FR3039202A1 - AIRCRAFT COMPRISING A TURBOMACHINE INTEGRATED WITH REAR FUSELAGE WITH VARIABLE POWER SUPPLY - Google Patents

Abstract

The invention relates to an aircraft propelled by a turbomachine with counter-rotating fans (7, 8), the turbomachine being integrated into the rear of a fuselage (1) of the aircraft, in the extension thereof and comprising at least two gas generators (2a, 2b) which feed a power turbine (3) via a common central duct (4), the turbine (3) comprising two counter-rotating rotors (5, 6) for driving two blowers (7, 8) disposed downstream of the gas generators (2a, 2b), characterized in that it comprises means (15) arranged for separating the gas flow in the power turbine (3) into at least two concentric veins ( 16, 17) and a device for distributing the gas flow (21-24) between these veins (16, 17) from the central vein (4), said dispensing device being configured to open or close the supply at least one vein (16), said occultable, among the veins (16, 17) of the power turbine (3).

Description

Field of the invention and state of the art:

The present invention relates to the field of aircraft such as aircraft, including civil aircraft, propelled by a turbomachine integrated in the extension of the fuselage downstream thereof. It relates more particularly to the means for adapting the turbomachine in this type of architecture to particular operating situations which must be taken into account for safety reasons.

It has been proposed in the patent application FR-A1-2 997 681, a new aircraft architecture for reducing noise and fuel consumption of the aircraft by limiting the aerodynamic drag by absorption of the boundary layer.

In such an architecture, an aircraft is propelled by a turbomachine with contra-rotating fans careened, the turbomachine being integrated into the rear of the fuselage of the aircraft. Generally, the turbomachine comprises at least two gas generators which feed a power turbine having two counter-rotating rotors for driving two blowers disposed downstream of the gas generators. The gas generators have separate side air intakes to supply each of them.

Downstream of the gas generators, the blowers are arranged in the extension of the fuselage of the aircraft and generally fed by an annular ring connected thereto, so as to absorb at least a portion of the boundary layer formed around the fuselage. The diameter of the blowers is of the order of that of the fuselage in its largest section. The speed of rotation of the blowers is generally lower than for conventional turbomachines, especially so that the speed at the head of the blade is subsonic.

The power turbine is sized to operate by being powered by all gas generators. In the case where one of the gas generators is extinguished for a voluntary or accidental cause, the flow entering the turbine is substantially reduced. For example, it is practically halved in the case of two gas generators, only one operating. The passage section of the vein of the turbine is then too large for the flow rate considered and the speed of the flow is greatly reduced. The power supplied by the turbine, depending on the flow rate and the speed of the flow that passes through it, becomes too low and is, for example, much smaller than that obtained for an aircraft with two separate turbomachines when one of the turbines is stopped. they.

The object of the present invention is to adapt the turbine to the case where one of the gas generators is extinguished and at least one remains in operation.

DESCRIPTION OF THE INVENTION For this purpose, the invention relates to an aircraft propelled by a turbomachine with counter-rotating fans, the turbomachine being integrated into the rear of a fuselage of the aircraft, in the extension thereof and comprising at least two gas generators which supply, by a common central stream, a power turbine, the turbine comprising two counter-rotating rotors for driving two blowers disposed downstream of the gas generators, characterized in that it comprises means arranged to separating the flow of gas in the power turbine into at least two concentric veins and a device for distributing the flow of gas between these two veins from the central vein, said dispensing device being configured to provide for opening or closing the supply at least one vein, said occultable, among the at least two concentric veins of the power turbine.

Said at least one occultable vein preferably has a passage section substantially equal to the total flow section in the turbine divided by the number of gas generators. In this way, when a gas generator is stopped and the other gas generators operate, the device makes it possible to maintain a counter-rotating turbine in degraded mode where only the section of each blade present in the unobstructed vein (s) is subjected to flux. As the overall passage section of the unobstructed vein or veins corresponds to the flow rate provided by the gas generator (s) remaining in operation, the speed of the flow to which the blades are subjected is substantially the same as in normal operation. The power supplied is decreased because of the lower flow rate but it does not add the effect of a significant drop in the speed of the incident flow on the blades.

