomachine, and the downstream end of the inner wall is fixed to an upstream flange of the exhaust casing by radial lugs leaving an annular ventilation space around the exhaust housing. The downstream end of the outer wall is connected to the exhaust casing by connecting rods allowing axial and radial expansion of this casing in operation. The downstream end of the outer wall can also be attached to the upstream end of a thrust reverser extending around the exhaust casing.

The invention also proposes a method for mounting a fan duct as described above in a turbomachine such as an airplane turbojet, this method consisting of:

- to fix the annular structure upstream of the internal wall to a flange of the intermediate casing, then to fix the annular structure downstream of the internal wall to a flange of the exhaust casing, and to connect these structures by longitudinal connecting elements, - then to set up the unitary support structure of the outer wall, by axial translation from downstream to upstream, and to fix the upstream end of this structure on an annular flange of the intermediate casing, - to mount the easements and their passage arms in the corresponding openings of the internal and external walls, - to position and fix on the downstream end of the external wall the ring comprising the connecting rods with the housing of exhaust, CA 02602172 2007-09-18 then to put in place and to fix the cover plates around the rods on the downstream end of the internal wall, - to put in place the panels of the internal wall, by passing through the openings in the outer wall, and securing them, then placing and securing the panels of the outer wall.

The invention will be better understood and other details, characteristics and advantages of the present invention will become apparent on reading the following description given by way of non-limiting example with reference to the appended drawings in which:

- Figure 1 is a partial schematic half view in axial section of a bypass turbojet equipped with a fan duct according to the prior art;

- Figure 2 is a schematic exploded perspective view of an outer cylindrical wall of a fan duct according to the invention;

- Figure 3 is a schematic exploded perspective view of an internal cylindrical wall of a fan duct according to the invention;

- Figure 4 is a schematic perspective view of the framework of the internal wall of Figure 3;

- Figures 5 to 10 are schematic perspective views of a bypass turbojet and illustrate steps for mounting the fan duct of Figures 2 to 4 on the body of the turbojet. double flow comprising from upstream to downstream, in the direction of the gas flow inside the turbojet, a fan 10, a compressor 12, a combustion chamber 14 and turbines 16, this turbojet being intended to be fixed by appropriate means under a wing of an airplane or on the rear part of the fuselage of an airplane.

The fan 10 comprises a plurality of blades 18 which are fixed at their radially inner ends on the periphery of the fan disc 20 of the turbojet which is itself fixed to the upstream end of a shaft (not shown) of the turbojet.

The fan blades 18 CA 02602172 2007-09-18 6 are surrounded on the outside by a retention casing mounted at the upstream end of the nacelle 22 of the turbojet engine which is substantially cylindrical and extends downstream around the compressor 12, of the combustion chamber 14 and of the turbines 16 of the turbojet.

This nacelle 22 makes it possible to channel the air flow 24 entering the turbojet.

Part 26 of this air flow, forming the primary flow or hot flow, enters the compressor 12, the last stage of which is centrifugal, then is mixed with fuel and burned in the combustion chamber 14, to be then injected. in the turbines 16 in order to supply energy to the rotor blades of the turbines and to drive the shaft of the compressor and the fan in rotation.

The other part 28 of the air flow entering the turbojet, forming the secondary flow or cold flow, flows around the body of the turbojet inside an intermediate casing 30 then a fan duct 32, and is used to supply ventilation and cooling circuits and to provide additional thrust in addition to that provided by the combustion gases ejected from the turbines 16.

The intermediate casing 30 comprises two cylindrical shells 34, 35 coaxial, respectively internal and external, which are connected to each other by radial arms or vanes 36.

The fan duct 32 is formed by two substantially cylindrical walls 38, 40 which extend coaxially one inside the other and which are interconnected by tubular radial arms 42 inside which pass tubes. easements such as fluid circulation conduits and electric cables.

The outer wall 40 of the fan duct is fixed at its upstream end to the downstream end of the outer shell 35 of the intermediate casing, and at its downstream end to the upstream end of a thrust reverser (not shown).

Its internal wall 38 is fixed at its upstream end on the downstream end of the internal ferrule 34 of the intermediate casing, and at its downstream end on an exhaust casing 44 mounted at the outlet of the CA 02602172 2007-09-18 7 turbines 16 .

Like the intermediate casing 30, the exhaust casing 44 comprises two coaxial cylindrical collars 46, 48, respectively internal and external, interconnected by radial vanes 50.

The internal wall 38 of the pipe comprises openings for access to equipment mounted on the body of the turbojet, such as, for example, fuel injectors 52 and actuators for controlling variable-pitch blades, these access openings being closed by removable panels (not shown).

In the embodiment of the invention shown in Figures 2 to 4, the outer wall 60 (Figure 2) of the pipe comprises a unitary and rigid support structure 62, comprising openings 64 of large size which can be closed by panels removable 66, and its internal wall 68 (Figures 3 and 4) is formed of sheet metal or composite material parts assembled to each other to form a frame 70 on which are removably fixed panels 72.

The support structure 62 of the outer wall is substantially cylindrical and comprises upstream 74 and downstream 76 rings rigidly connected to each other by uprights 78 of elongated shape extending parallel to the longitudinal axis 80 of the outer wall, these uprights delimiting between them and with the rings 74, 76 the aforementioned openings 64 of the support structure.

In the example shown, these uprights 78 are four in number and are regularly distributed around the axis 80 to define four substantially identical openings 64.

Two diametrically opposed longitudinal uprights 78 comprise orifices 81 for mounting the radially outer ends of tubular radial arms (not shown) for the passage of easements.

