Turbojet engine comprising an afterburner chamber having a diaphragm cooperating with a jacket to condition its passage cross-section to the axial expansion of the jacket

Abstract

The invention relates to a turbojet engine comprising a compressor and an afterburner chamber extending around a longitudinal axis, comprising: – a casing (8); – an inner sleeve (13) defining with this casing (8) an annular space, this sleeve (13) having its upstream end rigidly fixed to the casing, the downstream end of the sleeve (13) being movable in longitudinal translation under the effect of thermal expansions; – a diaphragm (31) in the shape of a crown extending in the annular space (14), comprising an upstream part (33) generally frustoconical with a cross-section decreasing upstream and having an upstream edge (39) defining with the sleeve (13) flow passage ports;– the sleeve (13) having protuberances (48) projecting from its outer face located longitudinally opposite and downstream of the ports so as to be closer to the ports the less the sleeve (13) is dilated, in order to close the ports more and more as the sleeve (13) is dilated. Figure for the abbreviation: Figure 6;

Description

Title of the invention: Turbojet comprising an afterburner chamber having a diaphragm cooperating with a liner to condition its passage section to the axial expansion of the liner Technical field

[0001] The invention applies to the field of military aeronautical propulsion, and more particularly to turbojets equipped with a convergent or convergent and divergent nozzle. STATE OF THE PRIOR ART

[0002] A turbojet engine has an inlet sleeve through which air is admitted to pass through a low pressure compressor before being split into a central primary flow and a secondary flow surrounding the primary flow.

[0003] The primary flow is then compressed in a high-pressure compressor before reaching a combustion chamber, after which it is expanded through a high-pressure turbine and a low-pressure turbine. The secondary flow is propelled directly backward by the low-pressure compressor in a vein delimited externally by an external casing.

[0004] The primary and secondary flows then pass through an exhaust casing located downstream of the low pressure turbine.

[0005] In the case of a military engine, an afterburner chamber located downstream of the exhaust casing receives the secondary flow joining the primary flow. In the so-called afterburner regime, fuel is injected into this chamber to generate additional combustion, whereas in a so-called dry regime no fuel is injected into the afterburner chamber.

[0006] To limit heating, the wall of the casing is lined at the level of the post-combustion chamber by an internal jacket delimiting with it an annular intermediate space. Part of the secondary flow is directed into this annular space to ventilate it in order to limit heating of the casing.

[0007] The afterburner chamber is extended by a nozzle with which it constitutes the afterbody of the turbojet. This nozzle comprises a convergent portion which can be extended by a divergent portion, the junction of the convergent portion with the divergent portion constituting the neck of this nozzle.

[0008] In operation, the air having passed through the annular space is evacuated at an outlet of this space located longitudinally at the downstream edge of the internal jacket, the air from this outlet being used to cool the shutters.

[0009] It may happen that the pressure of the flow circulating in the annular space is too low compared to that of the propulsion flow circulating in the afterburner chamber and in the nozzle.

[0010] In this case, hot gases from the propulsion flow can rise into the outlet of the annular space and circulate there counter-currently, which leads to heating of the downstream part of the casing which is detrimental in particular to the components which surround it.

[0011] In this context, a diaphragm is mounted in the annular space, in the vicinity of its outlet, to reduce the passage section at the outlet of this annular space, in order to increase the pressure in it to reduce the risks of hot gases rising.

[0012] A significant reduction in the passage section is necessary for certain engine speeds, but it is not necessary at other speeds, whereas the diaphragm is dimensioned to continually generate a significant reduction in the section. This significant reduction in the passage section thus penalizes the flow circulating in the annular space, and thereby its cooling and that of the flaps for all engine speeds.

[0013] The aim of the invention is to provide a solution to overcome this drawback. Statement of the invention

[0014] To this end, the invention relates to a turbojet engine afterbody comprising an afterburner chamber extending around a longitudinal axis and in which a gas flow circulates from upstream to downstream when this turbojet engine is in service, this afterburner chamber comprising:

[0015] - a casing;

[0016] - an internal sleeve delimiting with this casing an annular space, this sleeve having its upstream end rigidly secured to the casing, the downstream end of the liner being movable in translation along the longitudinal axis under the effect of thermal expansion;

[0017] - a crown-shaped diaphragm extending inside the annular space, this diaphragm comprising an upstream part having a generally truncated cone shape with a section decreasing towards the upstream, this upstream part having an upstream edge delimiting with the jacket flow passage openings;

[0018] - the shirt comprising protuberances protruding from its external face, these protuberances berances being located longitudinally opposite the lights and downstream of these lights to be all the closer to the lights as the jacket is slightly expanded, so that these protuberances block the lights all the more than the jacket is weakly dilated.

