FR3039218A1 - TURBOMACHINE WITH CONTRAROTATIVE BLOWERS COMPRISING TURBINE BLADES DETACHABLES - Google Patents

Abstract

The invention relates to a turbomachine with counter-rotating fan rotors (7, 8), the turbomachine comprising at least two gas generators (2a) which feed a power turbine (3) comprising two counter-rotating rotors (5, 6) to drive two fan rotors (7,8), characterized in that the counter-rotating rotors (5, 6) of the power turbine (3) comprise vanes configured to detach or break when at least one of said rotors (5) , 6) rotates at a speed greater than a nominal operating speed of said turbomachine multiplied by a determined breaking factor α greater than 1, and in that the turbomachine is configured to provide a minimum thrust determined from the gases from the generators gas (2a, 2b) when said blades are detached. The invention also relates to an aircraft comprising such a turbomachine and a method of putting into degraded mode of thrust in the event of total loss of propulsion by the fan rotors.

Description

Turbomachine with counter-rotating fans comprising detachable turbine blades

Field of the invention and state of the art:

The present invention relates to the field of aircraft such as aircraft, in particular civil aircraft, propelled by a turbomachine comprising a propeller stage with counter-rotating fans, the turbomachine being integrated in the extension of the fuselage downstream thereof. It relates more particularly to means to ensure minimal thrust in the event of total failure of the propulsion stage.

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.

It may be noted that with this type of architecture, the aircraft may comprise a single propellant formed by the two counter-rotating fans of the integrated turbomachine at the rear of the fuselage. Under these conditions, for reasons of safety and certification of the engine, it is necessary that the turbomachine can provide a minimum thrust in the event of total failure of its thruster.

The present invention aims to provide a solution to this problem by limiting the problems of size, mass and complexity.

SUMMARY OF THE INVENTION: For this purpose, the invention relates to a turbomachine with rotors of contra-rotating fans, the turbomachine comprising at least two gas generators which supply a power turbine comprising two counter-rotating rotors for driving two fan rotors, characterized in that the counter-rotating rotors of the power turbine comprise blades configured to disengage or break when at least one of said rotors rotates at a speed greater than a nominal operating speed of said turbomachine multiplied by a breaking factor a determined higher than 1, and in that the turbomachine is configured to provide a determined minimum thrust from the gases from said at least two gas generators when said blades are detached.

The overspeed of the counter-rotating rotors occurs naturally at least during an accidental sequence of total loss of propulsion by the fan rotors corresponding to the loss of connection between the blowers and the rotors of the power turbine.

In this case, the programmed rupture of the blades of the rotors of the turbine during an overspeed makes it possible to disengage the primary vein duct to let the gases of the gas generators pass freely and thus recover a thrust of the turbomachine corresponding to the ejection gases from the primary stream. A suitable operating mode gas generators, provided for this purpose, allows to recover a minimum thrust to allow emergency maneuvers to the aircraft.

Preferably, said breaking factor a has a value between 1.25 and 1.6. The minimum limit prevents accidental blade losses in case of overspeed. In addition, the upper limit corresponds to a speed rapidly reached during an accidental sequence to be able to quickly recover the minimum thrust.

Advantageously, said breaking factor α has a value which decreases as a function of the location of the blades, going from the upstream to the downstream direction along the direction of the gas flow.

The upstream and downstream terms refer to the flow of gases passing through the turbomachine.

This characteristic makes it possible for the breaking phenomenon to start downstream, which makes it easier to disengage the detached blades without becoming entangled.

Preferably, at least one of the fan rotors, and in particular the rotor of the downstream fan, comprises vanes whose setting can be placed at a given angle, determined so as to drive the corresponding power turbine rotor in overspeed, at a higher speed than the nominal speed multiplied by the breaking factor a.

This allows, in an accidental sequence of loss of propulsion by the fan rotors but where the connection with the rotors is maintained, to drive said rotors in overspeed so as to engage the detachment phenomenon of the blades.

Said at least one turbine rotor driven in overspeed may comprise blades fixed radially from the inside.

Preferably, the fan rotors are located at the outer periphery of the counter-rotating rotors of the power turbine.

Furthermore, the invention also relates to an aircraft propelled by a turbomachine described above, in particular an aircraft in which the power turbine is installed at the rear of a fuselage of the aircraft, in the extension thereof. The invention also relates to a method for controlling an aircraft as described above, comprising, in the case of a total loss of propulsion by the fan rotors, a first step in which at least one of said rotors of the power turbine is driven in overspeed so as to obtain rupture or detachment of the blades of said rotors of the power turbine, and a step of speeding the gas generators so as to obtain a determined minimum thrust of the turbomachine.

