Patent Application
20070251211
The invention is now described in greater detail with reference to the appended drawings in which:
The aircraft 1 represented in the figure is a nonlimiting example. It has a nose 2, two wings 3 and 4 and is propelled by one or two turbojets that are not visible. It is shaped so as to have the smallest possible RCS and IRS. Its rear portion in particular has no vertical stabilizer and terminates in a point 5 with an appropriate apex angle, for example 40°, to reject the radar waves to infinity. The exhaust assembly 10 participates in this requirement by being bifid. It distributes the main flow exiting the channel 12 at the entrance into two flows in two symmetrical channels 12A and 12B that terminate in two half-nozzles 14 and 16 of rectangular section. The channels 12, 12A and 12B have a shape suitable for ensuring the separation of the flow into two flows but also the transition from a circular or substantially circular section cylindrical shape to a rectangular section shape. Where appropriate, the channels include an additional elbow for masking the turbine. As may be seen in the figures, this masking is already at least partly provided by the separation between the half-nozzles 14 and 16.
According to the present application, the shape of the exhaust assembly is enhanced so as to ensure the masking of the turbine block irrespective of the position of a rear observer. As may be seen in
With reference to
This assembly comprises a transition element 21 with a cylindrical upstream portion 21A, of circular or other cross section. The transition element emerges in a first duct element 22 and a second duct element 23, these two being parallel. The two duct elements each terminate in a respective half-nozzle 24 and 26.
The upstream portion 21A communicates directly with the exit of the gas generator (not shown), such as a turbine of a gas turbine engine. As may be seen in
The shape of the transition element progressively changes toward the downstream so as to adopt the contour 21B consisting of two ellipses, which partially overlap, as may be seen in the plane of section VIIb-VIIb. The two ellipses, which are identical, here have a vertical major axis. On going toward the downstream, they progressively move apart until adopting the outline at 21C of two ellipses, as may be seen in the plane of section VIIc-VIIc. The separation takes place in the plane of section VIId-VIId.
On moving axially downstream, the two ducts each form an elbow. They progressively move apart and deflect the gas flows radially toward the outside until reaching a maximum separation at 22M and 23M where the flows become axial. Downstream, they converge on each other, deflecting the flows radially toward the axis until reaching 22N and 23N where they are returned to the axis. At this point, the separation between the two ducts is still sufficient to correspond substantially to that of the diameter of the entrance plane 21A. Each duct terminates in a half-nozzle, 24 and 26 respectively, which diverges downstream of the throat that lies in the plane 22N-23N. Here they have a rectangular cross section, but other shapes are possible. The shape of the cross sections of the ducts 22 and 23 progressively changes until they have the shape of the half-nozzles. The areas are determined according to the requirements of the fluid dynamics.
Preferably, the assembly has at least one of the following dimensional relationships:
Lelbow/Lchannel is between 0.5 and 0.7;
Lint/Lext≧½;
Lint/Lchannel close to ⅓;
Lext/Lchannel≧½;
Lseparation/Lchannel≦0.3,
where:
Lelbow is the length measured axially from the entrance plane 21A to the point where the elbow is at its maximum lateral deviation from the axis;
Lchannel is the length measured axially from the entrance plane 21A to the throat of the half-nozzles;
Lint is the width of the elbow measured transversely from the engine axis to the internal wall of the duct, at the point where the duct element is at its maximum departure;
Lext is the width of the elbow measured from the axis to the external wall of the duct, at the point where the duct element is at its maximum departure; and
Lseparation is the length measured along the engine axis from the entrance plane 21A to the plane of section VIId-VIId.
As illustrated by the straight lines D1 and D2, such a geometry allows effective masking of the hot zones of the engine and in particular the zones of the transition elements through which the gas flow passes. These straight lines constitute the limits of visibility of these zones.
The means of yaw guidance of the aircraft will now be described with reference to
The assembly is preferably optimized to provide, in the cases with no injection and no vectorization, a minimum transverse thrust of each half-nozzle. Specifically, the latter results in a loss of axial thrust that must be reduced to a minimum. The overall lateral thrust remains zero because of the symmetry of the system.
According to a feature of the invention, to provide the guidance of the aircraft 1 without a tail unit, control means are provided by which action is taken on the two flows. These control means may be mechanical or fluidic.
The convergent-divergent nozzle, for example 24, comprises the throat 24C and downstream the two divergent walls 24DI and 24DE. Here the nozzle comprises a fluid injector 28 placed on a wall at the throat and a fluid injector 29 situated on the wall 24DI of the divergence. The injector is preferably situated close to the end of the divergence.
In a symmetrical manner, the half-nozzle 26 is fitted with a fluid injector 28 at the throat 26C and a fluid injector 29 on the wall of divergence 26DI.
The injectors 28 and 29 are advantageously supplied with air tapped from the turbojet compressor that supplies the main flow, as appropriate.
Operation is as follows.
It is observed however that a sudden obstruction of the nozzle would instantaneously create an increase in pressure in the channel and a risk of pumping the compressor. According to a preferred operating mode, a nominal permanent injection is created. This is done at equal delivery rate tapped off in such a way that the generator does not undergo a sudden variation during the mission while regulating the nozzle at total equal effective section at the throat. The thermodynamic cycle of the engine is directly optimized under this constraint of constant tapping. In this manner, the system of regulating the tapped air operates continuously and does not undergo any transitional startup phase.
Therefore this operating mode in accordance with the invention provides, with a low impact on the performance of the engine, a vectored thrust that makes it possible to compensate for the absence of cell tail unit, particularly for cruising or slow transitional speeds.
The operation of the injection device situated in the divergence of the nozzles 24 and 26 is now described with reference to
The injectors 29, in this embodiment, are preferably placed at the end of the long wall of divergence. By injecting a fluid into the nozzle 24, the direction of which is represented by the arrow 29/24, a deviation of the thrust vector produced by the nozzle and shown by the arrow F′2 is induced. The thrust F′1 provided by the half-nozzle 26 remains axial since nothing disrupts its direction. This results in the creation of a yaw moment relative to the center of gravity of the aircraft. This operating mode provides a substantial vectored thrust in order to control the aircraft, to the detriment however of the performance of the generator. This deterioration is however controlled.
One embodiment of the invention has been described. However, many variants are possible without departing from the context of the invention. For example, a channel has been shown supplied by a single gas generator. In the case of a twin-engined aircraft, the two half-flows of exhaust are generated by two distinct engines whose regulation is synchronized. Preferably, only the injectors in the divergence are used.
Variants of the arrangement and operation of the control means comprise the presence of a single control means. It is possible to operate it at the same time as the other means or separately.
According to an embodiment not shown, the nozzles may be of the fluid type with ejector, that is to say a secondary flow emerging in or downstream of the main channel.
The control means according to the invention may be combined partly with mechanical means of orienting the flows.
| Number | Date | Country | Kind |
|---|---|---|---|
| 0651542 | Apr 2006 | FR | national |
