Blade for a turbomachine turbine, comprising internal passages for circulating cooling air

In order to limit the air flow required to cool its blade member, a blade for an aircraft turbomachine turbine includes, in its blade member, a main passage configured to collect a main air flow from an air inlet and to circulate it in the blade member, and a side passage located between the main passage and the pressure-side wall, while being separated from the main passage, wherein the side passage is configured to collect a side air flow from the air inlet and to circulate it in the blade member, the main passage having at least three portions connected together end-to-end to form a coil, the side passage being located between each of these portions and the pressure-side wall.

TECHNICAL FIELD

The present invention relates to the field of aircraft turbomachines and more particularly relates to the cooling of the moving blades of the turbines within such turbomachines.

STATE OF PRIOR ART

Within a turbomachine used to propel an aircraft, turbine blades are subjected to high heat of the combustion gases from the combustion chamber. It is in particular the case of blades directly arranged at the outlet of the combustion chamber, such as the blades of a high pressure turbine in a multi-spool turbomachine.

To protect the blades from these high temperatures, it is known to cool the blades by means of an airflow circulating in cavities provided inside the airfoils of the blades. These cavities are generally supplied with relatively fresh air taken at a compressor stage of the turbomachine.

Document FR 3 020 402 A1 describes an example of internal configuration of a blade provided with such cavities.

However, the constant improvement in the aircraft engine performance results in an increase in the temperature of the combustion gases.

But, progress made as regards materials and coatings used to form the blades do not make it possible on their own to compensate for the rise in the temperature of the combustion gases.

Consequently, it is desirable to improve the cooling of the blades.

However, it is desirable in the same time to limit the airflow rate used to cool the blades.

Indeed, taking cooling air for the blades at a compressor of the turbomachine (typically the high pressure compressor in the case of a two-spool turbomachine) tends to degrade the general turbomachine performance, in particular in terms of specific consumption.

DISCLOSURE OF THE INVENTION

The purpose of the invention is in particular to provide a simple, economic and efficient solution to this problem.

To that end, it provides a blade for an aircraft turbomachine turbine, comprising a blade root comprising at least one air inlet, and an airfoil carried by the blade root and extending along a span direction of the blade terminating in an apex of the blade, the airfoil having a pressure side wall, and a suction side wall connected to the pressure side wall at a leading edge and a trailing edge of the airfoil.

The airfoil comprises in particular a main cavity configured to collect a main airflow from the air inlet and circulate it in the airfoil and exhaust it from the main cavity by first outlet ports, and a side cavity interposed between the main cavity and the pressure side wall by being separated from the main cavity, the side cavity being configured to collect a side airflow from the air inlet and circulate it in the airfoil and exhaust it from the side cavity by second outlet ports.

According to the invention, the main cavity includes at least three portions connected to each other end-to-end so as to form a serpentine shape, the side cavity being interposed between each of the portions and the pressure side wall.

In use, the main cavity thus enables a main airflow circulating in the same to follow a relatively long path, which makes it possible to maximise heat exchanges between this main airflow and the walls delimiting the main cavity, whereas the side cavity enables a heat shield to be formed between the entire main cavity and the pressure side wall, which is generally particularly hot, so as to limit heating of the main airflow through the pressure side wall. Consequently, the main airflow remains relatively fresh at the outlet of the main cavity and hence can be exploited to cool another portion of the airfoil, as will appear more clearly in what follows.

The invention thus generally makes it possible to limit the flow rate necessary as regards the main airflow circulating in the main cavity.

The side cavity is preferably a so-called “thin” cavity, that is having a low thickness in comparison with the main cavity. Thus, the mean thickness of the side cavity is preferentially lower than one third of the mean thickness of the main cavity.

In one preferred embodiment of the invention, the portions of the main cavity are adjacent to the suction side wall.

The main airflow circulating in the main cavity thus allows cooling of the suction side wall.

