FR2947035A1 - Combustion chamber for use in gas turbine engine of vehicle i.e. aerial vehicle, has central part assembled with external annular ferrule, and cooling air inlet openings distributed on surface of central part - Google Patents

Abstract

The chamber has an external annular ferrule (11) provided with openings (11r) for supplying cooling air. A central part (16b) is curved towards an axis of the chamber. An annular metal band is welded with an internal wall of the external ferrule by the curved central part. The openings are opened at right side of lateral parts (16a, 16c). The welded central part is assembled with the ferrule, and cooling air inlet openings (16b1) are distributed on a surface of the central part.

Description

The present invention relates to gas turbine engines and more particularly the combustion chambers of these engines. A gas turbine engine comprises at least one assembly, the rotor rotating about an axis and a generally annular combustion chamber. The rotor is formed of at least one rotary compressor compressing the air admitted into the engine and supplying the combustion chamber with compressed air. The compressed air is mixed with a fuel entering the chamber to feed a combustion chamber. The energy produced by the combustion is recovered at least in part in the turbine stages located downstream of the chamber, in which the gas is allowed to relax. One of the turbine stages is connected by a shaft to the compressor that it drives. This type of engine is applied both in industry and for the propulsion of vehicles, including air. Depending on the applications the engine is likely to include a free turbine.

The combustion chamber is disposed between the compression section and the turbine section. According to the architecture of the engine, the gas flow is caused to undergo more or less significant changes of direction between the compressor outlet and the inlet of the first turbine stage.

The invention is more particularly directed to combustion chambers whose flow is reversed. Inside the chamber, the combustion gases flow in a direction opposite to that of the air flow from the compressor. The combustion gases leaving the flame tube undergo a change of direction again to be guided towards the inlet of the turbine. This type of combustion chamber is found in motors, for example used for driving helicopter rotors with axial-radial compression. These are compact motors of relatively small size, for which the radial space is compensated by a reduced axial length.

This type of combustion chamber generally comprises an annular flame tube around the axis of the engine; the chamber is disposed in a substantially cylindrical casing with which the outer wall of the flame tube provides an annular space for the flow of air to be admitted into the chamber. The dilution and combustion air is admitted into the chamber through holes provided along the walls, outer and inner, and by the fuel injectors. The latter, depending on the chamber model, are arranged on the outer wall of the flame tube or on the downstream transverse wall, called the dome, closing the annular space between the two annular rings of the flame tube.

For the cooling of the walls, in order to ensure a suitable circulation of the air along their internal faces and to form a film of cold air, it is known to dispose of the baffles at the right of orifices, made by micro-drilling, to deflect and guide the air parallel to the wall either upstream or downstream.

To simplify the manufacture of the chamber, it is known to make the deflector in the area immediately downstream of the combustion chamber to weld a metal plate on the outer shell of the flame tube. This blade is shaped so as to form two annular tabs free edge. One is facing upstream, the other facing downstream to guide two opposing cooling air films along the inner face of the outer shell. This deflector has in axial section a substantially omega-shaped curved section, with a central portion curved towards the axis of the flame tube and a lateral portion in the form of a tongue on either side of the central portion. It is welded by its central part to the shell. The temperature to which the parts are subjected affecting their life, we try as far as possible to keep it as low as possible given the operating conditions. However, the control of the temperature of the tongues of the deflector 10 presented above is delicate at the welded zone on the shell of the chamber. The small size of the engines of the type in question induces openings for holes and film thicknesses of small dimensions, with flow control difficult to achieve. This control is difficult to obtain especially in boilermaking which is the current manufacturing mode of these parts for its relatively low cost. For example, radial openings of relatively large diameter, possibly inclined tangentially, have already been made through the central welded part of the deflector to supply the zone of the chamber situated between the two blades with cooling air. The results are not sufficient in terms of limiting the heating of the wall. The object of the invention is to improve the service life of the combustion chambers having an omega-shaped deflector for guiding the cooling air.