Moreover, the fact that the dispensing device is in the central common vein causes that there is no vein dedicated to one or the other of the gas generators, their flows are mixed in the central vein. The dispensing device therefore makes it possible to adapt the gas passage section in the power turbine to the gas generators remaining in operation, regardless of which one is stopped.

Preferably, each rotor of the power turbine comprising at least one blade crown, the flow separation means in concentric veins, comprise fins extending circumferentially between the vanes of each ring, at an intermediate radius between the ends. radial of these blades.

Advantageously, the dispensing device comprises passages of the gas flow dedicated to the supply of gas flow coming from the central vein of each of the at least two concentric streams of the power turbine and integral with a fixed structure of the turbomachine. .

Preferably, said concentric radial vein the innermost in the power turbine is an occultable vein. It is therefore preferentially the outer part of the blades which remains subjected to an incident flow when the turbine operates in degraded mode.

According to a preferred embodiment, said at least one occultable vein comprises an inlet opening in the central vein defined between two edges substantially along the same curve about an axis of the turbomachine and offset along said axis.

Preferably, the dispensing device comprises a part movable in translation along an axis of the turbomachine and is configured such that said part opens or closes the inlet section of said at least one occultable vein of the power turbine according to its position in translation. Thus, having no rotating part, the dispensing device can have a simple design, with actuators that can be carried in the fuselage, where the problem of congestion is less severe.

Preferably, the aircraft comprises means configured to feed with a flow of ventilation air said at least one occultable vein of the power turbine when its supply of gas through the central vein is closed.

Indeed, the separation means concentric veins of the power turbine are not perfect. They can let through a leakage current. The ventilation flow purges the concealable vein when closed and counteracts the tendency of the primary gases to escape into the occludable vein through the interstices between the fins.

Advantageously, the aircraft comprises means configured to close the exit of said at least one occultable vein of the power turbine when its supply of gas through the central vein is closed. This prevents the ventilation flow from being sucked in by the rotation of the blades of the turbine. It can escape only through the interstices between the fins and thus fulfills its purging effect.

Advantageously, the control of the closing means at the outlet of said at least one occultable vein is a function of the difference between an elastic restoring force and a pressure of the gases leaving said at least one occultable vein. Indeed, the pressure exerted by the ventilation flow is much lower than that exerted by the primary flow when the vein is open. It is therefore easy to adjust the return force so that the means are closed when closing the supply of the vein and they open without impeding the operation of the turbine when the supply of the vein in primary flow is effective.

Advantageously, supporting arms of a downstream casing, rotating with a rotor of the power turbine, distributed in a ring at the outlet of said turbine, support said closing means.

In this case, said closure means may comprise flexible blades, each being attached to a support arm in the area of the leading edge and extending in the circumferential direction to a neighboring support arm when no force is exerted on a direction from said at least one occultable.

Finally, the invention relates more particularly to an aircraft in which the generators are powered by separate air inlets disposed at the periphery of the fuselage of the aircraft.

Brief description of the figures:

The present invention will be better understood and other details, features and advantages of the present invention will emerge more clearly on reading the description of a nonlimiting example which follows, with reference to the appended drawings in which:

FIG. 1 presents a schematic view in longitudinal section of the rear part of an aircraft equipped with a turbomachine according to the invention;

FIG. 2 shows a schematic view of a longitudinal half-section of the turbomachine according to the invention at the level of the power turbine in a normal operating mode;

FIG. 3 shows a schematic view from above of a longitudinal half-section of the turbomachine according to the invention at the level of the power turbine in a degraded operating mode;

FIG. 4 is a schematic view in longitudinal section of an embodiment of an adjustable distribution device for the primary flow at the inlet of the power turbine;

Figures 5a and 5b schematically show a circumferential view deployed between two adjacent support arms, a means for blocking the primary flow at the outlet of a vein of the power turbine in two operating situations.

