FR3070058A1 - AIRCRAFT TURBOMACHINE COMPRISING A COOLING ELEMENT IMPROVING CONVECTION COOLING AND PROVIDING AIR JET IMPACT COOLING OF AN ANNULAR COMBUSTION ROOM TERMINAL LINK FLANGE - Google Patents

Abstract

To improve the cooling of a combustion chamber annular wall end wall (18; 20), a turbomachine comprises a cooling element (50; 52) extending from a housing (30; 32) towards the combustion chamber so as to define with the terminal connection flange a downstream portion (36B; 38B) of an annular chamber bypass space having an air passage section ( S36B; S38B) smaller than that (S36A; S38A) of an upstream portion (36A; 38A) of this space. The cooling element defines with the housing a chamber (54; 80) having air inlet (56; 82) and air outlet (58B; 84B) ports formed in the cooling element.

Description

AIRCRAFT TURBOMACHINE COMPRISING A COOLING ELEMENT IMPROVING CONVECTION COOLING AND PROVIDING IMPACT AIR JET COOLING OF AN ANNULAR COMBUSTION CHAMBER WALL TERMINAL

DESCRIPTION

TECHNICAL AREA

The present invention relates to an aircraft turbomachine, such as an aircraft turbojet or turboprop, comprising an annular combustion chamber.

PRIOR STATE OF THE ART

In the state of the prior art, illustrated by patent applications FR2970512 and FR2904048 of the applicant, the combustion chamber comprises two walls of coaxial revolution, namely an external wall and an internal wall, having respective first ends connected to the '' to one another by a chamber bottom wall, and second respective opposite ends provided with respective terminal connecting flanges, and a first casing delimiting, between the latter and a first of the external and internal walls, an upstream portion a first annular space allowing the passage of a first air flow along and outside the annular combustion chamber.

Part of the so-called “bypass” air, circulating along the external and internal walls of the annular combustion chamber, outside the latter, participates in the cooling of these walls, and another part of this air continues. its circulation downstream to supply cooling circuits of other turbomachine components, in particular within a turbine.

However, the bypass air does not effectively cool the respective terminal connection flanges of the outer and inner walls of the annular combustion chamber.

STATEMENT OF THE INVENTION

The object of the invention is in particular to provide a simple, economical and effective solution to this problem.

This object is achieved, according to the invention, because the turbomachine further comprises a first cooling element, which has a shape of revolution and extends from the first casing towards the annular combustion chamber so as to delimit, between the first cooling element and the terminal connecting flange of said first of the external and internal walls, a downstream portion of the first annular space, having an air passage section smaller than an air passage section of the upstream portion of the first annular space, the first cooling element delimiting, between the latter and the first casing, a first annular chamber, and the first annular chamber comprising air inlet orifices and outlet orifices d formed in the first cooling element.

The first cooling element thus makes it possible to channel part of the first air flow within a space of reduced section, along the terminal connection flange considered, and thus makes it possible to increase the flow speed of the air and therefore improve the cooling of this terminal connection flange by convection.

The air inlet ports and the air outlet ports allow another part of the first air flow to continue to flow downstream directly, without undergoing considerable pressure drop due to the presence of the first cooling element.

In general, the respective air passage sections of the upstream and downstream portions of the first annular space not being constant, the sections considered above correspond to the minimum air passage sections of each of the upstream portions and downstream of the first annular space, that is to say at the air passage sections measured respectively at the respective narrowest regions of each of these upstream and downstream portions.

In a preferred embodiment of the invention, the first cooling element further comprises impact cooling holes configured to cool by impacts of air jets the terminal connecting flange of said first of the external and internal walls.

The air outlet ports typically have respective air passage sections considerably larger than respective air passage sections of the impact cooling holes.

In addition, the air outlet orifices are further away from the terminal connecting flange of said first of the external and internal walls than are the impact cooling holes.

The first cooling element thus also allows air jet impact cooling of the terminal connecting flange.

Preferably, the air outlet orifices are configured so that the sum of the respective air passage sections thereof and the air passage section of the downstream portion of the first annular space is strictly less than the air passage section of the upstream portion of the first annular space.