Preferably, the airfoil further comprises a leading edge cavity adjacent to the leading edge and separated from the air inlet.

In the preferred embodiment of the invention, the portions of the main cavity advantageously comprise a trailing edge side portion connected to the air inlet, a leading edge side portion connected to the leading edge cavity through the first outlet ports so as to supply the leading edge cavity with cooling air, and at least one intermediate portion connecting the trailing edge side portion to the leading edge side portion. In addition, the airfoil includes third outlet ports through which the leading edge cavity communicates with the outside of the blade.

The main airflow at the outlet of the main cavity can thus be exploited to cool the walls of the airfoil forming the leading edge.

Consequently, in the preferred embodiment of the invention, the pressure side wall, being relatively hot, is cooled by the side airflow which is exclusively used for cooling it, whereas the suction side wall, which is less hot, is cooled by the main airflow which is also used to cool the leading edge cavity.

Thus, the thermal heating of the airflow circulating in the main cavity is limited by the presence of the side cavity which protects the main cavity from the pressure side wall, being relatively hot. Controlling the thermal heating of the airflow circulating in the main cavity and cooling the airfoil walls forming the leading edge through impingement makes it possible to ensure a limited temperature of the airfoil at the leading edge.

Preferably, the first outlet ports are impingement cooling ports.

In this case, the main airflow comes out of the main cavity as jets projected on the walls of the airfoil forming the leading edge, thus allowing an impingement cooling of these walls at the leading edge.

In preferred embodiments of the invention, the airfoil further comprises a rear cavity, separated from the main cavity and from the side cavity, and configured to collect a rear airflow from the air inlet and circulate it in the airfoil and exhaust it from the rear cavity through fourth outlet ports.

Such a rear airflow, distinct from the main airflow and from the side airflow, can be exploited to cool the region of the airfoil close to the trailing edge.

Thus, the airfoil preferably comprises a first trailing edge cavity adjacent to the trailing edge, separated from the air inlet, and connected to the rear cavity through the fourth outlet ports so as to receive at least one portion of the rear airflow and circulate it in the airfoil and exhaust it from the first trailing edge cavity through fifth outlet ports.

In addition, the airfoil preferably includes a second trailing edge cavity adjacent to the trailing edge, adjacent to the first trailing edge cavity along the span direction, separated from the rear cavity, and configured to collect another rear airflow from the air inlet and circulate it in the airfoil and exhaust it from the second trailing edge cavity through sixth outlet ports.

The invention also relates to a turbine for an aircraft turbomachine, comprising at least one rotary disk provided with blades of the type described above.

The invention also relates to a turbomachine for an aircraft, comprising at least one turbine of the type described above.

The invention finally relates to a moulding kit for manufacturing a blade of the type described above, comprising a set of cores configured to form at least the main cavity and the side cavity of the airfoil.

DETAILED DISCLOSURE OF PREFERRED EMBODIMENTS

FIG. 1illustrates a turbomachine10for an aircraft, for example a bypass twin-spool turbofan engine, generally including a fan12for sucking an airflow divided downstream of the fan into a primary flow supplying a core of the turbomachine and a secondary flow bypassing this core. The core of the turbomachine generally includes a low pressure compressor14, a high pressure compressor16, a combustion chamber18, a high pressure turbine20and a low pressure turbine22. The turbomachine is shrouded with a nacelle24surrounding the secondary flow channel26. The rotors of the turbomachine are rotatably mounted about a longitudinal axis28of the turbomachine.

In the entire description, the axial direction X is the direction of the longitudinal axis28of the turbomachine, the radial direction R is at any point a direction orthogonal to the axial direction X and intersecting the same, and the tangential direction T is at any point orthogonal to both previous directions. On the other hand, the “upstream” and “downstream” directions are defined in reference to the general gas flow in the turbomachine.