[0019] The opening of the ports is thus maximum when the thermal expansion of the jacket is significant, and it is minimum when the thermal expansion of the jacket is low or zero. The invention thus makes it possible to automatically adjust the opening of the ports to the temperature prevailing in the combustion chamber, that is to say to adjust the cooling flow rate to the thermal situation of the postcombustion chamber.

[0020] The invention also relates to a turbojet afterbody thus defined, in which the diaphragm extends into the downstream end of the annular space.

[0021] The invention also relates to a turbojet afterbody thus defined, in which the upstream part of the diaphragm is extended by a cylindrical downstream fixing part.

[0022] The invention also relates to a turbojet afterbody thus defined, in which the protuberances are positioned longitudinally on the liner to be radially opposite the generally conical upstream part of the diaphragm when the thermal expansion of the liner is zero, and to be located downstream of this upstream part when the thermal expansion of the liner is maximum.

[0023] The invention also relates to a turbojet afterbody thus defined, in which the diaphragm delimits an alternation of ports and radial legs, and in which the jacket comprises a protuberance opposite each port.

[0024] The invention also relates to a turbojet afterbody thus defined, comprising protuberances of different dimensions to block certain ports differently.

[0025] The invention also relates to a turbojet afterbody thus defined, in which the protuberances are elements attached to the external face of the liner.

[0026] The invention also relates to a turbojet afterbody thus defined, in which the protuberances are integrated into the jacket.

[0027] The invention also relates to a turbojet afterbody thus defined, in which the part of the liner which is in contact with the radial legs is cylindrical.

[0028] The invention also relates to a turbojet comprising an afterbody thus defined. Brief description of the drawings

[0029] [Fig.l] is a schematic view in longitudinal section of the turbojet according to the invention;

[0030] [Fig.2] is a schematic view in partial longitudinal section showing the afterbody of the turbojet according to the invention in dry operation;

[0031] [Fig. 3] is a schematic cross-sectional view of a shirt fastener;

[0032] [Fig.4] is a schematic cross-sectional view of a bridge;

[0033] [Fig.5] is a schematic view in partial longitudinal section showing the afterbody of the turbojet according to the invention in post-combustion mode;

[0034] [Fig.6] is a cross-sectional view of the diaphragm shown alone at a radial lug when the engine is in dry operation;

[0035] [Fig.7] is a cross-sectional view of the diaphragm shown alone at a radial lug when the engine is in afterburner mode;

[0036] [Fig.8] is a top view of a portion of the diaphragm shown alone;

[0037] [Fig.9] is a general perspective view of the diaphragm shown alone;

[0038] [Fig. 10] is a cross-sectional view along a cutting plane located upstream of the diaphragm showing a radial leg of the diaphragm;

[0039] [Fig. 11] is a cross-sectional view along a cutting plane located upstream of the diaphragm showing three radial legs of the diaphragm;

[0040] [Fig. 12] is a perspective view of the rear part of the afterburner chamber with its diaphragm. DETAILED DESCRIPTION OF THE INVENTION

[0041] In [Fig.l], the turbojet engine 1 according to the invention comprises an inlet sleeve through which the air is admitted to pass through a low-pressure compressor 2 before being split into a central primary flow F1 and a secondary flow F2 surrounding the primary flow. These two flows circulate in the turbojet engine parallel to its longitudinal axis AX which corresponds to its axis of revolution, from its upstream AM to its downstream AV.

[0042] The primary flow is then compressed in a high pressure compressor 3 before reaching a combustion chamber 4, after which it is expanded through a high pressure turbine 6 and a low pressure turbine 7.

[0043] The secondary flow is propelled directly towards the rear by the low pressure compressor 2 in a vein delimited externally by an external casing 8.

[0044] The primary flow F1 and secondary flow F2 then pass through an exhaust casing 9 located downstream of the low-pressure turbine, which is extended by a confluence plate 11 extending downstream thereof.

[0045] This exhaust casing comprises a hub surrounded by a casing secured to each other by radial arms, the casing being carried by a structure of revolution integrated into the external casing 8, by means of radial yokes. The primary flow flows between the hub and the casing while the secondary flow flows around the casing.

[0046] This turbojet is equipped downstream of the casing 9 with a post-combustion chamber 12, in which the secondary flow F2 joins the primary flow F1, constituting a flow of propulsion noted Fp, and into which fuel is injected to generate additional combustion.

