GB2285097A

GB2285097A - Turbojet afterburn unit

Info

Publication number: GB2285097A
Application number: GB9422706A
Authority: GB
Inventors: Frederic Beule; Michel Andre Albert Desaulty; Jacques Dufau; Elisabeth Vilfeu
Original assignee: Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA; SNECMA SAS
Current assignee: Safran Aircraft Engines SAS
Priority date: 1993-11-24
Filing date: 1994-11-10
Publication date: 1995-06-28
Anticipated expiration: 2014-11-10
Also published as: FR2712962B1; FR2712962A1; GB9422706D0; GB2285097B

Abstract

A turbojet engine is provided with an afterburn unit having non-deformable flame catcher arms 10, in which each flame catcher arm 10 has a streamlined cross-section which is sharp at its upstream end 22 and convexly curved at its downstream end 23, and the final stage of vanes 7 of the turbine blading each has an upstream part 24 which is fixed relative to the engine structure and a downstream part 25 which is pivotable between a first configuration, in which it extends substantially in a radial plane 28 containing the engine axis, and a second configuration, in which it is oriented in a direction K which is oblique to the said radial plane 28. The arrangement provides for turbulence production to facilitate mixing of gases during afterburner operation, and a low-drag configuration when afterburner operation is not required. <IMAGE>

Description

TURBOJET AFTERBURN UNIT Multiflow turbojet engines used for military aircraft have an afterburn of "reheat" chamber serving as a second combustion chamber. This enables further heat energy to be injected into the gases between the time they leave the turbine and the time they are ejected through the nozzle, and helps to increase the gas ejection rate and hence engine thrust.

The primary flow gases leaving the main combustion chamber pass through the blading of the turbine, enter the afterburn chamber and are ejected through the nozzle. The secondary flow gases serve to cool the afterburn chamber and also supply unburned air thereto.

The afterburn unit comprises a number of fuel injection systems and flame catcher devices, for examples in the form of radial arms. The flame catcher devices serve to produce turbulence in the afterburn chamber in order to increase the dwell time of the fuel/air mixture therein.

Combustion is therefore stabilised and higher outputs are obtained.

The extra thrust provided by afterburn is useful to the pilot both for takeoff and for aerial combat. For the rest of the time the engine is operated "dry" - i.e. in a state in which fuel is not supplied to the afterburn chamber and in which there is therefore no combustion therein.

The flame catcher arms are essential for afterburn operation, but are useless and may even cause pressure losses in "dry" operation. The usual flame catcher arms, which are usually radial, may have the shape of a V with the tip pointing upstream, the arms of the V enhancing turbulence during afterburn operation. Means for retracting the arms have been proposed for dry operation but are complex, bulky and heavy. Also, since they operate in a very hot zone they are subject to expansion, which may lead to some arms jamming and thus impairing both afterburn and dry operation.

This invention aims to provide an afterburn unit with a simple flame-stabilising system which offers very low drag during "dry" operation and which can produce turbulence aerodynamically in the afterburn chamber during afterburn operation.

To this end, according to the invention there is provided a bypass turbojet engine having an afterburn unit, comprising an outer first wall and an inner second wall which are coaxially disposed about a common axis of revolution and which define between them an annular outer duct for a secondary flow of combustion-supporting agent, an inner exhaust duct located inwardly of the second wall and separated thereby from the outer duct, the exhaust duct being disposed downstream from the turbine blading of the engine for conducting the gases issuing therefrom, an afterburn chamber which is bounded by a third wall disposed axially downstream of the first and second walls and which forms a downstream axial extension of the outer and inner ducts, and flame catcher arms which are substantially non-deformable and are substantially fixed with respect to the second wall, the flame catcher arms being symmetrical with respect to radial planes containing the axis of revolution and each having a streamlined cross-section which is sharp at its upstream end and convexly curved at its downstream end, the said turbine blading comprising a final downstream stage of vanes of which at least some have two parts, namely an upstream first part which is fixed with respect to the second wall and a downstream second part which is mounted to pivot relative to the second wall about an axis contained in a radial plane passing through the axis of revolution and which is movable between a first configuration, in which it extends substantially in a radial plane containing the axis of revolution, and a second configuration, in which it is oriented obliquely with respect to the said radial plane.