Advantageously, in such a method, the first step comprises a placement of the blade wedging of said at least one blower rotor, whose blade setting is variable at the given angle to cause overspeed.

Brief description of the figures:

The present invention will be better understood and other details, characteristics and advantages of the present invention will appear more clearly on reading the description of a nonlimiting example which follows, with reference to the appended drawings in which: FIG. a schematic view in longitudinal section of the rear part of an aircraft equipped with a turbomachine according to the invention; FIGS. 2a and 2b show schematic front views of vanes of turbine rotors according to the invention; FIG. 3 is a diagrammatic view of a longitudinal half-section of the rear part of an aircraft equipped with a turbomachine according to the invention, with devices for adjusting the setting of the vanes of the fan rotors; and FIG. 4 shows a schematic view of a longitudinal half-section of the rear part of an aircraft equipped with a turbomachine according to the invention, in degraded operating mode.

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 gas flow direction, two separate gas generators 2a, 2b simultaneously supplying a single power turbine 3. The turbomachine is installed at the downstream end of the fuselage 1 of the engine. '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 fan rotors 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 fan rotors 7 are located. 8. The vanes 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. As shown in FIG. 1 , the blades of said first turbine rotor 5 are distributed in several, for example four, successive rings.

As shown in FIG. 2a, in each blade ring of the first turbine rotor 5, each blade of the first rotor comprises a blade 15 configured to form, with the blades of the other blades, a grid transforming a part of the energy of the turbine. the flow of primary flow in a pair driving the first rotor 5 in rotation. In particular, the assembly is configured such that the first turbine rotor 5 rotates at a nominal speed by driving the first fan rotor 7 at full power during cruising flight conditions.

Each blade also has a foot 16 for attaching it to the tubular body 5a and may include a platform 17 embodying the radially outer wall of the primary flow stream at the blade. Here, according to a first aspect of the invention, each blade may comprise at its foot 16 a frangible section 18. This frangible section 18 may be made by a longitudinal constriction in the stilt of the foot 16, whose shape is calibrated according to a chosen rupture coefficient with respect to the forces experienced by the blade during operation of the turbomachine.

In the case of said first turbine rotor 5, the blades being fixed by their outer radial ends, the centrifugal forces create buckling forces which must withstand the attachment at the foot 16 to prevent the blade from bending. Here, the frangible section 18 is configured to break when said first turbine rotor rotates at a higher speed than a nominal speed multiplied by a determined breaking factor a greater than 1. Advantageously, this breaking factor has a value between 1 , 25 and 1.6. Preferably, the value of this breaking factor a may be lower for the blades of the grids located downstream than for those of the grids located upstream.

On the same example of a turbomachine, the second rotor 6 may comprise a tubular body 6a radially internal to the primary stream in the turbine 3 and on which vanes are fixed. The blades of the second rotor 6 are also distributed in rings, which are interposed between the crowns of the blades of the first rotor 5 along the axis XX. As shown in FIG. 1, said second turbine rotor 6 may comprise the same number, for example four, of rings as the first rotor 5.

As represented in FIG. 2b, in each blade crown of the second turbine rotor 6, each blade comprises a blade 19 configured to form, with the blades of the other blades, a grid transforming a portion of the energy of the flow. primary flow in a pair driving the second rotor 6 in rotation. In particular, the assembly is configured such that the second turbine rotor 6 rotates at a nominal speed by driving the second fan 8 at full power during cruising flight conditions.

Each blade also has a foot 20 for attaching it to the tubular body 6a of the second rotor 6. The blade may comprise a platform 21 embodying the radially inner wall of the primary flow stream at the blade. Here, as in the case of the first rotor 5, the blade may comprise at its foot 20 a frangible zone 22. This frangible zone 22 can be achieved by a longitudinal constriction in the stilt of the foot 20, whose shape is calibrated according to a breaking factor a 'determined with respect to the forces experienced by the blade during operation of the turbomachine.

In the case of said second turbine rotor 6, the blades being fixed by their inner radial ends, the centrifugal forces create tensile forces which must withstand the fixing at the foot 20 to prevent the blade from being detached. Here, the frangible zone 22 is configured to break when said first turbine rotor 5 rotates at a higher speed than a nominal speed multiplied by a predetermined breaking factor a 'greater than 1. Advantageously, as previously, this factor a' has a value between 1.25 and 1.6. Preferably, the value of this breaking factor a 'is lower for the blades of the grids located downstream than for those of the grids located upstream. Since the resistance to tensile stresses is different from that to the buckling forces, the frangible zones 22 of the blades of the second rotor 6 are different from the frangible zones 18 of the blades of the first rotor 5 for the same values of the breaking factors a and at'.