FIG. 2illustrates a blade30for an aircraft turbomachine turbine according to one preferred embodiment of the invention, for example a blade for a high pressure turbine, generally comprising a blade root32carrying an airfoil34terminating with an apex35of the blade on the side opposite to the blade root32. The blade root includes a radially inner portion for enabling the blade to be retained by jointing in a rotor disk, in a known manner per se. In addition, the blade root is connected to the airfoil through an aerodynamic platform36for delimiting internally a primary flow channel within the turbine.

In reference toFIGS. 2 and 3, the airfoil34generally includes a pressure side wall40and a suction side wall42connected to each other at a leading edge44and at a trailing edge46of the airfoil.

InFIGS. 2 and 3, the axial direction X corresponds to the direction of the longitudinal axis28of the turbomachine when the blade30is mounted to a rotor disk within the high pressure turbine20. Under the same conditions, the direction from the blade root32to the apex35of the airfoil, called a span direction WS of the blade, coincides with the radial direction R.

The airfoil34generally comprises internal cavities each extending along the span direction WS of the airfoil, and configured for receiving and circulating cooling airflow taken in the blade through an air inlet48at the blade root32(FIG. 2).

The airfoil34comprises a main cavity50, more particularly visible inFIG. 5, configured to collect a main airflow F1from the air inlet48and circulate it in the airfoil34and exhaust from the main cavity through first outlet ports. The main cavity50includes three portions50A,50B,50C connected to each other end-to-end so as to form a serpentine shape. Such a configuration enables the path length of the main airflow F1to be maximised within the main cavity50, and thus heat exchanges between the main airflow F1and the walls delimiting the main cavity50to be maximised.

In the preferential example illustrated, each of the internal cavities has a dimension along the span direction WS higher than its extent according to any direction orthogonal to the span direction WS. Preferably, each of the internal cavities has a dimension along the span direction WS higher than several times its extent along any direction orthogonal to the span direction WS.

On the suction side, the three portions50A,50B,50C are each delimited by the suction side wall42, whereas on the pressure side, the three portions50A,50B,50C are each delimited by a first internal wall52of the airfoil (FIG. 3).

In addition, the portions50A and50B are separated from each other by a second internal wall53A connecting the suction side wall42to the first internal wall52. Analogously, the portions50B and50C are separated from each other by a third internal wall53B connecting the suction side wall42to the first internal wall52.

The airfoil34further comprises a side cavity54, more particularly visible inFIG. 4, configured to collect a side airflow F2from the air inlet48and circulate it in the airfoil34and exhaust from the side cavity through second outlet ports56. The side cavity54thus has a blade root side end54A connected to the air inlet48. In addition, the pressure side wall40includes the second outlet ports56enabling the side airflow F2to exhaust from the blade (FIG. 2), and the side cavity54has a closed apex side end54B.

The side cavity54is delimited, on the suction side wall side, by the first internal wall52, and on the pressure side, by the pressure side wall40(FIG. 3).

Thus, the side cavity54is separated from the main cavity50by the first internal wall52of the airfoil.

The first internal wall52is a solid wall, thus prohibiting any air exchange between the side cavity54and the main cavity50.

The side cavity54is in particular interposed between each of the three portions50A,50B,50C of the main cavity, and the pressure side wall40, so as to form a heat shield between the three portions50A,50B,50C of the main cavity and the pressure side wall.

The side cavity54thus enables heating of the main airflow F1to be limited within the main cavity50by the pressure side wall40, which is generally hotter than the suction side wall42in use.

To that end, the side cavity54is preferably a so-called “thin” cavity, that is having a low thickness in comparison with the main cavity50. Thus, the mean thickness of the side cavity54is preferentially lower than one third of the mean thickness of the main cavity50. By “mean thickness”, it is intended the extent measured perpendicular to the bend line58of the airfoil, averaged along said bend line and along the span direction WS.

Generally, the three portions50A,50B,50C of the main cavity are thus adjacent to the suction side wall42whereas the side cavity52is adjacent to the pressure side wall40.

