Propulsive assembly having decouplable casing portions

Abstract

A propulsive assembly including a turbine engine configured to be fastened to an aircraft by means of a suspension, having certain casing portions suitable for being decoupled so as to attenuate certain bending modes of the turbine engine in the event of an incident, the propulsive assembly comprising at least first and second casing portions (33a, 22) extending axially one after another, one of said casing portions being cantilevered out relative to said suspension, wherein a first casing portion (33a) possesses a first fastener portion (41) connected rigidly to the body (43) of the first casing portion (33a) and a second fastener portion (42) connected more flexibly to the body (43) of the first casing portion (33a), and wherein the second casing portion (22) is fastened to the second fastener portion (42) of the first casing portion (33a) in a manner that is permanent, and to the first fastener portion (41) of the first casing portion (33a) in a manner that is releasable.

Claims

1. A propulsive assembly including a turbine engine configured to be fastened to an aircraft by means of a suspension, the assembly comprising at least first and second casing portions extending axially one after another: one of said casing portions being cantilevered out relative to said suspension; wherein the first casing portion possesses a first fastener portion connected rigidly to a body of the first casing portion and a second fastener portion connected more flexibly to the body of the first casing portion; and wherein the second casing portion is fastened to the second fastener portion of the first casing portion in a manner that is permanent during operation of the propulsive assembly, and to the first fastener portion of the first casing portion in a manner that is releasable during operation of the propulsive assembly.

2. A propulsive assembly according to claim 1, wherein one of said first and second casing portions that is cantilevered out relative to the suspension is cantilevered out relative to a casing portion that is directly connected to the suspension.

3. A propulsive assembly according to claim 2, wherein said casing portion that is directly connected to the suspension is the other one of said first and second casing portions.

4. A propulsive assembly according to claim 1, wherein the turbine engine comprises a rotor and a stator, the rotor including a fan driven by a shaft mounted to rotate relative to the stator via at least two bearings, and wherein the center of gravity of said casing portion that is cantilevered out relative to the suspension is situated axially further upstream than the most upstream of said bearings or further downstream than the most downstream of said bearings.

5. A propulsive assembly according to claim 1, wherein the first fastener portion of the first casing portion is connected to the body of the first casing portion by means of a first connection portion, and the second fastener portion of the first casing portion is connected to the body of the first casing portion by means of a second connection portion; and wherein the second connection portion possesses bending stiffness that is at least 10% less than the bending stiffness of the first connection portion.

6. A propulsive assembly according to claim 1, wherein the first fastener portion of the first casing portion is a first fastener flange; wherein the second fastener portion of the first casing portion is a second fastener flange; and wherein the second casing portion is fastened to the first and second fastener portions of the first casing portion by means of a common fastener flange.

7. A propulsive assembly according to claim 6, wherein the first and second fastener portions of the first casing portion are festooned.

8. A propulsive assembly according to claim 1, wherein the second casing portion is fastened in releasable manner to the first fastener portion of the first casing portion by using breakable fastener means.

9. A propulsive assembly according to claim 1, further including a nacelle, wherein one of said first and second casing portions is a casing portion of the turbine engine, while the other one of the casing portions is a casing portion of the nacelle.

10. A propulsive assembly according to claim 9, wherein said casing portion of the nacelle is an exhaust cone, a nozzle, or an air intake duct.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) The drawings are diagrammatic and seek above all to illustrate the principles of the invention.

(2) In the drawings, from one figure to another, elements (or portions of an element) that are identical are referenced by the same reference signs.

(3) FIG. 1 is a plane section view of a propulsive assembly of the invention.

(4) FIG. 2 is a section view of the interface zone between the rear turbine casing and the exhaust cone.

(5) FIG. 3 is an enlarged view of box III in FIG. 2.

(6) FIG. 4 is a plane view of fastener flanges of the rear turbine casing.

(7) FIG. 5 is a section view on plane V-V of FIG. 4.

(8) FIG. 6 is a section view on plane VI-VI of FIG. 4.

(9) FIG. 7 is a section view of the interface zone between the rear turbine casing and the exhaust cone after the fusible fastener means have broken.

(10) FIG. 8 is a graph showing how the moment at the flanges between the inter-turbine casing and the LP turbine casing varies as a function of the speed of rotation of the engine.