The object of the invention is more particularly to improve the cooling of the chamber wall in the area concerned by the omega-shaped deflectors. According to the invention, the desired result is achieved with a gas turbine engine combustion chamber comprising an outer annular shroud, pierced with orifices for supplying cooling air to said chamber and an annular metal strip with axial section. omega forming a deflector with a central portion and two side portions, the central portion of the deflector being curved towards the axis and the side portions being tongue-shaped, said deflector being welded to the inner wall of the ferrule by its central portion curved, first orifices pierced in the ferrule opening to the right of the tabs for supplying opposing cooling films. The chamber is characterized in that said central portion of the deflector is pierced together with the outer shell of second cooling air intake ports between said two opposite films, in multi-piercing. As is known in the field of turbomachines, the orifices forming a multi-piercing are of small diameter and are evenly distributed over the surface to be cooled. Their diameter and their spacing being defined by the rate of permeability of the desired wall. The diameter of the holes is generally smaller than the space between the holes. Multi drilling eliminates hot spots and axial and azimuth temperature gradients. Thus, in operation, it has been found that the wall temperature in the zone considered was lower by more than 100 ° C. compared with a cooled chamber according to the technique of the prior art. The reduction of the thermal gradients leads to a reduction of thermomechanical stress generating cracks located on the outer casing of the flame tube. The lifetime of the component is increased. According to another characteristic, the second orifices are micro-drills with a diameter of between 0.3 and 0.8 mm, preferably between 0.3 and 0.6 mm, the number of bores per unit area being determined to obtain the desired permeability. According to a particularly advantageous characteristic, the micro-holes are inclined tangentially at an angle of between 50 ° and 70 °, preferably 60 ° relative to the normal of the wall. More particularly and in accordance with a second aspect of the invention the bores are inclined in the same direction as the gyration of gases generated by the rotation of the compressor. This has the effect of creating a film of gyratory air close to the wall. This film of air comes to protect the radiation of the flame, the inside of the wall of the envelope and the inner baffle. In addition, the inclination of the bores has the effect of doubling the thickness of material traversed by the air, which accordingly increases the convective exchange surface. The invention also relates to the gas turbine engine comprising a combustion chamber thus defined. The invention will be better understood and other objects, details, features and advantages thereof will appear more clearly in the explanatory description, detailed, which follows of a non-limiting embodiment with reference to the accompanying drawings, on FIG. 1 is a schematic axial sectional view of a gas turbine engine to which the invention applies. FIG. 2 shows in axial half-section the combustion chamber of the gas turbine engine. FIG. 3 is an enlarged view of the zone for forming the cooling films of the outer shell of the chamber of FIG. 2; FIG. 4 is a sectional view AA of the central portion of the zone of FIG. forming the cooling films, pierced with holes according to the invention. Referring to Figure 1, we see a gas turbine engine free turbine made in two parts, a first part comprising all the elements constituting the gas generator and a second part comprising the elements constituting the free turbine and the main gearbox. The first part comprises a compressor 3, shown here in the form of a centrifugal axial compressor in which the air enters through a suction nozzle 2 and which rejects the compressed air in an outlet volute 3 '. The volute is arranged in a ring around the compressor with a continuously increasing section to collect this compressed air and send it through an upstream transfer pipe 4 into an annular combustion chamber 10 where it participates in the combustion of a fuel injected by one or more injectors 6. The chamber is contained in an envelope 5.

The gases resulting from this combustion are guided by the channel 4 ', so as to be deflected at 180 °, and firstly relaxed in a turbine 7, concentric with the annular combustion chamber and connected to the compressor 3 by a shaft motor, then collected in an annular manifold 8 to pass into the second part of the engine. The rotating parts of the gas generator that require lubrication such as bearings and gears are contained in a crankcase 9. Figure 2 shows in axial half-section the combustion chamber only 10. This is formed of a outer annular shell 11, an annular inner ring 12 both being connected by a ferrule 13 in the form of toroidal portion. The latter constitutes the dome in which fuel injectors 14 are arranged. These elements form the annular flame tube of axis coaxial with that of the engine. The combustion air is admitted through openings in the walls of the tube, such as the radial openings 11d and 12d in the annular ferrules and the axial openings in the dome 13 and the injectors 14. The combustion gases produced in the tube are guided downstream by a wall 17 to the turbine, not visible in Figure 2. The walls are cooled by air films that are generated along the tube. The dome is thus cooled by a film formed by a deflector 13a extending axially on a portion of the dome a short distance from the wall. The film is fed by multiple holes, not shown, of the wall 13 which open to the right of the deflector 13a. The air is thus deflected to form a film along the inner wall of the dome. The tube comprises a plurality of other films such as those downstream, 12a and 12b, formed by the deflectors 13b, between the dome and the inner ferrule 12. The outer ferrule is cooled by means of a deflector in the form of a double tongue 16, which is enlarged, also in FIG. 3. This double tongue deflector is constituted by an annular metal strip 16 which has been deformed so as to take in axial section a form of omega capital, called omega form . There is a central portion 16b welded to the wall, with here a straight portion terminated on both sides by a curved connecting portion 16b '. The two connecting portions 16b 'each extend from a rectilinear portion 16a and 16c respectively forming a tongue. The two tabs are parallel to the shell 11, with a spacing sufficient to allow the formation of a film of air along the wall of the flame tube. The deflector 16 is welded to the shell 11 along its central portion 16b. According to the usual technique, it is a welding wheel. Microfoils 11r are formed in the wall 11 for the passage of cooling air from the outer channel to the chamber, formed between the chamber and the motor housing. The air passing through the holes 11r is deflected by the tongue 16a respectively 16c, in opposite directions, and forms two wall cooling films, one upstream the other downstream. According to the prior art openings are provided through the thickness formed by the central portion 16b of the deflector 16 and the shell 11. This air ensures the supply of the dilution air chamber downstream of the primary combustion zone located between the deflector 16 and the dome, upstream.

