Attachment pylon for an aircraft turbo-shaft engine with concentric hot air channels

A pylon for attaching a turbo-shaft aircraft engine designed to be laterally offset on a rear part of the aircraft structure, where the pylon possesses an aerodynamic profile which includes a leading edge, as well as a first hot air distribution channel passing along the leading edge in order to provide anti-icing, where this engine attachment pylon includes moreover a second hot air channel belonging to a system for supplying pressurized air to the aircraft. The second channel is housed inside the first channel.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates in general to an engine assembly for an aircraft, of the type designed to be installed laterally on a rear part of the aircraft structure and which incorporates a turbo-shaft engine, for example of the turbojet or turbo-prop type.

2. Discussion of the Background

Such an engine assembly includes an attachment pylon designed to provide the interface between the turbo-shaft engine and the rear part of the fuselage, where this pylon usually forms an aerodynamic profile whose leading edge is equipped with anti-icing means. These means draw off hot air at the turbo-shaft engine that is then blown towards the internal surface of the leading edge through a hot air distribution channel running along this leading edge in order to provide anti-icing of the latter. This principle is also referred to as an “anti-icing” effect.

In addition, the rear part of the aircraft also includes system for supplying pressurised air, which incorporates a channel for hot air also generally drawn off the turbo-shaft engine. This channel usually leads through the engine attachment pylon, away from the means used for anti-icing, to be connected to a heat-exchanger also fed by cold air drawn from the exterior of the aircraft.

The multiplicity of the number of these channels only accentuates the recurrent problems of space occupied within the engine attachment pylon, a part of the aircraft that is in fact crossed by a multitude of items of equipment. Furthermore, this confinement within the pylon creates a heat-risk zone in the event of one of the channels rupturing, or in the event of a hot air leak.

SUMMARY OF THE INVENTION

The purpose of the invention is therefore to propose an engine attachment pylon which remedies, at least in part, the above mentioned disadvantage associated with embodiments of the prior art.

To achieve this, the object of the invention is a pylon for attaching an aircraft turbo-shaft engine designed to be laterally offset on a rear part of the aircraft structure, where said pylon exhibits an aerodynamic profile which includes a leading edge, as well as a first hot-air distribution channel which passes along the leading edge in order to provide anti-icing of the latter, where said attachment pylon includes in addition a second hot air channel belonging to a system for supplying pressurised air for the aircraft. According to the invention the said second channel is housed inside the said first channel.

By placing the two channels one inside the other rather than in an adjacent manner, the occupied space associated with their presence is greatly reduced, and consequently space is freed up for other equipment to pass through the engine attachment pylon. In addition, within the engine attachment pylon, the heat-risk zone associated with the presence of hot air channels is less extensive.

Furthermore, by fitting these two channels one inside the other, the heat losses experienced by the hot air passing through the channels are appreciably reduced.

The first and second channels are each preferably cylindrical with circular cross-sections and are concentric, although other shapes can be adopted whilst still remaining within the scope of the invention.

The first channel for distribution of hot air preferably has multiple hot air distribution holes made in it, preferably arranged facing the leading edge.

The first and second channels are each preferably fed by hot air drawn off the turbo-shaft engine.

The first and second channels are preferably respectively fed by first and second pipes which emerge from the same channel which draws hot air off the engine. In such a case, means of controlling the apportionment of hot air flowing through the first and second pipes are preferably envisaged.

Another object of the invention is also an aircraft engine assembly which includes a turbo-shaft engine as well as an engine attachment pylon as described above which carries the said turbo-shaft engine and which is designed to be laterally offset on a rear part of the aircraft structure.

The turbo-shaft engine of the assembly may be a turbojet engine or a turbo-prop engine with a single propeller or two contra-rotating propellers, for example of the “open rotor” type.

Furthermore, the invention relates to a rear part of the aircraft which includes at least one engine assembly as described above, laterally offset on the fuselage of the aircraft.

Finally, one object of the invention is an aircraft which includes a rear part as described above.

Other advantages and characteristics of the invention will appear in the detailed non-restrictive description below.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

With reference toFIG. 1, a rear part of the aircraft can be seen which is in the form of a preferred embodiment of this invention.

Throughout the entire description that follows, by convention the longitudinal direction of the aircraft, which is parallel to a longitudinal axis of this aircraft, is called X. Moreover, the direction which is aligned transversely in relation to the aircraft is called Y, and the vertical direction or height is Z, with these three directions X, Y and Z being orthogonal to each other.

Moreover, the terms “forward” and “rear” are to be regarded in relation to a direction of forward motion of the aircraft experienced as a result of the thrust exerted by the engines, with this direction being schematically represented by arrow4.

Overall, the rear part1includes a fuselage6with an approximately circular, elliptical or similar transverse section, whose centre passes through the longitudinal axis2and which delimits an internal space of the aircraft8.

In addition, it includes two engine assemblies100, arranged on either side of a vertical median plane P passing through the axis2, where each assembly100includes a turbo-shaft engine10with propellers, preferably a turbo-prop engine of the “open rotor” type which includes a pair of contra-rotating propellers. Each of these has a longitudinal axis12which is approximately parallel to the direction X. Moreover, each engine assembly100is arranged laterally in relation to the fuselage6, it being specified as regards to this that an angle may be envisaged between the median horizontal plane P′ of the aircraft and the plane passing through the longitudinal axes2,12of the turbo-shaft engine and of the aircraft. Typically, this angle may be between 10 and 35°. In any event, each engine assembly100is laterally offset on the structure of the aircraft, more precisely on a rear part of the latter, on the fuselage6behind the main wing.