The flame catcher arms preferably extend radially relative to the axis of revolution, and the pivot axis of the second part of each two-part vane is preferably substantially perpendicular to the axis.

Each flame catcher arm is preferably ventilated to achieve the cooling needed for satisfactory thermal behaviour, particularly during afterburn operation. This ventilation may be achieved by internal channels which form labyrinths and which open out in that part of the outer surface which is most exposed to heat radiation, or simply by the flame catcher arm being hollow and bounded by a wall through which a plurality of apertures extend, the apertures being oriented transversely relative to the axis of revolution or having an upstream-to-downstream orientation, the interior of the arm also communicating, via an inlet opening, with the annular outer duct. The radially outer end of the wall of each flame catcher arm may be fixed to the second wall and preferably also defines the inlet opening.

Surface fuel ducts may be provided in the extrados face of the first part of at least some of the two-part vanes, these fuel ducts having apertures which are oriented transversely with respect to the axis of revolution and which form fuel injection orifices.

Alternatively or in addition, substantially radially extending fuel ducts may be interposed between the trailing edges of the second parts of the two-part vanes and the flame catcher arms, and have apertures which are oriented substantially parallel to the axis of revolution and which form fuel injection orifices.

The main advantage of the arrangement in accordance with the invention is that the drag caused by the flame catcher arms disposed in the afterburn chamber is very reduced by the streamlining of these arms, and that the blading which produces turbulence during afterburn operation is located outside the afterburn chamber in a zone where thermal stresses are less than in the afterburn zone and, therefore, in a zone where there is less risk of deformation and jamming of moving members and their actuating system.

Various embodiments of the invention will now be described, by way of example only, with reference to the accompanying drawings, in which: Figure 1 is an axial section through part of a turbojet engine having an afterburn unit in accordance with the invention; Figure 2 is a transverse section on the line II-II in Figure 1; Figure 3 is a partial "developed" view of part of the engine shown in Figure 1; Figure 4 is a cross-section through a flame catcher arm of the engine shown in Figure 1; and, Figures 5 and 6 are views similar to that of Figure 3 but showing two variants of the afterburn unit.

The turbojet engine shown in Figure 1 comprises an outer first wall 2 and an inner second wall 8 disposed coaxially about a general axis 1 of revolution and extending from upstream A to downstream B substantially as far as a transverse plane T perpendicular to the axis 1, the walls 2 and 8 defining an annular duct 12 between them.

An annular combustion chamber 3 is disposed coaxially around the axis 1, and a stage 4 of turbine inlet guide vanes are secured to the inner second wall 8 downstream from the combustion chamber. A stage 5 of mobile turbine blades are secured to the body of a rotor 6 rotatable about the axis 1 relative to the walls 2 and 8, and a final stage 7 of turbine outlet guide vanes are mounted on the wall 8.

An outer third wall 9, which is an extension of the first wall 2 in the downstream direction, defines an afterburn chamber 18.

A plurality of hollow radial arms 10, each defined by a wall 11, are secured at their radially outer end llA to the downstream end of the inner second wall 8, the end llA of each arm bounding an opening in the downstream end of the second wall 8 which provides a communication between the interior of the arm and the duct 12 bounded by the first and second walls 2 and 8, and the inner other end llB of each arm being closed. The rotor 6 is bounded by a wall 13 which extends to near the ends llB of the arms 10.

A ring 14, concentric with the axis 1, is secured to the downstream end of the inner second wall 8, and has a cross-section in the shape of a V having its tip 15 pointing upstream. An annular fuel duct 16 is disposed inside the V-section of the ring 14 and is formed with a plurality of fuel injection orifices 17 which open in the downstream direction B into the afterburn chamber 18 and are oriented parallel to the axis 1, the duct 16 being connected to a fuel supply line 19.

The annular duct 12 serves as a duct for the flow of secondary combustion-supporting agent and is connected upstream to a source of compressed combustion-supporting agent diagrammatically represented by the arrows F and usually taking the form of an air compressor, the dct 12 opening at its downstream end into the afterburn chamber 18.