The removable panels 66 of the external wall 60 are fixed by screws or bolts on the support structure 62 so that their radially internal faces come to the same level as the internal surface CA 02602172 2007-09-18 = 8 of the support structure to limit the pressure drops during the flow of the secondary flow.

The fixing of the support structure 62 to the intermediate casing 30 and to the thrust reverser, and its connection to the exhaust casing 44, will be described in more detail in the following with reference to FIGS. 5 to 10.

This support structure 62 is sufficiently rigid in operation, on the one hand, to withstand the forces transmitted by the engine and by the thrust reverser, and on the other hand, to ensure the transmission of forces between the turbojet and the engine. airplane equipped with this turbojet.

The section of the internal wall 68 (Figures 3 and 4) increases from its upstream end to its middle part, located at the level of the centrifugal stage of the turbojet compressor, then decreases to its downstream end, the general shape of the internal wall being barrel-shaped.

The framework 70 of the internal wall 68 of the pipe is formed of two annular structures 82, 84 coaxial, respectively upstream and downstream, which are connected to each other by elongated elements 86 which extend in parallel. to the longitudinal axis 88 of the internal wall.

These elements define between themselves and with the annular structures 82, 84 windows 90 for access to the aforementioned equipment mounted on the body of the turbojet.

In the example shown, these longitudinal elements 86 are four in number and are regularly distributed around the axis 88 of the internal wall so as to form four substantially identical windows 90, intended to be substantially aligned radially with the openings 64 of the outer wall.

Two of these longitudinal elements 86 comprise orifices 96 aligned radially with the orifices 81 of the external wall 60 for mounting the radially internal ends of the radial arms for the passage of easements.

The internal wall 68 of the pipe further comprises a downstream tapered cowling 94 made of sheet metal or composite material formed of three CA 02602172 2012-09-14 9 sectors whose upstream ends are fixed to the downstream annular structure 84 of the frame 70 .

This cowling 94 comprises at its downstream end radial notches 95 for the passage of means for suspending the exhaust casing 44 to the support structure 62 of the outer wall.

The removable panels 72 of the internal wall 68 are fixed by screws or bolts or any other rapid fixing system on the framework 70 and some of them include holes 97 for taking air from the secondary flow.

The dimensions of these panels 72 are such that they can be dismantled and removed from the inner wall through the openings 64 in the outer wall.

As will be described in more detail in what follows, the upstream annular structure 82 is intended to be fixed on the inner shell of the intermediate casing 30 and the downstream annular structure comprises inner radial lugs 92 for fixing to the outer shell of the casing. exhaust 44.

These legs are for example three in number and are regularly distributed around the axis 88, at a distance from each other, to allow air intended to ventilate the exhaust casing 44 to pass between them.

FIGS. 5 to 10 illustrate steps of the method for mounting the fan duct according to the invention on a bypass turbojet.

The first step shown in Figure 5 consists in fixing by means of the screw / nut type the upstream annular structure 82 of the internal wall to a corresponding annular flange of the internal ferrule 34 of the intermediate casing 30, in fixing the fixing lugs 92 of the downstream annular structure 84 of the inner wall on corresponding means provided on the outer shell 48 of the exhaust casing 44, then to fix the longitudinal connecting elements 86, at their ends on the upstream structure 82 and on the downstream structure 84. The framework 70 of the internal wall is thus assembled and fixed to the body of the turbojet.

The method then consists (FIG. 6) in bringing the support structure 62 of the external wall around the framework 70 of the internal wall, by moving it axially from the downstream around the framework of the wall CA 02602172 2012- 09-14 _ - 10 internal, then to fix by bolting an annular flange upstream of the support structure on a corresponding annular flange of the outer shell 35 of the intermediate casing 30.

The downstream end of the support structure 62 is fixed by bolting to the upstream end of a thrust reverser, schematically represented by dotted lines 99 in FIG. 7.

The downstream end of the support structure 62 is also connected to the exhaust casing 44 by force transmission rods 98, these rods being fixed at one of their ends on the exhaust housing 44, and connected to the other of their ends to a ring 100 attached and fixed coaxially to the downstream end of the support structure 62.

In operation, these rods 98 provide support and centering of the exhaust casing, and allow its axial and radial expansion in operation.

In the example shown, the rods are six in number and all extend in the same transverse plane being joined in pairs, the radially outer ends of the rods of each pair being at a circumferential distance from each other. which is smaller than the circumferential distance between their radially inner ends.

Tubular arms 102 are engaged radially from the outside in the openings 81 and 96 of the outer and inner walls and are fixed to these walls by appropriate means.

Easements 104 can be mounted inside these arms.

The sectors of the sheet metal cowling 94 (FIG. 8) are then fixed at their upstream ends to the downstream annular structure 84 of the frame 70 of the internal wall, so that the rods 98 pass through the aforementioned notches 95 of this cowling.

This cowling 94 externally surrounds the upstream part of a mixer 106 which is shaped to ensure the mixing of the gases ejected from the turbines, around an exhaust cone 108, with the air flow leaving the fan duct.

CA 02602172 2007-09-18 11 The process then consists (figure 9) in placing and fixing the removable panels 72 on the framework 70 of the internal wall so as to close the windows 90 of this framework, then finally to close by the removable panels 66 the openings 64 of the support structure 62 (Figure 10).

The method for disassembling the fan duct according to the invention consists in repeating the aforementioned steps in the reverse order.

To carry out maintenance operations on the equipment mounted on the body of the turbojet, such as in particular fuel injectors 110 and actuators 112 for controlling variable-pitch blades (FIG. 5), it suffices only to dismantle and remove the panels 66 of the outer wall and then dismantle and remove the panels 72 of the inner wall through the openings 64 of the outer wall.