Such a configuration makes it possible to guarantee the acceleration of the air flow within the downstream portion of the first annular space, while ensuring, if necessary, that the air supplying the impact cooling holes has a pressure d '' a level allowing impact cooling under optimal conditions.

In the preferred embodiment of the invention, the first cooling element forms at least one bend partially delimited by the terminal connecting flange of said first of the external and internal walls, within the first annular space.

Preferably, the first cooling element is a plate having a first part comprising the air inlet orifices, a second part comprising the air outlet orifices and connected to the first part so that the first and second parts form a nose.

Preferably, the nose has the impact cooling holes.

In the preferred embodiment of the invention, the first part of the first cooling element has a free end in pre-stressed support on the first casing.

Such a configuration allows relative movement of the free end of the first cooling element and the first casing and thus makes it possible to effectively overcome possible phenomena of differential thermal expansion between the cooling element and the casing.

Preferably, the second part of the first cooling element is fixed to a fixing flange of the first housing.

Preferably, the terminal connecting flange of said first of the external and internal walls is formed of a flange, and the nose of the first cooling element is oriented radially inwards and arranged upstream relative to the flange.

As a variant, the terminal connecting flange of said first of the external and internal walls may comprise a bent end portion convex in the upstream direction having a free end in radial abutment on the first casing or on another casing extending the first casing, and the nose of the first cooling element can be oriented downstream and arranged between said first of the external and internal walls and the bent end part.

In the preferred embodiment of the invention, the turbomachine further comprises a second casing defining, between the latter and a second of the external and internal walls, an upstream portion of a second annular space allowing the passage of a second air flow along and outside the annular combustion chamber, and a second cooling element, which has a shape of revolution and extends from the second casing towards the annular combustion chamber to delimit, between the second cooling element and the terminal connecting flange of said second of the external and internal walls, a downstream portion of the second annular space, having an air passage section smaller than a passage section air from the upstream portion of the second annular space.

In addition, the second cooling element defines, between the latter and the second casing, a second annular chamber comprising air inlet orifices and air outlet orifices formed in the second cooling element.

The second cooling element operates similarly to the first cooling element, but being applied to the cooling of the terminal connecting flange of said second of the external and internal walls. This second cooling element can thus have all or part of the preferential characteristics set out above in relation to the first cooling element.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be better understood, and other details, advantages and characteristics thereof will appear on reading the following description given by way of non-limiting example and with reference to the appended drawings in which:

- Figure 1 is a partial schematic view in axial section of a turbomachine of a known type;

- Figure 2 is a view similar to Figure 1, of a turbomachine according to a preferred embodiment of the invention;

- Figure 3 is a partial schematic perspective view of an external cooling element belonging to the turbomachine of Figure 2, seen from downstream;

- Figure 4 is a partial schematic perspective view of the external cooling element of Figure 3, seen from upstream;

- Figure 5 is a partial schematic perspective view of an internal cooling element belonging to the turbomachine of Figure 2, seen from downstream;

- Figure 6 is a partial schematic perspective view of the internal cooling element of Figure 5, seen from upstream;

- Figure 7 is an enlarged view of detail VII of Figure 2, illustrating the external cooling element of Figures 3 and 4;

FIG. 8 is an enlarged view of detail VIII of FIG. 2, illustrating the internal cooling element of FIGS. 5 and 6.

Throughout these figures, identical references may designate identical or analogous elements.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

FIG. 1 shows an annular combustion chamber 10 within a turbomachine of a known type, such as an airplane turbojet or turboprop, as well as the environment close to this combustion chamber. The rest of the turbomachine will not be described and may be of a conventional type.

Throughout this description, the axial direction X is the direction of the longitudinal axis of the turbomachine, the radial direction R is at all points a direction orthogonal to the axial direction X and passing through the longitudinal axis of the turbomachine , and the tangential direction T is at any point orthogonal to the two previous directions so as to form a cylindrical coordinate system. Furthermore, in the absence of an indication to the contrary, the “upstream” and “downstream” directions are defined by reference to the general flow of the gases in the turbomachine.