In a known manner, the combustion chamber 3 also communicates with the source of compressed combustion-supporting agent, as indicated by the arrows G. The exhaust gases leaving the turbine stages 4, 5, 7 are guided by the vanes of the final stage 7 through the exhaust duct 20 which is defined between the inner second wall 8 and the rotor wall 13 and extends as far as the arms 10, where the duct 20 opens into the afterburn chamber 18.

Each arm 10 extends radially relative to the axis 1, and its wall 11 has a plane of symmetry Pll which contains the axis 1. The wall 11 is streamlined in a direction from upstream to downstream, its cross-section forming a point 22 which is oriented towards the upstream direction A and having a convex rounded surface 23 facing in the downstream direction B. Also, the wall 11 has a plurality of apertures 21 extending through it, these apertures 21 preferably having an upstream-to-downstream orientation as shown in Figure 4.

At least some of the vanes of the final turbine guide stage 7 have an upstream part 24 secured to the inside surface of the second wall 8, and a downstream part 25 which is mounted to pivot relative to the second wall 8 on a shaft 26 disposed in a radial plane P26. In the embodiment shown all the vanes of the stage 7 are constructed in this way, the pivot shaft 26 of each of them being substantially perpendicular to the axis 1.

Each shaft 26 is also connected to a pivot controller 27 adapted to position the second part 25 either in a first configuration, shown in solid lines in Figure 3 and broken lines in Figure 5, wherein the intrados face 25A is substantially contained in a radial plane 28, or in a second configuration, shown in broken lines in Figure 3 and solid lines in Figure 5, wherein the intrados face 25A forms an angle A25 with the radial plane 28 and therefore with the general gas flow direction H as well.

The stage 7 of vanes is disposed upstream of the radial arms 10.

Referring to Figure 5, the extrados face 24B of the first part 24 of each vane 24-25 of the stage 7 has a fuel duct 29 therein which has a substantially radial extent and which opens into the exhaust gas duct 20 through a plurality of fuel injection orifices 30 all oriented substantially in a transverse plane T30 perpendicular to the axis 1.

Referring to Figure 6, fuel ducts 31 extend substantially radially downstream of the trailing edges 25C of the second parts 25 of the vanes 24-25, and open into the duct 20 through a plurality of fuel injection orifices 32 all oriented substantially parallel to the axis 1. The ducts 31 are positioned axially between the trailing edges 25C of the two part vanes and the flame catcher arms 10.

In the three embodiments shown in Figures 3, 5 and 6 the arrows K denote the direction defined by the intersection of the plane perpendicular to the radial plane 28 and parallel to the axis 1 with the plane of the intrados face 25A of the part 25 of each vane 24-25 when in the second configuration, and reference 33 indicates an eddy in the flow of gases starting from the tip 22 of each radial arm 10.

The assembly formed by the two-part vanes 24-25 and the radial arms 10 form the flame catcher system of the afterburn unit of the engine.

When the second parts 25 of the vanes are placed in their first position (solid lines in Figure 3), the exhaust gases guided by the intrados faces 25A issue from the final guide vane stage 7 parallel to the radial planes 28 - i.e., substantially parallel to the radial symmetry planes Pll of the arms 10. This is the configuration adopted during "dry" operation, i.e. when the afterburner is not in use and there is no need to create eddies in the afterburn chamber 18 for prolonging the dwell time and increasing mixing of the gases, and when instead it is necessary to facilitate a gas flow free from load loss. This is exactly what is obtained, the gases flowing parallel to the direction H and approaching the arms 10 without any angle of incidence, and flowing therealong with minimal friction by virtue of the streamlined shape of the arms 10.

When the second parts 25 of the vanes are placed in their second configuration (solid lines in Figures 5 and 6), the afterburner is operative and eddies 33 are created to enhance the mixing of the gases with the fuel injected through the orifices 17 and 30 or 17 and 32 and hence increase combustion efficiency. The gases leave the vanes 24-25 in the directions K which are inclined to the symmetry planes P11 of the arms 10, and produce the required eddies 33 when reaching the arms.

Furthermore, the ducts 29, 31 provide an injection of fuel which is correctly distributed in the exhaust duct 20 so as to help ensure a thorough and high-efficiency combustion.