The airfoil34further comprises a leading edge cavity60adjacent to the leading edge44(FIG. 3). This leading edge cavity60is separated from the air inlet48, in that this cavity is not directly connected to the air inlet48but receives cooling air from the main cavity50.

More particularly, the three portions of the main cavity50comprise a trailing edge side portion50A connected to the air inlet48for taking in the main airflow F1, a leading edge side portion50C connected to the leading edge cavity60so as to supply the same with cooling air, and an intermediate portion50B connecting the trailing edge side portion50A to the leading edge side portion50C (FIGS. 3 and 5).

The trailing edge side portion50A thus has a blade root side end50A1(that is a radially inner end) connected to the air inlet48, and an apex side end50A2opening into a first connection50D connected to an apex side end50b2of the intermediate portion50B (FIG. 5).

In addition, the intermediate portion50B has a blade root side end50b1opening into a second connection50E connected to a blade root side end50C1of the leading edge side portion50C.

Finally, the leading edge side portion50C has a closed apex side end50C2.

It is to be noted that the apex side end54B of the side cavity54extends facing the first connection50D and the apex side end50C2of the leading edge side portion50C of the main cavity50.

The leading edge cavity60is separated from the leading edge side portion50C through a third internal wall62, which connects the pressure side wall40to the suction side wall42, and to which a leading edge side end52A of the first internal wall52is connected (FIG. 3).

In the embodiment illustrated, communication of the leading edge side portion50C with the leading edge cavity60is ensured by impingement cooling ports64formed in the third internal wall62(FIG. 5), and thus making up the aforementioned first outlet ports, in the terminology of the invention. The leading edge cavity60has a closed blade root side end60A, and an apex side end60B which is also closed.

In addition, the leading edge cavity60communicates with the outside of the blade by means of third outlet ports66formed in at least one of the pressure side40and suction side42walls in the proximity of the leading edge44(FIG. 2).

On the other hand, the airfoil34comprises a rear cavity70(FIGS. 3 and 4) separated from the main cavity50and from the side cavity54, and configured to collect a rear airflow F3from the air inlet48and circulate it in the airfoil and exhaust it from the rear cavity70by fourth outlet ports. The rear cavity70thus has a blade root side end70A connected to the air inlet48.

The separation between the rear cavity70on the one hand and the main cavity50and the side cavity54on the other hand is ensured by a fifth internal wall72of the airfoil, which is a solid wall connecting the pressure side wall40to the suction side wall42, and to which a trailing edge side end52B of the first internal wall52(FIG. 3) is connected.

The rear cavity70is thus delimited by the pressure side40and suction side42walls.

In preferred embodiments, the airfoil34finally includes at least one trailing edge cavity adjacent to the trailing edge46, and communicating preferably with the outside of the blade through slots74opening into the trailing edge46, in a known manner per se (FIG. 3).

In the example illustrated, the airfoil34includes two trailing edge cavities80,82which are adjacent along the span direction WS of the blade. These two trailing edge cavities80,82are separated from the rear cavity70by a sixth internal wall84of the airfoil connecting the pressure side wall40to the suction side wall42(FIG. 3).

Among both cavities, a trailing edge first cavity80(FIG. 4), disposed on the apex side35of the blade, is connected to the rear cavity70through calibrated ports86formed in the sixth internal wall84, and thus making up the aforementioned fourth outlet ports, in the terminology of the invention. By “calibrated ports”, it is intended ports the passage cross-section of which is chosen upon design to obtain a desired airflow rate, which is determined by the thermal behaviour of the blade in this region. The trailing edge first cavity80has a closed blade root side end80A, and an apex side end80B which is also closed. The trailing edge first cavity80is thus configured to receive at least one portion of the rear airflow F3and circulate it in the airfoil and exhaust it from the trailing edge first cavity80by fifth outlet ports, made up by some of the slots74.