DESCRIPTION OF AN EMBODIMENT The invention applies in particular to an aircraft such as an airplane comprising a turbomachine of the type shown in FIG.

As shown in FIG. 1, the turbomachine is centered on the longitudinal axis XX of the fuselage 1 of the aircraft. This turbomachine comprises, from upstream to downstream in the direction of gas flow, two separate gas generators 2a, 2b simultaneously supplying a single power turbine 3. The turbine engine is installed at the downstream end of the fuselage 1 of the aircraft.

In this document, the axial and radial designations refer to the axis XX of the fuselage and the turbomachine. Similarly, upstream and downstream terms refer to the direction of the main flow along this axis.

In a manner known per se, each gas generator 2a, 2b comprises at least one compressor, a combustion chamber and a turbine (not shown in the figures).

Each gas generator 2a, 2b is housed inside a primary flow vein 3a, 3b. Separate air inlets 4a, 4b are provided for these veins 3a, 3b to supply each gas generator 2a, 2b. In the example shown, the air inlets 4a, 4b are connected to the fuselage 1 of the aircraft, upstream of the gas generators 2a, 2b, and their inner wall is directly integrated with the fuselage 1. They thus absorb a portion of the boundary layer formed around the fuselage 1 of the aircraft. In another configuration, not shown, the lateral air inlets feeding each of the gas generators may, on the contrary, be spaced from the fuselage 1 of the aircraft, so as to minimize this phenomenon of absorption of the boundary layer and to facilitate the operation of gas generators. It is also conceivable to use more than two gas generators, for example three to supply the power turbine 3.

Preferably, the two primary flow veins 3a, 3b of the gas generators 2a, 2b converge on the longitudinal axis XX and form between them an open V upstream, the opening angle is preferably included between 80 ° and 120 °.

The two primary flow streams 3a, 3b of the gas generators 2a, 2b converge in a central primary stream 4 which feeds the power turbine 3. A mixer (not shown in the figures) is preferably positioned at the level of the zone convergence of the two veins 3a, 3b, housing the gas generators 2a, 2b. This mixer has the function of mixing the gas flows from the two gas generators 2a, 2b to create a single homogeneous gas stream at the outlet of the primary central vein 4.

The power turbine 3, which is fed by this primary flow output of the central vein 4, is provided with two rotors 5, 6 counter-rotating turbine to drive contrarotatively two blowers 7, 8. These turbine rotors 5, 6 are coaxial and centered on the longitudinal axis XX. They revolve around a central casing 9 fixed to the structure of the aircraft.

Here, a first turbine rotor 5 corresponds to blades connected to a tubular body 5a separating the primary flow stream, in the power turbine 3, from the secondary flow duct, in which the fans 7.8 are located. . The blades and the tubular body 5a of the first rotor 5 are connected to the support bearings of the rotor 5 on the inner casing 9 by support arms 10 which pass through the primary vein upstream of the power turbine 3.

In the same example, the second rotor 6 corresponds to blades connected to a radially inner wall of the primary stream in the turbine 3 and inserted longitudinally between the vanes of the first rotor 5.

Downstream of the power turbine 3, the radially inner portion of the second rotor 6 is extended by a central body 11. On the other hand, it is connected, by support arms 12, to a ring 13 for supporting the blades of the In addition, this ring 13 extends the tubular body 5a of the first rotor 5 and has a rearward extension, so as to form, with the central body 11, a primary discharge nozzle, at the outlet of the power turbine 3.

In the example presented, a first fan 7, upstream, is positioned at the inlet of the power turbine 3. It is connected to the first rotor 5 at the arms 10 which support upstream cylindrical outer body 5a.

This upstream fan 7 therefore rotates at the same speed as the first rotor 5 of the power turbine 3.

In the same example, the second fan 8, downstream, is positioned at the output of the power turbine 3. It is connected to the second rotor 6 at the level of the support ring 13 and arms 12 which support it. This downstream fan 8 therefore rotates at the same speed as the second rotor 6 of the power turbine 3.