DETAILED DESCRIPTION OF EMBODIMENT(S)

(11) In order to make the invention more concrete, there follows a description in greater detail of an example propulsive assembly, which description is given with reference to the accompanying drawings. It should be recalled that the invention is not limited to this example.

(12) FIG. 1 is a section view of a propulsive assembly 1 of the invention on a vertical plane containing its main axis A. The assembly comprises a bypass turbojet 10 and a nacelle 20 surrounding the turbojet 10.

(13) From upstream to downstream in the flow direction of the air stream, the turbojet 10 comprises: a fan 12; a low pressure (LP) compressor 13; a high pressure (HP) compressor 14; a combustion chamber 15; a high pressure (HP) turbine 16; and a low pressure (LP) turbine 17. These various members are protected by a plurality of casing portions that are generally cylindrical and that are fastened to one another via fastener flanges. Among these various casing portions, the turbojet 10 has a fan casing 31 surrounding the fan 12, an inter-turbine casing 32 extending between the HP turbine 16 and the LP turbine 17, and a rear turbine casing 33 extending behind the LP turbine 17 and including an inner shroud 33a and an outer shroud 33b.

(14) The nacelle 20 has various casing portions serving in particular to guide the air stream: these various casing portions include in particular an air intake duct 21 located upstream from the fan casing 31 and guiding the air stream at the intake of the propulsive assembly 1 towards the fan 12; an exhaust cone 22 mounted downstream from the inner shroud of the rear turbine casing 33 and defining the inner envelope of the exhaust passage for the primary flow through the turbojet 10; and a nozzle 23 mounted downstream from the outer shroud of the rear casing of the turbine 33 and separating the exhaust passage for the primary flow from the turbojet 10 from the parallel secondary passage.

(15) Typically, the turbojet 10 is fastened to the aircraft via a suspension 9 that is connected to the turbojet via at least two points 9a and 9b: the suspension is thus often connected to the upstream end of the intermediate casing 8 and to the downstream end of the rear turbine casing 33 or to the inter-turbine casing 32.

(16) FIG. 2 shows the interface zone between the rear turbine casing 33 and the exhaust cone 22 in greater detail. For simplification purposes, the nozzle 23 is omitted from this figure.

(17) As can be seen better in FIG. 3, the inner shroud 33a of the rear turbine casing 33 possesses a first fastener flange 41 and a second fastener flange 42 at its rear end. More precisely, the body 43 of the inner shroud 33a is extended firstly in rectilinear manner to the first fastener flange 41 marking the downstream end of the inner shroud 33a; and secondly the second fastener flange 42 is connected to the body 43 via a ring 46 fitted to the inner surface of the casing body 43, or branched therefrom.

(18) Thus, on going from upstream to downstream, it can be said that the casing body 43 reaches a bifurcation 44 leading firstly to the first fastener flange 41 via a first connection portion 45 that is in fact constituted by the rectilinear extension of the casing body 43, and to the second fastener flange 42 via a second connection portion 46 that is constituted by the ring.

(19) In this example, the second connection portion 46 possesses thickness that is less than the thickness of the casing body 43 and thus less than the thickness of the first connection portion 45. Under such circumstances, the second connection portion 46 possesses bending stiffness that is less than the bending stiffness of the first connection portion 45. The second connection portion 46 also possesses a rounded portion 46a that also contributes to reducing its stiffness.

(20) The exhaust cone 22 is provided at its upstream end with a single fastener flange 51 that is fastened simultaneously but in independent manner to the first and second fastener flanges 41 and 42 of the inner shroud 33a of the rear turbine casing 33: these fastenings are described below in greater detail with reference to FIGS. 4, 5, and 6.

(21) As can be seen in FIG. 4, the first and second fastener flanges 41 and 42 of the rear turbine casing 33 are festooned, i.e. they present a circumferential succession of festoons 41a and 41b extending radially in disjoint manner. Each festoon 42a of the second flange 42 is thus separated from its neighbor by a setback 42b where it is possible to place a festoon 41a of the first flange; in the same way, each festoon 41a of the first flange 41 is separated from its neighbor by a setback 41b coinciding with a festoon 42a of the second flange 42. Each festoon 41a, 42a is provided with at least one bore 47, 48.