The combustion chamber is arranged at the outlet of a bent diffuser 12 forming the outlet of a centrifugal type compressor (partially visible), and this combustion chamber 10 is followed by a high pressure turbine 14, of which only the distributor d entry 16 is shown.

The combustion chamber 10 comprises two coaxial frustoconical walls of revolution, namely an external wall 18 and an internal wall 20, which delimit between them the internal space 21 of the combustion chamber. These two walls are arranged one inside the other and typically have respective sections which reduce downstream.

The outer 18 and inner 20 walls have respective first ends, in this case upstream ends, connected to each other by a chamber bottom wall 22, respective second opposite ends, in this case ends downstream, provided with respective terminal connection flanges 24 and 26. The orientation of the first and second ends may be different, for example in the case of a reverse flow combustion chamber.

The turbomachine further comprises two coaxial casings, namely an external casing 30 and an internal casing 32, delimiting between them an enclosure 34 in which the combustion chamber 10 is housed. In the example illustrated, it should be noted that the casing internal 32 is fixed to an internal downstream casing 33 of the diffuser 12.

As shown in FIG. 1, the external casing 30 delimits, with the external wall 18, an upstream portion 36A of an external annular space 36, while the internal casing 32 delimits, with the internal wall 20, an upstream portion 38A d an internal annular space 38.

In the example illustrated, the terminal connecting flange 24 of the external wall 18 is formed of a collar extending radially outward and provided with a free end 24A bent downstream and extending at a distance radially of the outer casing 30. In addition, the terminal connecting flange 26 of the inner wall 20 comprises a bent end portion 26A convex in the upstream direction and having a free end 26B in radial abutment on an inner casing 40 of the turbine distributor 16, extending downstream of the internal casing 32, substantially in an extension of the latter, and connected to the latter. Other conformations of terminal connecting flanges 24, 26 are however possible within the scope of the present invention.

In operation, the air delivered by the diffuser 12 is divided into a central flow F1 supplying injection systems 42 and / or separate air inlet orifices arranged in the bottom wall of chamber 22, as well as an external flow F2 and an internal flow F3 circulating around the combustion chamber 10 respectively within the external annular spaces 36 and internal 38.

A part of each of the external F2 and internal F3 flows supplies air with various orifices formed in the external 18 and internal 20 walls of the combustion chamber, while the remaining part of each of these flows continues to flow downstream to supply ventilation or cooling circuits of other components of the turbomachine such as turbine blades or external turbine rings.

More specifically, the remaining part F2A of the flow F2 flows downstream beyond the terminal connecting flange 24 of the external wall 18 via an external passage 44 formed between the flange 24 and the external casing 30, while the remaining part F3A flow F3 flows to a downstream internal space through channels 46.

The role of the external F2 and internal F3 streams is furthermore to cool the external 18 and internal 20 walls of the combustion chamber and to supply, by microperforations formed in these walls, cooling films along the internal faces of these same walls. With flow velocities which are typically of the order of 30 meters per second, these flows F2 and F3 thus make it possible to reduce the temperature of the walls 18, 20 by around 300 degrees Celsius.

The external flow F2 and internal F3 also cool the terminal connection flanges 24 and 26.

In this regard, the inventors have noted that, on the one hand, the techniques of cooling by microperforations cannot be applied effectively to such connecting flanges, and that, on the other hand, the flow rates of the flows. F2 and F3 at the connection flanges, which are typically of the order of 15 meters per second, are too weak to allow convection cooling which is of a sufficient level, in particular with regard to the increase in temperatures of combustion chamber operation. However, such an increase in temperature is necessary for improving the performance of turbomachinery, and tends to increase the thermal gradients within the connection flanges which connect the hot walls of the combustion chamber to the relatively cold casings.

According to the invention, in order to improve the cooling of at least one of the terminal connection flanges 24 and 26, the turbomachine 11 illustrated in FIG. 2, similar to the turbomachine of FIG. 1, differs from the latter by the presence of at least one cooling element having two main functions: channeling part of the corresponding air flow within a space of relatively small section along the terminal connection flange considered, so as to increase the flow speed of this flow, and allow the flow of another part of this air flow downstream without considerable pressure drop.