The two blowers 7, 8 are careened by a nacelle 14 fixed to the structure of the aircraft. This nacelle 14 is in particular fixed here to the vertical tail of the aircraft, not shown in the figures. The blowers have an outer diameter D which corresponds substantially to the outermost diameter of the fuselage 1 of the aircraft. The air entering the blowers 7, 8 is partly composed of the fuselage boundary layer of the aircraft, the input speed is low compared to conventional turbomachine blowers and the output speed is also lower at identical compression ratio, which improves the propulsive and acoustic performances of these blowers. Moreover, the large outer diameter D of the blowers 7, 8 causes their rotational speed, like that of the rotors 5, 6 of the power turbine 3, will also remain low compared to a conventional turbomachine. According to a first aspect of the invention, with reference to FIG. 2, the turbine blades of the two rotors 5, 6 as well as the support arms 10, 12 of these rotors which pass through the primary flow, are equipped with fins 15 having a circumferential extension substantially at the halfway point of the radial ends of the primary flow vein.

The shape of the fins 15 makes it possible to separate the vein from the primary flow into a radially internal vein 16 and a radially external vein 17, the two veins being of revolution about the longitudinal axis XX. FIG. 4 makes it possible to represent the shape of these fins 15 on a support arm 10 of the first rotor 5. Such fins 15 form here the separation between the inner and outer veins 16 and 17, in a continuous manner in the circumferential direction but leave longitudinally a space between them, so as to allow the differential rotation of the successive stages of the turbine. In this, they are comparable to the ferrules 18, 19, which are located at the foot and at the end of the blades or arms in the turbine 3 and which reconstitute the inner and outer walls of the passage of the primary flow in the power turbine 3 These ferrules 18, 19 and the fin 15 are shown on the support arm 10 of the first rotor 5 in FIG.

Referring to Figure 4, a device is set up to distribute the primary flow from the central vein 4 between the internal veins 16 and outer 17 of the power turbine 3.

Here, a ring of structural arms 20, inclined relative to the longitudinal axis XX, keeps the central casing 9 of the power turbine 3 to the fixed structure of the upstream turbomachine. The direction perpendicular to these structural arms defines a cone of divergence from the central vein 4 to the veins 16, 17 of the power turbine 3.

This ring of structural arms 20 makes it possible to hold a series of frustoconical ferrules 21, 22, 23 flaring towards the rear. An outer frustoconical ferrule 22 forms the connection between the outer radial wall of the central vein 4 and the outer radial wall of the external vein 17 of the turbine 3. An inner frustoconical ferrule 21 forms, starting from an upstream edge 29, here substantially circular and having a determined radius R, is connected to the radially inner wall of the internal duct 16 of the turbine 3. An intermediate frustoconical ferrule 23 is also fixed on the support arms 20 so as to separate the vein between the two rings 21, 22 end of two veins of substantially equal section and to be connected in its downstream portion to the fins 15 on the arm 10 of the first rotor 5. In this way, the frustoconical ferrules 21, 22, 23 extend upstream. internal veins 16 and outer 17 of the power turbine 3.

Preferably, the intermediate frusto-conical ferrule 23 extends forwards in the central vein 4, at a constant inclination, to an upstream edge 30 located at an axial distance L upstream of the upstream edge 29 of the inner ferrule. 21. The upstream edge 30 of the intermediate shell 23 is substantially circular, with a radius R substantially equal to that of the upstream edge 29 of the inner shell 21.

Moreover, a dome-shaped piece of revolution 24 closes the central vein 4 to the determined radius R at which the inner frustoconical ferrule 21 stops. In correspondence with this characteristic, the part of revolution 24 is movable in translation along of the longitudinal axis XX. It can be moved by actuators 25, shown in Figures 2 and 3 and can find in the fuselage 1 of the aircraft in front of the turbomachine. It can be positioned, as is illustrated in FIG. 4, either in a position where it is connected with the upstream edge 29 of the inner shell 21, or in a position where it connects with the upstream edge 30 of the intermediate ferrule 23.