In association with the frangible zones 18, 22 formed on the blades of the rotors 5, 6, the tubular body 5a of the first rotor 5 is equipped with a shield designed to retain the blades 15, 19 which are detached from the rotors 5, 6 and 5. avoid causing damage by being radially ejected out of the power turbine 3.

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 rotor. of the downstream fan 8. 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 23, in output of the power turbine 3.

In the example shown, a first upstream fan rotor 7 is positioned at the inlet of the power turbine 3. It is connected to the first turbine rotor 3 at the arms 10 which support upstream the cylindrical outer body 5a. This upstream fan rotor 7 therefore rotates at the same speed as the first rotor 5 of the power turbine 3.

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

The two fan rotors 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.

Advantageously, the blades of the rotor of the downstream fan 8 are mounted with a device making it possible to vary their angular setting with respect to a meridian plane vis-à-vis the longitudinal axis XX. For this, with reference to Figure 3, they are mounted here with a system of rolling rings 24 on the support ring 13 and each rotate about a radial rod 25 which passes through a support arm 12 of the ring 13 A system 26 arranged on the axis XX and controlled from the upstream, for example by a system of rods passing inside the fixed housing 9, makes it possible to actuate the radial rods 25 to vary the angular setting of the blades. of the downstream fan rotor 8.

According to one aspect of the invention, the device 26 and the rotor blades of the downstream fan 8 are configured to allow them to be fixed angularly at a value corresponding to a very small pitch. A pitch angle corresponding to a very small pitch positions the vanes of the rotor of the fan 8 almost perpendicularly to the longitudinal axis XX of the turbomachine but the blades of the fan rotor 8 do not close the passage of the secondary flow. This very low pitch therefore corresponds to a very high rotational speed of the fan rotor 8 for a given longitudinal velocity of the secondary flow.

Similarly, the wedging angle of the rotor blades of the upstream fan 7 can be varied between a nominal value, corresponding to the normal operation of the rotor, the fan 7 in the propulsion mode, and a value providing a very small pitch, corresponding to The rotor blades of the upstream fan 7 are, for example, mounted on movable rings 28, at the level of the connection between the arms 10 and the cylindrical outer body 5a of the turbine. 3. Unlike the rotor of the downstream fan 8, the setting of the blade timing is not necessarily continuous. The mobile rings 28 may be equipped with a locking system associated with elastic return means. In nominal operation, the locking system maintains the setting of the blades so that the rotor of the upstream fan 7 performs its thruster function. In accidental mode, the system can be unlocked and the return means force the blades to stall in the position forming a very small pitch.

Some characteristics of the turbomachine according to the invention will now be detailed according to different modes of operation of the turbomachine.

In normal operation, the two gas generators 2a, 2b deliver gases that drive the two rotors 5, 6 counter-rotating the power turbine and rotate the two fan rotors 7, 8 to propel the aircraft. The gas generators 2a, 2b are thus configured to deliver a determined flow of gas so as to drive the counter-rotating rotors 5, 6 at a nominal speed corresponding to a reference thrust of the counter-rotating fan rotors to propel the aircraft. In this case, the primary flow gases escaping from the exhaust nozzle 23 at the outlet of the power turbine 3 provide only a small contribution to the thrust of the turbomachine.

In a first accidental sequence considered for the turbomachine, the connection between the two fan rotors 7, 8 and the corresponding turbine rotors 5, 6 is broken. Under these conditions, there is a total loss of propulsion normally provided by the fan rotors 7, 8. Furthermore, the turbine rotors 5, 6 no longer being engaged on the fan rotors are driven in overspeed by the gases from the gas generators 2a, 2b.

The speed of rotation of the rotors 5, 6 of the power turbine 3 quickly reaches values well above the nominal value of rotation multiplied by the breaking factor a for which are dimensioned the frangible zones 18, 22 of the blades of these rotors. Under these conditions, there occurs a phenomenon of "blade clearance" in the Anglo-Saxon literature and recalled, for example, in the patent application FR-A1-2 964 145. In this process , the blade of one of the rotor blades 5, 6 of the power turbine 3 is detached and then causes the blade rupture of the entire corresponding blade crown. The detached blades are retained radially in the turbine 3 by the shielding of the tubular body 5a and are then expelled from the primary vein, downstream.

According to an alternative embodiment, it is possible for the blades of only one rotor, for example the inner rotor 6, to be equipped with frangible zones 22 because the blades 19 detached from a rotor 6 will cause the blades 15 of the blades of the rotor blade to rupture. the other rotor 5, which are interposed axially.

