BE1029507B1 - STRUCTURE OF THREE-FLOW TURBOMACHINE - Google Patents

Abstract

The invention relates to a turbine engine (2) of the type with an unducted propeller (4), comprising: a splitter nozzle (10) splitting an air flow (F) into a primary flow (F1) and a flow secondary (F2); a compressor (14) compressing the primary stream (F1); and an air/oil heat exchanger (24); remarkable in that the exchanger (24) is positioned in a channel (26) traversed by a tertiary flow (F3), the tertiary flow (F3) being drawn from the secondary flow (F2) upstream of the exchanger (24) and meeting downstream of the exchanger (24) at least one annular row of rotor blades (20, 22) of the compressor (14).

Description

Description

STRUCTURE OF A THREE-FLOW TURBOMACHINE Field The invention relates to the design of a turbomachine and more particularly a turbomachine of the unducted propeller type. The invention relates to the arrangement of the heat exchanger intended for cooling the oil of the turbomachine. PRIOR ART In a ducted turbomachine (turbojet), it is known to arrange one or more heat exchanger(s) in the secondary flow, that is to say downstream of the fan. The lack of fairing around the secondary flow in turboprops makes this arrangement impossible and the exchangers are therefore supplied with cold air by a fan external to the engine. This solution is neither compact nor light, and reduces motor efficiency since part of the motor's energy is needed to drive this additional fan. Application WO 2020/084271 A1 describes a so-called fast propeller, or “open-rotor” motor (for example of the CROR “counter-rotating open rotor” or USF “unducted single fan” type). In this type of structure, the airflow velocity is too low at low aircraft speed for cooling to be sufficient, and the airflow velocity is too high at cruising speed, resulting in energy losses resulting from the drag of the exchangers. The known solution for turbojets is therefore not suitable for this type of engine, and a very large external fan is necessary to ensure the significant heat exchange needs.

In short, the known technical solutions take air directed towards the exchanger which is either air too hot to effectively cool the oil, or air having too high a speed for the aerodynamic or thrust losses to be negligible.

Added to this is the constraint linked to the fragility of the heat exchanger which cannot be placed directly upstream of the turbomachine, in particular because of the risk of impact with foreign bodies which may enter the turbomachine.

Summary of the invention Technical problem The invention aims to solve the drawbacks of the design/manufacture of turbomachines of the state of the art. In particular, the invention aims to propose a structure which allows effective cooling in a restricted space without hindering the performance of the turbomachine. Technical solution The invention relates to a turbine engine of the unducted propeller type, comprising: a splitter nozzle splitting an air flow into a primary flow and a secondary flow; a compressor compressing the primary stream; and an air/oil heat exchanger; noteworthy in that the exchanger is positioned in a channel traversed by a tertiary flow, the tertiary flow being drawn from the secondary flow upstream of the exchanger and meeting downstream of the exchanger at least one annular row of rotor blades of the compressor.

The turbomachine is, for example, in the form of a turboprop or an open-rotor turbomachine (for example of the CROR “counter-rotating open rotor” or USF “unducted single fan” type).

The airflow is generated by the propeller and/or by the movement of the aircraft on which the turbine engine is mounted.

… The unducted propeller can be arranged upstream of the splitter or downstream. The compressor can be a low pressure compressor, or “booster”.

The primary flow is compressed at least over part of its stroke by at least some compressor blades. The secondary flow does not “see” the compressor. The secondary flow is virtually unrestricted in size as there is no shroud surrounding it (radially outward). The tertiary flow is distinct from the secondary flow. The tertiary flow can converge with and/or diverge from the primary flow.

At least one of the rows of rotor blades of the compressor is located downstream of the exchanger, which, in other words, means that the compressor draws air from the secondary flow to drive it through the exchanger.

This so-called downstream part of the compressor is located in an annular zone where the primary flow and the tertiary flow converge.

This arrangement makes it possible to supply the heat exchanger with air that is cold enough and slow enough so that both the effectiveness of oil cooling is ensured and that the aerodynamic losses linked to the presence of the heat exchanger are limited.

According to an advantageous embodiment of the invention, the compressor comprises an upstream part with at least one annular row of blades and a downstream part with at least one annular row of blades, and only the downstream part of the compressor is traversed by air. from the tertiary flow from the exchanger.

This makes it possible to keep the primary flow, whose pressure and speed increase rapidly, distinct from the tertiary flow.