In the preferred embodiment of the invention illustrated in Figures 2-8, the turbomachine 11 comprises two cooling elements of this type, namely an external cooling element 50 and an internal cooling element 52, aimed respectively at improving the cooling of the two terminal connection flanges 24 and 26.

These external cooling elements 50 and internal 52 can each constitute a "first" or a "second" cooling element, according to the terminology adopted above and in the appended claims. In particular, each of the external 50 and internal cooling elements 52 may have all or part of the preferential optional characteristics set out above and in the dependent claims appended with reference to the first cooling element, as will appear more clearly in the following. Similarly, any element or component qualified as internal or external in the present description may constitute a "first" or a "second" corresponding element or component according to the terminology adopted above and in the appended claims.

Thus, the external cooling element 50 has a shape of revolution and extends, from the external casing 30, in the direction of the annular combustion chamber 10, so as to delimit with the end connection flange 24 of the wall outer 18, a downstream portion 36B of the outer annular space 36, this downstream portion 36B having an air passage section S36B smaller than an air passage section S36A of the upstream portion 36A of the annular space external 36.

The narrowing of the air passage section in the downstream portion 36B relative to the upstream portion 36A of the outer annular space 36 makes it possible to accelerate the air flow flowing along the terminal connecting flange 24 and therefore to improve the cooling thereof.

Generally, the respective air passage sections of the upstream 36A and downstream 36B portions of the external annular space 36 are not constant, the sections S36A and S36B considered above correspond to the passage sections of minimum air from each of the upstream portions 36A and downstream 36B, that is to say at the air passage sections measured respectively at the respective narrowest regions of each of the upstream portions 36A and downstream 36B, as illustrated in FIG. figure 2.

In addition, as shown in FIGS. 3, 4 and 7, the external cooling element 50 defines, with the external casing 30, an external annular chamber 54 comprising air inlet orifices 56 and outlet orifices d 58B formed in the external cooling element 50.

The external cooling element 50 further comprises impact cooling holes 58A configured to cool by impacts of air jets the terminal connecting flange 24 of the external wall 18 of the chamber.

Referring to Figures 3 and 4, the air outlet ports 58B have respective air passage sections S58B larger than the respective air passage sections S58A of the impact cooling holes 58A, and are further apart of the terminal connection flange 24 that these impact cooling holes 58A are, the latter being arranged directly opposite the terminal connection flange 24.

The air outlet openings 58B allow part of the air circulating in the external annular space 36 to continue its downstream circulation directly, without undergoing a considerable pressure drop due to the presence of the external cooling element 50.

In addition, the sum of the respective air passage sections S58B of the air outlet orifices 58B and the air passage section S36B of the downstream portion 36B of the external annular space 36 is strictly less than the section air passage S36A of the upstream portion 36A of the outer annular space 36.

Such a configuration makes it possible to guarantee the acceleration of the air flow within the downstream portion 36B, and to ensure that the air supplying the impact cooling holes 58A has sufficient pressure to allow impact cooling in optimal conditions.

In the example illustrated, the number of air inlet ports 56 is equal to the number of air outlet ports 58B, and the diameter of the air inlet ports 56 is substantially equal to twice the diameter of the air outlet ports 58B.

In the preferred embodiment of the invention, the external cooling element 50 is a generally V-shaped plate of revolution, sectorized or in one piece, comprising:

a first part 60 of revolution arranged on the upstream side facing the general direction of air flow within the external annular space 36, and a second part 62 of revolution arranged on the downstream side facing the general direction of air flow within the external annular space 36.

The first part 60 comprises the air inlet orifices 56, and has, on the radially external side, a free end 60A with a curved edge downstream in pre-stressed support on the external casing 30.

The second part 62 has a conformation such that, viewed in axial section and traversed in a direction going radially from the inside to the outside, the second part 62 first comprises a radially internal portion 62A, then the second part 62 bends downstream in an intermediate portion 62B, then bends radially outward in a radially outer portion 62C.

The radially internal portion 62A comprises the impact cooling holes 58A, while the intermediate portion 62B comprises the air outlet orifices 58B.