However, it is found with this embodiment of the prior art that the cooling of the wall and the tabs 16b and 16c does not prevent local overheating parts. According to a first aspect of the invention, the orifices of relatively large diameter have been replaced in the central part of the baffle by multi-drilling operations. These are holes (16b1) of small diameter distributed in the central zone. The diameter is between 0.3 and 0.8 mm, preferably between 0.3 and 0.5 mm and more particularly in the context of an example of a combustion chamber of 0.4 mm diameter bores. The holes are evenly distributed over the entire surface of the central zone in a pattern formed of rows, circumferentially, perpendicular to the axis of the chamber, the spacing between the orifices being advantageously between 1 and 10 mm. By the multiplication of the orifices, a more homogeneous cooling has been obtained in this zone. The number, the diameter and the distribution of the bores are determined so as to ensure a sufficient flow of cooling air and the evacuation of heat in the wall while reducing the axial and longitudinal thermal gradients.

According to a second aspect of the invention, the protection was further improved with bores tilted tangentially at an angle of between 50 and 70 ° more particularly 60 °, as seen in Figure 4, in the direction of rotation gases produced by the compressor. This creates a cooling film close to the wall that protects it from hot flows.

This increased protection results from an improvement in the temperature distribution on the circumference of this central zone. Tests were carried out on a combustion chamber of the type used on the engine of FIG. 1. The outer shell of the flame tube has a thickness of the order of 0.8 mm for a diameter of 360 mm. Three circular rows were drilled, each 340 holes 0.4 mm in diameter, so as to have a uniform distribution of the latter. The inclination was 60 °. It was found that the maximum temperature measured on the central zone was decreased by 200 ° C and that the axial temperature gradient was divided by three compared to a chamber without inclination of the bores. Insofar as the corrosion and deterioration of the walls of the chamber are essentially a function of the resistance of the material to the temperature of use. The maximum temperature of the wall in this zone is lower, the incidences related to corrosion and oxidation are reduced. Compared to the solution with radial multi-drills, the inclination does not penalize the operation of the engine in terms of flow rate and pressure drop, in particular the pumping margin throughout the flight range is retained.

Claims (5)

CLAIMS1) Combustion chamber (10) of a gas turbine engine comprising an outer annular shroud (11) pierced with orifices for supplying cooling air to said chamber and an annular metal strip (16) with axial section omega forming a deflector with a central portion (16b) and two side portions (16a, 16c), the central portion being curved toward the axis of the chamber and the side portions being tongue-shaped, said band (16) being welded to the inner wall of the outer shell (11) by its curved central part Zo, first orifices (11r) drilled in the shell (11) opening to the right of the lateral parts (16a, 16c), characterized in that said welded central portion (16b) is pierced, together with the ferrule (11), second orifices (16b1) cooling air intake, in multi-holes distributed on the surface of the central portion. 15 2) Combustion chamber according to the preceding claim, the second orifices (16b1) are micro-drilled with a diameter between 0.3 and 0.8mm, preferably between 0.3 and 0.5mm. 3) Combustion chamber according to the preceding claim, the micro-holes are inclined tangentially at an angle of between 50 and 70 °, preferably 60 °. 4) Combustion chamber according to the preceding claim wherein the combustion gases flow with a direction of gyration around the motor axis, said micro holes being inclined in the same direction of gyration. 5) A gas turbine engine comprising a combustion chamber 25 according to one of the preceding claims.