[0047] To limit heating, the wall of the casing 8 is lined at the level of the chamber 12 by an internal jacket 13 delimiting with it an annular space 14. A part of the secondary flow is directed into this annular space to ventilate it in order to limit heating of the casing 8.

[0048] This jacket 13 provides thermal protection for the casing 8: by lowering the temperature of this jacket, its thermal radiation towards the wall of the casing 8 is reduced.

[0049] Additionally, air circulation is established around the engine, i.e. between the engine and structural elements of the aircraft which carry it, to contribute to its cooling, this additional flow, called tunnel air, being noted Ft.

[0050] As visible in [Fig. 1], the afterburner chamber is extended by a nozzle 16 with which it constitutes the rear body of the turbojet. This nozzle comprises a convergent portion 17 which can be extended by a divergent portion 18, the portion located at the downstream end of the convergent portion 17 constituting the neck 20 of the nozzle 16.

[0051] The converging portion 17 is delimited by a series of converging flaps 19 distributed around the axis AX and extending in an oblique direction relative to this axis AX to jointly delimit a conical shape. Each flap 19 has an upstream end secured to a downstream edge of the external casing 8 by an articulation 21 with an orthoradial axis relative to the axis AX.

[0052] Similarly, when the nozzle comprises a diverging portion 18, this diverging portion is delimited by a series of diverging flaps 22 distributed around the axis AX of oblique orientation relative to the axis AX to constitute a conical shape. Each diverging flap 22 then has its upstream end secured to a downstream edge of a corresponding converging flap 19 by an articulation 23 of orthoradial axis relative to AX.

[0053] These flaps are actuated by hydraulic cylinders to modify their inclinations, so as to increase or decrease the section of the nozzle neck, being controlled by a control system not shown.

[0054] This engine can thus operate at a so-called dry regime corresponding to figures 2 and 6, in which the afterburner chamber is not supplied with fuel, and in which the nozzle throat 20 has a minimum section. It can also operate at a so-called afterburner regime corresponding to figures 5 and 7 in which the supply of fuel to the afterburner chamber makes it possible to temporarily increase its thrust, and in which the throat 20 has a maximum section.

[0055] As visible in [Fig.2], the air flow Fs circulating in the annular space 14 is evacuated through holes 24 formed in the jacket 13 and through the outlet 26 of this annular space which corresponds to the crown delimited internally by the downstream edge of the jacket 13 and externally by that of a thermal protection ferrule 27.

[0056] This air passes through the holes 24 following oblique trajectories to cool the jacket 13, and it passes through the outlet 26 following a longitudinal trajectory to leave the annular space 14 before ensuring cooling of the flaps.

[0057] The outlet 26 exposes the rear part of the casing 8 to the thermal radiation of the flow leaving the post-combustion chamber, so that the thermal protection shroud 27 is provided to protect this rear part of the casing 8.

[0058] The ferrule 27 comprises a cylindrical body 28 of diameter smaller than that of the casing 8, extended upstream by a rim 29 extending radially outwards to join the casing 8 by which this ferrule 27 is carried. It thus constitutes a reduction in the passage section of the annular space 14 at its outlet 26.

[0059] The annular space 14 is equipped with a diaphragm 31 in the form of a sheet metal ring delimiting openings which further reduce the passage section. This diaphragm 31 makes it possible to increase the pressure in the annular space 14 to reduce the risk of hot gases rising through the outlet 26.

[0060] The diaphragm 31 comprises a generally conical upstream portion 33 provided with undulations extending radially inwards to be in contact with the sleeve 13, extended by a downstream fixing portion 34. This downstream portion 34 is cylindrical and has a diameter corresponding to that of the ferrule 27, and it is fixed to this ferrule 27 for example by riveting or by bolting.

[0061] In post-combustion mode, the temperature increases in the post-combustion chamber, so that the liner 13 expands essentially longitudinally, and to a lesser extent radially. This liner 13 is rigidly secured to the casing 8 at its upstream end by fasteners 36, and it is held radially in its longitudinally median region and in its downstream region by bridges 37 carried by the casing 8. These bridges 37 thus ensure the radial retention of the liner while constituting a sliding connection allowing its longitudinal expansion.

[0062] These fasteners 36 are distributed circumferentially in the inlet of the annular space 14, that is to say at its upstream end. Each fastener 36 has its radially external end fixed to the casing 8 and its radially internal end fixed to the sleeve 13 as visible in [Fig. 3], for example by riveting or by bolting.

[0063] The bridges 37 are also distributed circumferentially inside the space 14, on the one hand at its median region, and on the other hand near its downstream end, also being fixed to the casing 8 by riveting or bolting. They thus ensure the radial retention of the median part and the downstream part of the liner 13.