(22) The flange 51 of the exhaust cone 22 is a conventional flange extending radially continuously all around the upstream periphery of the exhaust cone 22. It has bores 52 coinciding with the bores 47, 48 in the festoons 41a, 42a.

(23) As can be seen in FIGS. 5 and 6, it is then possible to fasten the fastener flange 41 of the exhaust cone 22 in independent manner to the first fastener flange 41 of the casing 33 by using fastener means 55 of a first type, and to the second fastener flange 42 of the casing 33 by using fastener means 56 of a second type.

(24) The first fastener means 55 have the feature of being breakable, i.e. they break when they are subjected to a stress exceeding a certain threshold.

(25) Conversely, the second fastener means 56 are not breakable and cannot break or become undone while the propulsive assembly is in operation.

(26) In the present example, the second fastener means are conventional bolts 56, while the first fastener means are break bolts 55 each possessing a segment of reduced thickness that is capable of breaking when the moment on the flanges 41 and 51 exceeds a predetermined threshold, e.g. a threshold lying in the range 10.10.sup.3 newton-meters (N.m) to 50.10.sup.3 N.m for a two-spool bypass turbojet.

(27) Thus, in the event of an unusual moment appearing that acts on the interface between the rear turbine casing 33 and the exhaust cone 22, e.g. excited by an unbalance of the fan 12 caused by ingesting a bird, the break bolt 53 breaks so that the exhaust cone 22 then remains fastened to the rear turbine casing 33 via only the second fastener flange 42. The connection of the exhaust cone 22 to the body 43 of the rear turbine casing 33, and thus to the remainder of the propulsive assembly 1, then takes place exclusively via the flexible force path of the second connection portion 46: as can be seen in FIG. 7, the lower stiffness of this connection then gives the exhaust cone 22 greater freedom of movement. Furthermore, this reduction of stiffness changes the dynamic response of the system, and in particular causes the modes of vibration of the propulsive assembly to be shifted and attenuated.

(28) FIG. 8 shows this phenomenon by means of a graph plotting the moment that acts on the flanges forming the interface between the inter-turbine casing 32 and the LP turbine casing 17 as a function of the engine speed in the event of a limiting unbalance in the fan 12 (where the value of this limiting unbalance is defined by certification regulations).

(29) The curve in dashed line 61 corresponds to the situation in which the exhaust cone 22 remains rigidly fastened to the rear turbine casing 33, i.e. in the absence of the releasable fastening provided by the break bolts 55. In contrast, the curve in continuous line 62 corresponds to the situation after the break bolts 55 have broken, and thus after the exhaust cone 22 has been decoupled relative to the rigid force path 45 of the rear turbine casing 33. The vertical line 69 defines the upper limit of the normal operating range of the engine. Certification tests need to be carried out at such a limiting speed.

(30) It can thus be seen in FIG. 8 that in the absence of decoupling of the present invention, the moment at the interface between the inter-turbine casing 32 and the LP turbine casing 17 presents a large peak 61a in the operating range of the engine. In contrast, after decoupling, this mode 62a is offset to a lower frequency, and above all it presents much lower magnitude. There can also be seen a significant drop B1 of about 30% in the maximum moment that occurs in the operating range of the engine, thus making it possible to lighten the structures of the turbine engine, and in particular to reduce the thicknesses of the casings, to reduce the number of fastener bolts used, or indeed to use different materials.

(31) In the present disclosure, it is the interface between the rear turbine casing 33 and the exhaust cone 22 that is described in detail by way of example. Nevertheless, it should be observed that significant improvements can likewise be obtained by making use of such a configuration between other pairs of portions of the casing of the turbine engine and/or of the nacelle. Thus, by way of example, it is possible to install an analogous configuration at the interface between the outer shroud 33b of the rear turbine casing 33 and the nozzle 23, or indeed between the air intake duct 21 and the fan casing 31, to mention only those two examples.

(32) In any event, the embodiments described in the present disclosure are given by way of non-limiting illustration, and in the light of this disclosure, a person skilled in the art can easily modify these embodiments or envisage others while remaining within the ambit of the invention.

(33) Furthermore, the various characteristics of these embodiments may be used singly or they may be combined with one another. When they are combined, the characteristics may be combined as described above or differentially, the invention not being limited to the specific combinations described in the present disclosure. In particular, unless specified to the contrary, a characteristic described with reference to any embodiment may be applied in analogous manner to some other embodiment.