The exchanger does not "see" the upstream part of the compressor and thus retains its ability to cool.

According to an advantageous embodiment of the invention, the blades of the downstream part have a much greater radial height than the blades of the upstream part, preferably the blades of the downstream part are between 1.5 and 4 times higher radially than the blades of the upstream of the compressor.

The tertiary flow is thus not subjected to a significant speed gradient at the outlet of the exchanger, which would result in aerodynamic turbulence and therefore loss of efficiency in both cooling and engine performance.

According to an advantageous embodiment of the invention, the downstream part of the compressor that the air encounters downstream of the exchanger comprises a single annular row of rotor blades.

The overall compactness is improved.

These blades can be arranged in the form of a bladed wheel.

In addition, the downstream part can include one or two rows of stator vanes.

The number of blades in the downstream part can be established according to the size that can be afforded and/or according to the desired compression ratio to both suck in the tertiary flow at the desired speed and not hamper the compression of the primary flow, a key parameter of the efficiency of the turbomachine.

According to an advantageous embodiment of the invention, the air flow leaving the exchanger crosses all the annular rows of blades of the compressor.

According to an advantageous mode of the invention, an opening makes it possible to suck in part of the secondary flow to form the tertiary flow, the opening being non-bailing.

Thus, in a certain way, the opening can be delimited by an upstream edge and a downstream edge, the upstream and downstream edges having an identical radial position.

Alternatively, the upstream edge is farther from the axis of the turbomachine than the downstream edge, for example 1.1 times farther.

Alternatively, viewed in an axial section, the tangent to the air-guiding surface upstream of the opening describes a direction approaching the axis and the opening is contained between the axis and this tangent.

In all cases, this prevents the opening creating the tertiary flow from favoring the penetration of foreign bodies into the exchanger like a scoop.

According to an advantageous embodiment of the invention, the channel and/or the exchanger extend circumferentially over 360° around the axis of the turbomachine.

Alternatively, the exchanger and/or the channel does not extend over 360°. Independently of the angle described by the exchanger around the axis, the opening in the secondary flow can be partial (i.e. not describe 360°) and/or the mouth of the channel on the blades of the compressor can be partial (i.e. not describe 360°). According to an advantageous embodiment of the invention, a propeller is arranged upstream of the nozzle and/or blades straightening the secondary flow overlap axially the downstream part of the compressor.

Alternatively the propeller can be arranged downstream of the nozzle.

Alternatively, two propellers in the direction of rotation opposite to each other are arranged at an axial position downstream of the exchanger.

According to an advantageous embodiment of the invention, a bypass is arranged downstream of the downstream part of the compressor to divert part of the flow leaving the compressor towards the secondary flow.

The bypass is thus directly downstream of the convergence zone of the primary and tertiary flows (when they converge). This allows an acceleration of the flow to create thrust and compensate for the volume of air extracted in the tertiary flow upstream of the exchanger.

Engine efficiency is thus improved.

The bifurcation of the by-pass towards the secondary flow also makes it possible to evacuate the foreign bodies potentially sucked in upstream of the primary and tertiary flows, to prevent them from heading towards the high-pressure compressor or the combustion chamber downstream of the primary stream.

According to an advantageous embodiment of the invention, means are provided to isolate the primary flow from the tertiary flow, the latter flowing in the channel, through the downstream part of the compressor and then into the bypass.

Thus, the tertiary flow is completely isolated from the primary flow, limiting disturbances downstream of the exchanger.

The two streams are however not completely dynamically independent since they both “see” a respective part of the same row of rotor blades downstream of the exchanger.

According to an advantageous embodiment of the invention, at least the last annular row of blades of the downstream part has an intermediate circumferential ring.

This last row can be made of rotating or fixed blades.

The circumferential ring allows the flow to be guided towards the by-pass or towards the high-pressure compressor.

In connection with the paragraph above, the ring can also participate in the isolation between the tertiary flow and the primary flow.

The invention also relates to a method for cooling the oil of a turbine engine with an unducted propeller, comprising: the separation of an air flow into a primary flow and a secondary flow by means of a separation, the primary flow being compressed by one or more compressor(s) of the turbomachine and the secondary flow being external to the compressor(s); and the generation of a tertiary flow passing through an air/oil exchanger by means of at least one annular row of rotor blades of a compressor, the tertiary flow being drawn from the secondary flow.