Finally, the radially external portion 62C includes orifices 64 (FIGS. 3 and 4) for the passage of fixing members 65 which ensure the fixing of a fixing flange 66 of the external casing 30 (FIG. 7) to a fixing flange 68 of an external casing 70 of the turbine distributor 16 arranged downstream of the external casing 30. The fixing flanges 66 and 68 thus sandwich the radially external portion 62C.

The first and second parts 60, 62 are interconnected so as to form a nose 72 oriented radially inwards and arranged upstream of the terminal connecting flange 24, so that the radially internal portion 62A of the second part 62 and the terminal connecting flange 24 of the external wall 18 delimit between them a constriction 74 of the downstream portion 36B of the external annular space 36 (FIG. 7).

In the illustrated embodiment, when the turbomachine is seen in axial section, the external cooling element 50 thus forms, within the external annular space 36, an elbow 75 partially delimited by the terminal connecting flange 24.

Preferably, the first part 60 and the radially internal portion 62A of the second part 62 form an acute angle between them, so that the elbow 75 is at least 90 degrees, the angle of the elbow being measured in the direction of the air flow within the outer annular space 36.

Similarly, the internal cooling element 52 has a shape of revolution and extends, from the internal casing 32, towards the annular combustion chamber 10, so as to delimit with the connecting flange 26 of the internal wall 20, a downstream portion 38B of the internal annular space 38, this downstream portion 38B having an air passage section S38B smaller than an air passage section S38A of the upstream portion 38A of the space internal ring 38.

The narrowing of the air passage section in the downstream portion 38B relative to the upstream portion 38A of the internal annular space 38 makes it possible to accelerate the air flow flowing along the terminal connecting flange 26 and therefore to improve the cooling thereof.

The respective air passage sections of the upstream 38A and downstream 38B portions of the internal annular space 38 are not constant, the sections S38A and S38B considered above correspond to the minimum air passage sections of each of the portions upstream 38A and downstream 38B, that is to say at the air passage sections measured respectively at the narrowest regions of each of the upstream 38A and downstream 38B portions, as illustrated in FIG. 2.

In addition, as shown in FIGS. 5, 6 and 8, the internal cooling element 52 defines, with the internal casing 32, an internal annular chamber 80 comprising air inlet orifices 82 and outlet orifices d 84B formed in the internal cooling element 52.

The internal cooling element 52 further comprises impact cooling holes 84A configured to cool by impacts of air jets the terminal connecting flange 26 of the internal wall 20 of the chamber.

Referring to Figures 5 and 6, the air outlet ports 84B have respective air passage sections S84B larger than the respective air passage sections S84A of the impact cooling holes 84A, and are further apart of the terminal connecting flange 26 than are these impact cooling holes 84A.

The air outlet orifices 84B allow part of the air circulating in the internal annular space 38 to continue its downstream circulation directly, without undergoing a considerable pressure drop due to the presence of the internal cooling element 52.

In addition, the sum of the respective air passage sections S84B of the air outlet orifices 84B and the air passage section S38B of the downstream portion 38B of the internal annular space 38 is strictly less than the section air passage S38A of the upstream portion 38A of the internal annular space 36. Such a configuration makes it possible to guarantee the acceleration of the air flow within the downstream portion 38B, and to ensure that the air supplying the impact cooling holes 84A have sufficient pressure to allow impact cooling under optimal conditions.

In the example illustrated, the number of air inlet ports 82 is equal to the number of air outlet ports 84B, and the diameter of the air inlet ports 82 is substantially equal 5/3 of the diameter of the air outlet orifices 84B.

In the preferred embodiment of the invention, the internal cooling element 52 is a revolutionized plate, sectorized or in one piece, comprising:

a first part 90 of revolution arranged on the upstream side facing the general direction of air flow within the internal annular space 38, and a second part 92 of revolution arranged on the downstream side facing the general direction of air flow within the internal annular space 38.

The first part 90 extends substantially axially and has, on the upstream side, a free end 90A with a slightly curved edge radially outwards in pre-stressed support on the internal casing 32. This first part 90 comprises the inlet orifices of air 82.