[0064] As visible in [Fig.6] to 9, the upstream part 33 of the diaphragm 31, which has a generally in the shape of a truncated cone whose diameter decreases towards the upstream AM, comprises undulations extending radially inwards, that is to say towards the axis AX. These undulations are in contact with the external face of the jacket 13 to constitute the radial tabs 38 of this diaphragm 31.

[0065] Each radial tab 38 is thus constituted by a portion of wall 42 of the upstream part 33 which is curved towards the axis AX, the downstream end of this portion of wall 42 connecting to the cylindrical downstream part 34, while the upstream end of this portion of wall 42 is terminated by an undulating portion of the upstream edge 39.

[0066] Each portion of wall 42 which is contiguous to the upstream edge 39 constitutes a key 44 in contact with the sleeve 13, as visible in Figures 6 and 7. When this sleeve 13 elongates, it slides longitudinally on these keys 44 while remaining radially in contact with them, which corresponds to the situation in Figures 5 and 7. In this context, and as visible in Figures 3, 4, 6 and 7, the sleeve 13 advantageously has a cylindrical shape devoid of undulations in its downstream part by which it is in contact with the radial tabs 38 of the diaphragm.

[0067] The radial legs 38 of each pair of circumferentially consecutive radial legs are connected to each other by a panel 43 of the upstream part 33, each panel 43 having an oblique orientation approaching the axis AX towards the upstream, and a cone portion shape.

[0068] As visible in Figures 10 and 11, each region extending between two consecutive radial tabs 38 thus constitutes a light 46 allowing the passage of air through the diaphragm. Each light 46 is delimited externally by the upstream edge 39 of a panel 43, internally by the external face of the jacket 13, and laterally by the upstream edges of the two consecutive radial tabs 38 which it separates.

[0069] The diaphragm 31 thus defines an alternation of radial legs 38 and slots 46 over its entire circumference, the radial legs 38 having an angular extent around the axis AX significantly less than that of the slots 46.

[0070] Additionally, each curved wall portion 42 includes a window 47 which allows the air circulating in the annular space 14 to pass through the radial tab 38 which it constitutes.

[0071] The external diameter of the upstream edge 39 (which is corrugated) is less than the diameter of the casing 8, this external diameter corresponding for example to the average of the internal diameter of the shell 27 and the external diameter of the jacket 13. The passage openings which the diaphragm 31 delimits, that is to say the ports 46 and the windows 47, thus define a passage section whose area is of the order of half that of the passage section of the annular space 14 at the level of the shell 27.

[0072] The sizing of the area of ​​the passage openings is a compromise between a need for sufficient flow for cooling, and the reduction of the risk of backflow. of hot gases. It conditions the pressure difference between the annular space and the post-combustion chamber, and it is chosen so that this difference remains below an admissible maximum conditioning the risk of hot gases rising.

[0073] The presence of the windows 47 ensures a distribution of the air flow passing through the diaphragm 31, which is uniform over its entire circumference.

[0074] According to the invention, the external face of the downstream end of the jacket 13 carries protrusions 48 each located longitudinally opposite a slot. These protrusions 48 are located downstream of the slots 46 and they protrude radially beyond the external face of the jacket 13 by a height in the radial direction which is slightly less than that of the slots 46.

[0075] When the temperature is low, the jacket 13 undergoes low thermal expansion, so that the protrusions 48 are longitudinally brought closer to the openings 46 to close them. The protrusions 48 are then opposite the panels 43 to delimit with them a small passage section, which is represented schematically by the circle marked C in [Fig.6].

[0076] When the temperature is high, the jacket 13 undergoes significant thermal expansion in the axial direction, so that the protrusions 48 move away from the ports 46 so as not to block them. The protrusions 48 are then located radially opposite the cylindrical downstream part 34 to delimit with this part a large passage section, which is represented schematically by the circle C' in [Fig.7].

[0077] Due to the inclined orientation of the panels 43 forming the upstream part 33 of the diaphragm 31, the ports 46 visible in [Fig. 1 1] are all the more blocked the lower the expansion of the jacket. In other words, the ports are all the more blocked the lower the temperature in the post-combustion chamber.

[0078] The invention thus makes it possible to ensure spontaneous control of the passage section through the annular space 14 which increases this passage section when the temperature increases, that is to say when the need for cooling becomes greater, and which decreases it when the temperature is low, that is to say when the need for cooling is low.

[0079] Generally speaking, the thermal expansion of the jacket 13 is low in dry mode, and it is maximum in post-combustion mode, but the invention makes it possible to increase the passage section, and therefore the cooling, including in the event of an increase in temperature in dry mode, and to reduce it in post-combustion mode if the thermal situation requires it.