The invention finally relates to a method of using the turbomachine according to one of the embodiments described above, comprising a propeller rotation step during which the primary flow and the secondary flow have a Mach number of 0.5, and the tertiary flow has a Mach number much lower than 0.3. Advantages of the invention The invention is particularly advantageous in that it makes it possible to circulate in the exchanger an air which is cold and at an adequate speed, thus ensuring cooling efficiency without hindering the efficiency of the engine or requiring additional means. bulky.

Good cooling efficiency allows the use of less imposing and therefore less bulky, less heavy and less expensive exchangers.

Description of the drawings Figure 1 shows a first embodiment of the invention;

Figure 2 illustrates a second embodiment of the invention; FIG. 3 represents a third embodiment of the invention; Figure 4 illustrates a fourth embodiment of the invention. Description of an embodiment In the description that follows, the terms "internal" and "external" refer to positioning relative to the axis of rotation of a turbomachine. The axial direction corresponds to the direction along the axis of rotation of the turbomachine. The radial direction is perpendicular to the axis of rotation. Upstream and downstream refer to the flow direction of a stream in the turbomachine.

The figures show the elements schematically and are not drawn to scale. In particular, certain dimensions are enlarged to facilitate reading of the figures.

FIG. 1 illustrates a turbomachine 2 according to a first variant. A propeller 4 attached to a hub 6 rotates around an axis 8.

The turbomachine 2 evolves in an air flow F whose movement relative to the turbomachine 2 is generated by the rotation of the propeller 4 and the advancement of the aircraft on which the turbomachine 2 is mounted.

In a variant not shown which may be similar to one of the examples described in document WO 2020/084271 A1, the propeller 4 is arranged in a downstream part of the turbomachine 2 and may optionally be supplemented by a second propeller having a direction of opposite rotation.

The air flow F is split into a primary flow F1 and a secondary flow F2 at the level of a splitter 10. The primary flow F1 encounters a rectifier inlet vane ("IGV") 11 and enters a vein 12 while the secondary flow F2 remains radially outside any fairing. The casing 13 delimits the stream 12 on the outside. Fixed shrouds and rotating wheel hubs delimit the stream 12 on the inside. The vein 12 is also crossed by structural arms (“struts”) (see 13.1 in figure 2) which take up the forces of the casing 13.

A compressor 14 is arranged to compress the primary flow F1. To do this, the compressor 14 is equipped with alternating rotor blades 16, 18, 20, 22 and stator vanes 17, 19, 21, arranged in annular rows around the axis 8.

FIG. 1 represents only the upstream part of the turbomachine. Downstream of the compressor 14, the primary flow F1 continues its path towards a second compressor, a combustion chamber and one or more turbines (not shown). The rotation of the turbine(s) causes the rotation of the hub 6, the propeller 4 and the rotor blades 16, 18, 20, 22.

The rotating elements are supported by bearings and the turbomachine may include a reduction gear between various rotating elements. The bearings and the reduction gear of the turbomachine 2 are lubricated by an oil which it is necessary to maintain within a given operating temperature range. An exchanger 24 is thus provided to cool the oil by causing the oil to travel through pipes which are cooled by a flow of air.

The exchanger 24 is arranged in a channel 26 in which circulates a so-called tertiary flow F3. The channel 26 can extend circumferentially over all or part of the turbomachine, that is to say 360° around the axis 8 or less. Similarly, the exchanger 24 can occupy all or part of the channel 26 and therefore extend over a large angular part around the axis 8 and in particular over 360°. The channel 26 has an opening 28 which leads to the space traversed by the secondary flow F2. The opening 28 is delimited by the fairing upstream by an upstream edge 30 and downstream by a downstream edge 32. The upstream 30 and downstream 32 edges are substantially at the same radial height to limit the fact that foreign elements present in the flow F are deflected towards the exchanger 24. The opening 28 is thus such that the speed of the flow F3 has a significant radial component when the latter is formed upstream of the channel 26. This allows in particular that the speed of the flow F3 when it crosses the exchanger is much smaller than the speed of the primary flow F1 in the stream 12. Thus, the geometry can be such that when the flow F1 is at a Mach number in an interval of 0.45 to 0.6 (usually 0.5), the flow F3 that the exchanger sees is much lower than a number of

Mach of 0.3. These values ensure efficient cooling at cruising speed. The opening 28 can receive a protective grid (not shown) or a valve opening or blocking the access of the secondary flow F2 to the channel 26.