A second trailing edge cavity82, disposed on the side of the blade root32, is separated from the rear cavity70by a solid portion of the sixth internal wall84, and is configured to collect another rear airflow F4directly from the air inlet48and circulate it in the airfoil and exhaust it from the second trailing edge cavity82through sixth outlet ports, made of other slots74. The second trailing edge cavity82thus has a blade root side end82A connected to the air inlet48, and a closed apex side end82B.

The blade30can be made by moulding, in a known manner per se.

To that end, a moulding kit for manufacturing a blade according to the invention comprises a set of cores configured to form the different cavities of the airfoil respectively.

Such a set of cores, visible inFIG. 6, thus includes a serpentine-shaped main core N1(more clearly visible inFIG. 7) having the form of the main cavity50, a side core N2having the form of the side cavity54, a leading edge core N3having the form of the leading edge cavity60, a rear core N4having the form of the rear cavity70, a trailing edge first core N4having the form of the trailing edge first cavity80, and a second trailing edge second core N5having the form of the second trailing edge cavity82.

In use, the main airflow F1(FIG. 5) penetrates the trailing edge side portion50A of the main cavity50through its blade root side end50A1, and circulates radially outwardly up to the first connection50D, and then penetrates the intermediate portion50B and circulates radially inwardly in the same, up to the second connection50E. The main airflow F1then radially rises outwardly in the leading edge side portion50C and penetrates the impingement cooling ports64(or first outlet ports), of which it comes out as impingement jets F1A within the leading edge cavity60. The main airflow F1finally exhausts from the blade by passing through the third outlet ports66close to the leading edge44.

The main airflow F1thus allows cooling of the suction side wall42by circulating along the three portions50A,50B,50C of the main cavity, as well as impingement cooling of the walls40and42of the airfoil at the leading edge44, by taking the form of a plurality of impingement jets F1A at the outlet of the impingement cooling ports64.

On the other hand, the side airflow F2(FIG. 4) penetrates the side cavity54through the blade root side end54A thereof, and then this flow circulates radially outwardly and is distributed into the second outlet ports56(FIG. 2) so as to exhaust from the airfoil34. The side airflow F2enables the first internal wall52to be cooled and thus a heat shield to be formed between the pressure side wall40and the main cavity50, contributing in that the main airflow F1remains relatively fresh when it reaches the impingement cooling ports64.

Hence, the invention enables the necessary flow rate to be limited as regards the main airflow F1, while ensuring an optimum cooling of the suction side wall42and the region of the pressure side40and suction side42walls close to the leading edge44, through the main airflow F1, as well as an optimum cooling of the pressure side wall40through the side airflow F2.

Further, the rear airflow F3(FIG. 4) penetrates the rear cavity70through the blade root side end70A thereof, and then this flow circulates radially outwardly and is distributed into the different calibrated ports86(or fourth outlet ports) through which this flow penetrates the trailing edge first cavity80. The rear airflow F3then exhausts from the airfoil by passing through the slots74or fifth outlet ports (FIG. 3) communicating with the trailing edge first cavity80.

The other rear airflow F4penetrates the second trailing edge cavity82through the blade root side end82A thereof, and circulates radially outwardly so as to be distributed into the slots74or sixth outlet ports communicating with the second trailing edge cavity82, and then to thus exhaust from the airfoil34.

Other configurations of the blade30are possible without departing from the scope of the present invention defined by the set of appended claims.

For example, the blade30can include three intermediate portions connected in series so as to form themselves a serpentine shape connecting the trailing edge side portion50A to the leading edge side portion50C. In this case, a median intermediate portion among these three intermediate portions can be connected to the air inlet48much like the trailing edge side portion50A, so as to improve controlling heating of the airflow circulating in the serpentine shape. The connection of the median intermediate portion to the air inlet48is preferably configured to let an airflow rate higher than the flow rate of the connection of the trailing edge side portion50A to the air inlet48pass.