[0080] The protrusions 48 which are visible in Figures 6, 7, 10 and 11 can be either elements attached to the external face of the jacket 13 at its downstream end. These protrusions can also be integrated into the jacket 13 by being by obtained by deformation by stamping or similar of the sheet metal of this jacket, at its downstream end.

[0081] Each protuberance 48 has a bead shape extending around the circumference of the jacket 13 over an angular extent corresponding to the angular extent of the slot 46 located longitudinally opposite this bead.

[0082] In practice, as certain angular sectors of the engine may require greater cooling than others, some of the protrusions 48 may be provided with larger (or smaller) dimensions than the others, so as to create more (or less) obstruction for the corresponding ports. Thus, the thickness (in the radial direction) of each protrusion may be adjusted to the cooling requirement of the corresponding port.

[0083] As visible in [Fig.8], each panel 43 has a slot 49 extending longitudinally from the upstream edge 39 to a region which adjoins the cylindrical downstream part 34. Thanks to these slots 49, the radial tabs 38 are independent of each other, each tab thus having a certain radial flexibility.

[0084] The diaphragm 31 which is formed from a single piece of sheet metal in the example of the figures can also be made up of several distinct diaphragm sectors, each comprising one or more radial tabs 38. These sectors are then fixed independently of each other to the ferrule 27, by riveting or by bolting, to jointly constitute an assembly having essentially the same geometric shape and the same characteristics as the diaphragm 31 of the figures.

[0085] As illustrated in [Fig. 12], a flange 51 can advantageously be attached to the downstream end of the afterburner chamber 12 by being fixed to the downstream edge of the casing 8 to close the annular space delimited by the ferrule 27 and the casing 8. This flange 51 then advantageously comprises a cylindrical internal rim 52 extending upstream and running along the internal face of the downstream part 34 of the diaphragm 31, the diaphragm 31 can then be fixed by its downstream part 34 to this cylindrical internal rim 52, for example by bolting or by riveting.

Claims

Claims

1. Turbojet engine afterbody comprising an afterburner chamber (12) extending around a longitudinal axis (AX) and in which a gas flow circulates from upstream (AM) to downstream (AV) when this turbojet engine is in operation, this afterburner chamber (12) comprising: - a casing (8); - an internal liner (13) delimiting with this casing (8) an annular space (14), this liner (13) having its upstream end rigidly secured to the casing (8), the downstream end of the liner (13) being movable in translation along the longitudinal axis (AX) under the effect of thermal expansions; - a crown-shaped diaphragm (31) extending inside the annular space (14), this diaphragm (31) comprising an upstream part (33) having a generally truncated cone shape with a section decreasing towards the upstream, this upstream part (33) having an upstream edge (39) delimiting with the jacket (13) flow passage openings (46);- the jacket (13) comprising protuberances (48) protruding from its external face, these protuberances (48) being located longitudinally opposite the slots (46) and downstream of these slots (46) to be all the closer to the slots (46) the more the jacket (13) is slightly expanded, so that these protuberances (48) block the slots (46) all the more the jacket (13) is slightly expanded.;

2. A turbojet afterbody according to claim 1, wherein the diaphragm (31) extends into the downstream end of the annular space (14).

3. Turbojet afterbody according to one of the preceding claims, in which the upstream part (33) of the diaphragm (31) is extended by a cylindrical downstream fixing part (34).

4. Turbojet afterbody according to one of the preceding claims, in which the protuberances (48) are positioned longitudinally on the liner (13) to be radially opposite the generally conical upstream part (33) of the diaphragm (31) when the thermal expansion of the liner (13) is zero, and to be located downstream of this upstream part (33) when the thermal expansion of the liner (13) is maximum.

5. Turbojet afterbody according to one of the preceding claims, in which the diaphragm (31) delimits an alternation of ports (46) and radial legs (38), and in which the jacket (13) comprises a protuberance (48) opposite each port (46).

6. Turbojet afterbody according to claim 5, comprising protrusions (48) of different dimensions to differently close certain ports (46).

7. Turbojet engine afterbody according to claim 5, in which the protrusions (48) are elements attached to the external face of the liner (13).

8. A turbojet engine afterbody according to claim 5, wherein the protrusions (48) are integrated into the jacket (13).

9. A turbojet afterbody according to one of the preceding claims, in which the part of the liner (13) which is in contact with the radial legs (38) is cylindrical.

10. Turbojet comprising an afterbody according to one of the preceding claims.

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