An engine attachment pylon14is envisaged for suspending the turbo-shaft engine of each assembly, where this includes a rigid structure15, also called the primary structure, through which the forces produced by the engine10are transmitted, where the rigid structure is covered by aerodynamic fairings, in particular a front fairing16which forms a leading edge. Analogously, a rear fairing17is envisaged which forms a trailing edge of the pylon.

In a conventional manner, the pylon14has means of fixing (not shown) interposed between the engine10and the rigid structure15, as well as other means of fixing (not shown) interposed between the rigid structure15and the structure of the aircraft.

InFIG. 1it can be seen that the pylon14presents an external surface20which forms an aerodynamic profile which incorporates the leading edge16, as well as a trailing edge17located to the rear of the rigid structure15. Thus this external surface20is formed, from the front towards the rear, of the front fairing16which forms a leading edge, the rigid structure15, preferably fitted with a fairing, and the rear fairing17, which forms the trailing edge.

It should be noted that the leading edge16, and more precisely the fairing which defines it, here take the form of a skin, possibly a double skin.

With reference now toFIG. 2which shows one of the two engine assemblies100in a schematic exploded manner, it is shown that the engine assembly includes means which provide anti-icing of the leading edge16, which essentially include a first hot air distribution channel, reference30, which passes along this leading edge16and internally with reference to the leading edge. This first channel30is approximately cylindrical and of circular cross-section, therefore parallel to the leading edge16, with a number of hot air distribution holes32made in it spaced apart from each other in the direction of the wingspan, and which are made facing the front end of the leading edge16.

Furthermore,FIG. 2shows a pressurised air supply system for the aircraft, reference34. This system includes a second hot air channel36which also passes through the engine attachment pylon14, along the leading edge16. This second channel36supplies hot air to a heat exchanger38belonging to the system34, which is also supplied by an intake of cold air drawn from outside the aircraft. In a known manner, at the outlet42of this system34, connecting with the cold air intake40, pressurisation air is supplied which is destined to be introduced to the interior of the aircraft, in particular inside the aircraft's passenger cabin. In addition, the pressure, flow and temperature are regulated at the outlet42′ of the heat exchanger which is connected to the second hot air channel.

In a known manner, this heat exchanger38may be housed in the structure of the aircraft.

As is shown schematically inFIG. 2, the first channel30is supplied with hot air by a first pipe30a, whilst the second hot air channel36is supplied by a second pipe36a, with both these pipes30a,36a, merging at a downstream end of a channel for drawing off hot air44connected to the turbo-shaft engine10. Thus the design that is used advantageously envisages only a single channel44for drawing off hot air connected to the turbo-shaft engine, and which at its downstream end has a fork which allows simultaneous supply of both pipes30a,36a, which themselves respectively supply the first hot air distribution channel30which provides anti-icing, and the second hot air channel36which belongs to the system34for supplying pressurised air to the aircraft. One envisaged alternative (not shown) involves providing two draw-off channels which respectively supply the two pipes30a,36a. In both cases, one or more valves may be envisaged which are used to control the amount of air passing through the pipes30a,36a.

As may be seen more clearly inFIGS. 3 and 4, one of the specific features of this invention rests in the fact that the second channel36is housed inside the first hot air distribution channel30, with these two channels preferably being concentric. Thus the hot air which arrives through the pipe30ais introduced through one or more inlets50into an annular space, formed between the internal wall of the first channel30and the external wall of the second channel36. This flow of air52passing through the annular space54is blown in the direction of the forward end of the leading edge, inside the latter, by means of distribution holes32made through the first channel30.

In parallel, the hot air from the pipe36ais introduced into the channel36, thus creating a flow56which extends through the engine attachment pylon up to the heat exchanger38. Thus the hot air flows from the distal end of the engine attachment pylon towards the proximal end of the latter, within both channels30,36.

Moreover, it should be noted that one or more air outlets (not shown) made in the leading edge16allow the hot air leaving the distribution holes32to leave once this hot air has contributed to the anti-icing of this leading edge16.

InFIG. 4an annular space54is shown, deliberately magnified in relation to that preferentially used. It is in fact arranged for the second channel36to be inserted into the first channel30, with the clearance allowing the one to be assembled inside the other, thus forming the annular space through which the flow52passes before being blown through the distribution holes32. In other words, the second channel36is mechanically supported in relation to the first channel30by the fitting of the one inside the other, preferably without any additional fastening components.

InFIG. 3a flow regulation valve located on the pipe30ahas been shown. This valve58is used to control the apportionment of hot air flows through the first and second pipes30a,36a. In particular, when anti-icing is not required for the leading edge16, the valve58is fully closed so that all the hot air coming from the supply channel44is discharged into the pipe36awhich supplies the second channel36which belongs to the aircraft's pressurised air supply system.

Naturally, various modifications may be made by those working in this field to the invention that has just been described in the form of non-restrictive examples only.