It should also be noted that it is essential to avoid moving mechanisms inside or near the afterburn chamber 18 in view of the very high temperatures which can be reached therein, and in this connection the fixing of the arms 10 without moving mechanisms on the inner second wall 8 is an important feature of the invention.

Preferably the arms 10 are hollow, a feature which facilitates their cooling and thus increases their working life and also makes possible a distributed secondary air supply boosting complete combustion of the injected fuel. However, the arms need not be hollow, nor be formed with apertures 21.

In order to limit the angle of turn induced by the steered movable blading during afterburn operation, radial arms with slightly asymmetrical profiles may be used for the sake of achieving a satisfactory compromise between pressure losses during "dry" operation, and pressure losses with flow deviation and flame stabilisation efficiency during afterburn operation.

Ventilation for each flame catcher arm to obtain satisfactory thermal behaviour, particularly during afterburn operation, may be achieved by providing the flame catcher arm with internal ducts which open in its outer surface, and in particular in that part of the outer surface which is most exposed to heat radiation.

In this variant the internal ducts communicate with the annular outer duct and preferably form labyrinths.

As another variant, in the embodiment in which a plurality of apertures 21 extend through the wall of each hollow arm 10, these apertures may be oriented transversely with respect to the axis of revolution.

Claims

1. A bypass turbojet engine having an afterburn unit, comprising an outer first wall and an inner second wall which are coaxially disposed about a common axis of revolution and which define between them an annular outer duct for a secondary flow of combustion-supporting agent, an inner exhaust duct located inwardly of the second wall and separated thereby froin the outer duct, the exhaust: duct being disposed downstream from the turbine blading of the engine for conducting the gases issuing therefrom, an afterburn chamber which is bounded by a third wall disposed axially downstream of the first and second walls and which forms a downstream axial extension of the outer and inner ducts, and flame catcher arms which are substantially non-deformable and are substantially fixed with respect to the second wall, the flame catcher arms being symmetrical with respect to radial planes containing the axis of revolution and each having a streamlined cross-section which is sharp at its upstream end and convexly curved at its downstream end, the said turbine blading comprising a final downstream stage of vanes of which at least some have two parts, namely an upstream first part which is fixed with respect to the second wall and a downstream second part which is mounted to pivot relative to the second wall about an axis contained in a radial plane passing through the axis of revolution and which is movable between a first configuration, in which it extends substantially in a radial plane containing the axis of revolution, and a second configuration, in which it is oriented obliquely with respect to the said radial plane.

2. A turbojet engine according to claim 1, in which the flame catcher arms extend radially with respect to the axis of revolution.

3. A turbojet engine according to claim 1 or claim 2, in which the pivot axis of the second part of each two-part vane is substantially perpendicular to the axis of revolution.

4. A turbojet engine according to any one of claims 1 to 3, in which each flame catcher arm has internal ventilating and cooling ducts which open out in that part of the outer surface which is most exposed to heat radiation.

5. A turbojet engine according to claim 4, in which the internal ducts form labyrinths.

6. A turbojet engine according to any one of claims 1 to 3, in which each flame catcher arm is hollow and is bounded by a wall through which a plurality of apertures extend, the interior of the arm also communicating, via an inlet opening, with the annular outer duct.

7. A turbojet engine according to claim 6, in which the apertures are oriented transversely with respect to the axis of revolution.

8. A turbojet engine according to claim 6, in which the apertures have an upstream-to-downstream orientation.

9. A turbojet engine according to any one of claims 6 to 8, in which the radially outer end of the wall of each flame catcher arm is fixed to the second wall and defines the inlet opening.

10. A turbojet engine according to any one of the preceding claims, in which surface fuel ducts are provided in the extrados face of the first part of at least some of the two-part vanes and have apertures which are oriented transversely with respect to the axis of revolution and which form fuel injection orifices.

11. A turbojet engine according to any one of the preceding claims, in which substantially radially extending fuel ducts are interposed between the trailing edges of the second parts of the two-part vanes and the flame catcher arms, and have apertures which are oriented substantially parallel to the axis of revolution and which form fuel injection orifices.

12. A turbojet engine according to claim 1, substantially as described with reference to Figures 1 to 4, Figure 5 or Figure 6 of the accompanying drawings.