Referring to Figure 2, when the part 24 is positioned rearwardly by the actuators 25, the primary air flow from the central vein 4 separates, as indicated by the two arrows from the central vein, in two flows distributed between the inner veins 16 and outer 17 of the power turbine 3. This corresponds to the case, where the flow from the central vein 4 is the output mixture of the two gas generators 2a, 2b in operation.

With reference to FIG. 3, when the part 24 is positioned forwardly, in the position 24 ', by the actuators 25, the primary air flow coming from the central vein 4 passes only into the external vein 17 of FIG. the power turbine 3, as indicated by the arrow from the central vein 4. This corresponds to the case, where the flow from the central vein comes only from a gas generator, 2a or 2b.

According to another aspect of the invention, the power turbine 3 is arranged to introduce a flow of ventilation air Fv in the internal vein 16 when its inlet is closed. With reference to FIG. 3, this flow of ventilation air Fv can be taken upstream of the turbine 3 by an unrepresented air inlet and be conveyed to an inlet 26 in the upstream portion the internal vein 16 while passing through the structural support arms of the central casing 9. As shown by the arrow in FIG. 2, this inflow of ventilation air into the internal vein 16 is closed when the internal vein 16 is in communication with the vein central primary 4.

Furthermore, this ventilation air inlet 26 can be controlled by a valve system which opens or closes automatically, depending on the pressure applied by the primary flow, depending on whether or not it passes into the vein. internal 16.

Preferably, when the primary air inlet in the internal vein 16 of the turbine 3 is closed, closing means 27 of the outlet of the internal vein 16 are actuated. In this way, the ventilation air can escape only via the space between the fins 15, and can thus fulfill its purge function.

Of course, these means 27 must be open when the power turbine 3 operates with its two veins 16, 17 through which the primary flow flows.

A nonlimiting example of a device for producing these closure means 27 so that they close and open automatically as a function, respectively, of the closure and opening upstream of the internal vein 16, is described on Figures 5a and 5b.

With reference to FIG. 5a, flexible blades 28 of the height of the radial extension of the internal vein 16 are fixed to each support arm 12 of the second rotor 6, at the rear of the turbine 3. These flexible blades 28 have a length along the longitudinal axis XX of the same order as that of the arms 12 and a value greater than the spacing between two successive arms 12 in the circumferential direction. They are placed on the underside of each arm 12 relative to the direction of rotation of the primary flow, as shown by arrows in thick lines in Figure 5a, and they are attached to each arm 12 at its level. leading edge 12a.

As shown in FIG. 5a, when the internal vein 16 is open and the primary flow, represented by arrows in thick lines, arrives on these blades 28, this flow has enough energy to press the blades 28 against the arms 12 of the second rotor 6, and thus release the space required for its passage.

By cons, with reference to Figure 5b, when the internal vein 16 is closed and it is the ventilation air, represented by arrows in fine lines, which arrives on these blades 28, this air does not have enough These blades 28 are thus relaxed from the leading edge 12a of each arm 12 and their rear end bears against the arm 12 adjacent. They thus form an obstruction of the downstream outlet of the internal vein 16 and force this ventilation air to escape through the spaces between the fins 15, towards the external vein 17, in the power turbine 3.