According to another variant, if the failure factor decreases going from upstream to downstream, the blades 15, 19 of the vanes on the downstream rings start to detach first. The blades 15, 19 can thus be expelled from downstream to upstream, avoiding entanglement risks that could hinder their evacuation.

When most of the rotor blades 5, 6 are evacuated the primary stream in the power turbine 3 can pass freely the gases produced by the gas generators 2a, 2b. This ensures a residual thrust by the ejection nozzle 23 of the primary flow of the turbomachine.

According to another aspect of the invention, the gas generators 2a, 2b may also be designed to be able to operate at a higher speed than the speed necessary to ensure the thrust of the turbomachine in normal operation, with the fan rotors 7, 8. This higher speed of the gas generators 2a, 2b is determined so that the residual thrust obtained with the primary flow out of the exhaust nozzle 23 is equal to a minimum thrust allowing the aircraft to perform emergency maneuvers.

FIG. 4 illustrates such a degraded operating mode of the turbomachine where, optimally, the blades of the vanes of the counter-rotating rotors 5, 6 have all been disengaged. The only obstacles remaining in the primary vein of the power turbine 3 are the arms 10, 12, which hold the fan rotors 7, 8, as well as the tubular body 5a of the upstream rotor 5, thus guaranteeing the integrity of the vein primary to the ejection nozzle 23.

In another accidental sequence considered, the loss of propulsion by the fan rotors 7, 8 is total but the link is not broken with the rotors 5, 6 of the power turbine 3. In this case, according to the provisions described above, the setting of the vanes of the fan rotors 7, 8 is adjusted, by the corresponding means 26, 28, to the position imposing a small step. By the speed of the aircraft, the air flow passing through the fan rotors 7, 8 then drives them in overspeed with the turbine rotors 5, 6 which are linked to them. When the rotational speed exceeds the nominal speed of the rupture factor a corresponding to the sizing of the frangible zones 18, 22 of the vanes of the rotors 5, 6, the phenomenon of "disengagement of the vanes" starts, as in the accident sequence previously described.

Thus, with the same variants as previously mentioned for the configuration of the turbine rotors 5, 6, it is possible to release the gas passage of the gas generators 2a, 2b through the primary stream of the power turbine 3 and to recover a minimum thrust by the ejection nozzle 23 of the primary flow, in the event of total loss of propulsion by the fan rotors 7, 8.

Claims (9)

claims 1. Turbomachine with rotors of contra-rotating blowers (7, 8), the turbomachine comprising at least two gas generators (2a, 2b) which feed a power turbine (3) having two counter-rotating rotors (5, 6) for driving two rotors of blowers (7, 8), characterized in that the counter-rotating rotors (5, 6) of the power turbine (3) comprise blades (15, 19) configured to detach or break when at least one said rotors (5, 6) rotate at a higher speed than a nominal operating speed of said turbomachine multiplied by a breaking factor a determined greater than 1, and in that the turbomachine is configured to provide a minimum thrust determined from the gas from said at least two gas generators (2a, 2b) when said blades (15, 19) are detached. 2. Turbomachine according to the preceding claim, wherein said breaking factor a has a value between 1.25 and 1.6. 3. A turbomachine according to the preceding claim, wherein said breaking factor has a value which decreases as a function of the location of the blades (15,19), going from upstream to downstream in the direction of the flow. gases. 4. A turbomachine according to one of the preceding claims, wherein at least one of the fan rotors (7, 8), and in particular the rotor of the downstream fan (8) comprises vanes whose setting can be placed at a given angle, determined to drive the corresponding power turbine rotor (6) in overspeed, at a higher speed than the nominal speed multiplied by the breaking factor a. 5. A turbomachine according to the preceding claim, wherein said at least one turbine rotor (6) driven in overspeed has blades (19) fixed radially from the inside. 6. Turbomachine according to one of the preceding claims, wherein the fan rotors (7, 8) are located at the outer periphery of the counter-rotating rotors (5, 6) of the power turbine (3). 7. Aircraft propelled by a turbomachine according to one of the preceding claims, wherein the power turbine (3) is installed at the rear of a fuselage (1) of the aircraft, in the extension thereof. 8. A method of driving a turbomachine according to one of claims 1 to 6, comprising, in the case of a total loss of propulsion by the fan rotors (7, 8), a first step in which at least l one of said rotors (5, 6) of the power turbine (3) is driven in overspeed so as to obtain breaking or detaching of the blades (15, 19) of said rotors (5, 6), and a step of setting in operation of the gas generators (2a, 2b) so as to obtain a determined minimum thrust of the turbomachine. 9. Method according to the preceding claim, according to claim 4, wherein the first step comprises a placement of the blade wedging of said at least one fan rotor (7,8) at the given angle to cause overspeed.