The second part 92 has a conformation such that, viewed in axial section and traversed in a direction going radially from the outside to the inside, the second part 92 firstly comprises a radially external portion 92A extending substantially radially, then the second part 92 bends in the downstream direction into an intermediate portion 92B extending substantially axially, then the second part 92 bends radially outward into a radially internal portion 92C extending substantially radially. At the radially internal end of the radially internal portion 92C, the internal cooling element 52 curves in the upstream direction in its first part 90.

The radially internal portion 92C comprises the impact cooling holes 84A, while the radially external portion 92A comprises the air outlet orifices 84B and orifices 94 (FIGS. 5 and 6) for passage of fastening members 95 which provide the fixing of a fixing flange 96 of the internal casing 32 (FIG. 8) to a fixing flange 98 of the internal casing 40 of the distributor 16 arranged downstream of the internal casing 32. The fixing flanges 96 and 98 thus sandwich the radially internal portion 92C, for example with the interposition of a washer 100. In the example illustrated, the fixing flange 98 integrates the channels 46 allowing the circulation of the remaining part of the air flow downstream beyond the internal annular space 38.

As shown in FIGS. 5, 6 and 8, the intermediate 92B and radially internal 92C portions form, with a downstream portion of the first part 90, a nose 102 oriented downstream and arranged upstream of the terminal connecting flange 26 , so that the radially internal portion 92C of the second part 92 and the terminal connecting flange 26 of the internal wall 20 delimit between them a constriction 104 of the downstream portion 38B of the internal annular space 38 (FIG. 8).

In the illustrated embodiment, when the turbomachine is seen in axial section, the internal cooling element 52 thus forms, within the internal annular space 38, a succession of elbows 105 partially delimited by the terminal connecting flange 26 .

The succession of elbows 105 comprises for example a 180 degree elbow followed by a 90 degree elbow.

In a typical application example, the cooling elements 50 and 52 are made of a superalloy of nickel and chromium, and have a thickness of between 1 and 2 millimeters.

In order to guarantee satisfactory air passage sections as regards the downstream portions 36B, 38B of the annular spaces 36, 38 and as regards the air inlet orifices 56, 82 and air outlet 58B , 84B, the dimensions of the cooling elements 50 and 52 are determined when hot, that is to say taking into account the operating temperatures and the differential thermal expansion which result therefrom.

In operation, as shown in FIG. 7, the air of the external flow F2 is divided into a convective cooling flow F2B which circulates in the downstream portion 36B of the external annular space, along the corresponding terminal connection flange 24 , being accelerated by the narrowing of section induced by the external cooling element 50, and a flow F2C penetrating into the external annular chamber 54 by the air inlet orifices 52 of the external cooling element 50. This flow F2C is itself divided into an impact cooling flow F2D passing through the impact cooling holes 58A, and an bypass flow F2E which continues to flow downstream directly through the air outlet orifices 58B, without undergoing a considerable pressure drop while passing through the external cooling element 50.

Similarly, as shown in FIG. 8, the air from the internal flow F3 is divided into a convective cooling flow F3B which circulates in the downstream portion 38B of the internal annular space, along the terminal connecting flange 26 corresponding, being accelerated by the narrowing of section induced by the internal cooling element 52, and a flow F3C penetrating into the internal annular chamber 80 by the air inlet orifices 82 of the internal cooling element 52. This flow F3C is itself divided into an impact cooling flow F3D passing through the impact cooling holes 84A, and a bypass flow F3E which continues to flow downstream directly through the outlet orifices of air 84B, without undergoing a considerable pressure drop passing through the internal cooling element 52.

Claims (10)