Thus, a method for passing, on such an aircraft, in operation of the turbomachine with a single gas generator, for example in the event of accidental loss of one, may comprise essentially, at least with respect to the turbine of power 3, a displacement control of the central piece 24 forward, in position 24 ', to close the primary flow inlet in the internal vein 16. In this case, the ventilation air flow Fv is set up, with the opening of the air inlet 26 and the obstruction of the outlet of the internal vein 16 by the closing means 27, which are automatically controlled by the flow conditions. The invention has been presented in the case of two gas generators. In the case, not shown of several gas generators, for example three, a variant of the invention consists in modifying the distribution of the passage sections between the inner and outer veins 16 and 17 of the power turbine 3, so that that the remaining passage section when the internal vein 16 is concealed corresponds to the flow rate of the number of gas generators minus one. In the case of three gas generators, the ratios of the passage sections with respect to the overall passage section are respectively one-third and two-thirds, for the internal vein 16 and for the external vein 17. In this case, the ferrule intermediate 23 on the structural arms at the inlet of the turbine is brought closer to the inner ferrule 21 and the axial distance L from its upstream edge 30 to the upstream edge 29 of the inner ferrule 21 is reduced accordingly. To conceal the internal vein, the frustoconical piece 24 must therefore move a lesser distance but the operating principle of this variant remains identical to the previous one.

Both variants can operate with more than two gas generators. In particular, in the case of three generators, the second variant can operate if two gas generators stop, it allows a fall in the speed of the flow in the outer vein 16 which is less important than with the prior art and therefore decreases the power loss.

Finally, for example in the case of three generators, one can generalize the third variant by considering the turbine operating with two gas generators and the outer vein 16 and by applying to it the variant of the invention provided for two generators, it is that is, by subdividing the outer vein into two concentric veins.

Claims (10)

claims 1. Aircraft propelled by a turbomachine with contra-rotating blowers (7, 8), the turbomachine being integrated into the rear of a fuselage (1) of the aircraft, in the extension thereof and comprising at least two generators of gas (2a, 2b) which feed, by a central central duct (4), a power turbine (3), the turbine (3) having two counter-rotating rotors (5, 6) for driving two blowers (7, 8) arranged downstream of the gas generators (2a, 2b), characterized in that it comprises means (15) arranged for separating the gas flow in the power turbine (3) into at least two concentric veins (16, 17). ) and a device for distributing the gas flow (21-24) between these at least two veins (16, 17) from the central vein (4), said dispensing device being configured to open or close the supply of at least one so-called occultable vein (16), among the at least two concentric veins (16 17) of the power turbine (3). 2. Aircraft according to the preceding claim, wherein each rotor (5, 6) of the power turbine (3) having at least one blade crown, the concentric flow flow separation means (16, 17) comprise fins (15) extending circumferentially between the vanes of each ring, at an intermediate radius between the radial ends of these vanes. 3. Aircraft according to one of the preceding claims, wherein the radially innermost concentric vein in the power turbine is an occultable vein. 4. Aircraft according to one of the preceding claims, wherein said at least one occultable vein (16) comprises an inlet opening in the central vein (4) defined between two edges (29, 30) defined substantially by the same curve around an axis (XX) of the turbomachine and offset along said axis. 5. Aircraft according to the preceding claim, wherein the dispensing device comprises a part (24) movable in translation along the axis of the turbomachine, and it is configured such that said piece opens (24) or closes (24 ' ) the inlet section (29, 30) of said at least one occultable vein (16) of the power turbine (3) according to its translational position. 6. Aircraft according to one of the preceding claims, comprising means (26) configured to feed with a ventilation air flow (Fv) said at least one occultable vein (16) of the power turbine (3) when its gas supply through the central vein (4) is closed. 7. Aircraft according to the preceding claim, comprising means (27) configured to close the outlet of said at least one occludable vein (16) of the power turbine (3) when its supply of gas through the central vein (4) is closed. 8. Aircraft according to the preceding claim, wherein the control of the closure means (27) output of said at least one occultable vein (16) is a function of the difference between an elastic restoring force and a pressure of the gases at the output of said at least one occultable vein. 9. Aircraft according to the preceding claim, comprising supporting arms (12) of a downstream housing (13), rotating with a rotor (6) of the power turbine (3), distributed in a ring at the outlet of said turbine said support arms (12) supporting said closure means (27). 10. Aircraft according to the preceding claim, wherein said closure means (27) comprise flexible blades, each being attached to a support arm (12) in the area of the leading edge (12a) and extending into the circumferential direction to a neighboring support arm (12 ') when no force is exerted thereupon in a direction from said at least one occultable vein (16).