1. Turbomachine (11) for aircraft, comprising: - an annular combustion chamber (10), comprising two coaxial walls of revolution, namely an external wall (18) and an internal wall (20), having respective first ends connected to each other by a wall of chamber bottom (22), and respective second opposite ends provided with respective terminal connecting flanges (24, 26), and - a first casing (30; 32) delimiting, between the latter and a first (18; 20) of the external and internal walls, an upstream portion (36A; 38A) of a first annular space (36; 38) allowing the passage of a first air flow (F2; F3) along and outside the annular combustion chamber (10), characterized in that it further comprises a first cooling element (50; 52) , which has a shape of revolution and extends from the first casing (30; 32) towards the annular combustion chamber so as to delimit, between the first cooling element (50; 52) and the connecting flange terminal (24; 26) of said first (18; 20) of the external and internal walls, a downstream portion (36B; 38B) of the first annular space, having an air passage section (S36B; S38B) smaller than an air passage section (S36A; S38A) of the upstream portion (36A; 36B) of the first annular space, the first cooling element ment (50; 52) delimiting, between the latter and the first casing (30; 32), a first annular chamber (54; 80), and the first annular chamber (54; 80) having air inlet ports (56; 82) and air outlet ports (58B; 84B) formed in the first cooling element. 2. The turbomachine according to claim 1, in which the first cooling element (50; 52) further comprises impact cooling holes (58A; 84A) configured to cool by impact of air jets the terminal connecting flange ( 24; 26) of said first (18; 20) of the outer and inner walls, the air outlet ports (58B; 84B) having respective air passage sections (S58B; S84B) larger than respective air passage sections (S58A; S84A) of the impact cooling holes (58A; 84A), and the air outlets (58B; 84B) being further away from the flange terminal link (24; 26) of said first (18; 20) of the outer and inner walls than are the impact cooling holes (58A; 84A). 3. The turbomachine according to claim 1 or 2, in which the air outlet orifices (58B; 84B) are configured so that the sum of the respective air passage sections (S58B; S84B) of these and of the air passage section (S36B; S38B) of the downstream portion (36B; 38B) of the first annular space is strictly less than the air passage section (S36A; S38A) of the upstream portion (36A; 38A) of the first annular space. 4. A turbomachine according to any one of claims 1 to 3, in which the first cooling element (50; 52) forms at least one elbow (75; 105) partially delimited by the terminal connecting flange (24; 26) of said first (18; 20) of the external and internal walls, within the first annular space (36; 38). 5. A turbomachine according to any one of claims 1 to 4, in which the first cooling element (50; 52) is a plate having a first part (60; 90) comprising the air inlet orifices (56; 82), a second part (62; 92) having the air outlet orifices (58B; 84B) and connected to the first part so that the first and second parts form a nose (72; 102). 6. A turbomachine according to claim 5, in which the first part (60; 90) of the first cooling element (50; 52) has a free end (60A; 90A) in pre-stressed support on the first casing (30; 32). 7. A turbomachine according to claim 5 or 6, in which the second part (62; 92) of the first cooling element (50; 52) is fixed to a fixing flange (66; 96) of the first casing (30; 32) . 8. A turbomachine according to any one of claims 5 to 7, in which the terminal connecting flange (24) of said first (18) of the external and internal walls is formed of a flange, and the nose (72) of the first cooling element (50) is oriented radially inwards and arranged upstream relative to the flange. 9. A turbomachine according to any one of claims 5 to 7, in which the terminal connecting flange (26) of said first (20) of the external and internal walls comprises a bent end portion (26A) convex in the direction of the upstream having a free end (26B) bearing radially on the first casing (32) or on another casing (40) extending the first casing (32), and the nose (102) of the first cooling element (52) is oriented downstream and arranged between said first (20) of the external and internal walls and the bent end part (26A). 10. Turbomachine according to any one of claims 1 to 9, further comprising: - a second casing (32; 30) delimiting, between the latter and a second (20; 18) of the external and internal walls, an upstream portion (38A; 36A) of a second annular space (38; 36) allowing the passage of a second air flow (F3; F2) along and outside the annular combustion chamber (10), and - a second cooling element (52; 50), which has a shape of revolution and extends from the second casing (32; 30) in the direction of the annular combustion chamber (10) so as to delimit, between the second cooling element (52; 50) and the terminal connecting flange (26; 24) of said second (20; 18) of the external and internal walls, a downstream portion (38B; 36B) of the second annular space (38; 36 ), having an air passage section (S38B; S36B) smaller than an air passage section (S38A; S36A) of the upstream portion (38A; 36A) of the second annular space (38; 36), and in which: the second cooling element (52; 50) delimits, between the latter and the second casing (32; 30), a second annular chamber (80; 54), and - The second annular chamber (80; 54) has air inlet orifices (82; 56) and air outlet orifices (84B; 58B) formed in the second cooling element (52; 50).