The compressor 14 is composed of an upstream part 14.1 which includes the blades arranged in the stream 12 and a downstream part 14.2. The downstream part 14.2 comprises at least one rotor blade 20, 22 which generates the formation of the flow F3, creating a depression at the level of the opening 28.

Thus, after crossing the exchanger 24, the tertiary flow F3 joins the downstream part

14.2 of the compressor. In this first embodiment, the stream F3 converges with the primary stream F1.

The blades 20, 21, 22 of the downstream part 14.2 have a radial height H which is between 1.5 and 4 times the radial height h of the blades 16, 17, 18, 19 of the upstream part 14.1.

Figure 1 shows in dotted lines the possible position of an annular row of blades 34, fixed around the axis 8 and straightening the flow F2. The blades 34 as well as the blades of the propeller 4 can have a variable orientation (around the direction of their largest dimension).

The number of compressor blades 14 which form the downstream part 14.2, that is to say the number of blades seen by the tertiary flow F3 leaving the channel 26 can vary. Thus, in the embodiment illustrated in FIG. 2, the entire compressor is arranged in the downstream part 14.2, no blade of the compressor being placed in the section 12.

In comparison with FIG. 1, the stream 12 can for example be shorter axially and can comprise only one row of inlet vanes 11 and support arms 13.1.

In a third embodiment of the invention, a diversion channel or bypass 36 is arranged to divert part of the flow to the secondary flow F2. The bypass 36 is arranged downstream of the compressor 14 to divert part of the flow leaving the compressor to the secondary flow F2. Thus, a second separation nozzle 38 separates the flow which is seen by the downstream part 14.2 of the compressor into a primary flow F1 which continues its path towards a high pressure compressor and towards the combustion chamber, and a quaternary flow F4 which returns to the secondary stream F2. An outer fairing 40 delimits the channel 26 and the bypass 36 radially on the outside, and delimits the flow F2 on the inside. In addition to its function of heat exchange with the exchanger, the quaternary flow F4 therefore has here a thrust function complementary to the flow F2. Any losses of pressure in the flow F3 due to the interaction with the exchanger 24 are compensated by the blades of the downstream part 14.2 to restore sufficient pressure in the quaternary flow F4 before it finds the secondary flow F2. A compromise is thus obtained between a sufficiently low pressure to promote heat exchange in the channel 26 and a sufficiently high pressure in the by-pass 36 to promote thrust.

FIG. 4 illustrates a variant in which a single row of vanes occupies the downstream part 14.2 of the compressor 14. To structurally support the outer fairing 40 in the absence of stator vanes in the downstream part 14.2, structural arms 42 can be placed in bypass 36, and/or in channel 26 (not shown).

FIG. 4 also illustrates an aspect which may be applicable to the embodiment of FIG. 3, namely the isolation between the primary flow F1 and the tertiary flow F3. The quaternary F4 of FIG. 3 is therefore here the tertiary flow F3: each particle of air passing through the exchanger 24 continues its path in the by-pass 36. Thus, means are provided to isolate the primary flow from the tertiary flow , such as seals or rings.

FIG. 4 represents as such a circumferential ring 44 circumferentially connecting step by step the blades of the row of rotor blades of the downstream part 14.2. This ring is arranged radially at the level of the beak 38 and the fairing 13.

In a variant not shown, where the downstream part 14.2 comprises several rows of blades (as in FIG. 3), a ring of the same type can be placed on all the rows of blades. Alternatively, a ring can be provided only on the last row of blades of the downstream part 14.2 in order to prepare the separation of the flow between the primary flow led to the high pressure compressor and the bypass 36, thus only partially isolating the primary flow from the tertiary flow.

It should be noted that the invention is not limited to the examples described in the figures.

The teachings of the present invention are in particular also applicable to a turbine engine with a ducted propeller.

Each technical characteristic of each illustrated example is applicable to the other examples.

In particular, the number of blades in the downstream part, the presence or not of a by-pass, the position of the propeller or of the straightening blades, the presence of a circumferential ring, etc., can be drawn from one embodiment and be applied to another.

Claims (16)

Claims 1. Turbomachine (2) of the type with unducted propeller (4), comprising: - a splitter nozzle (10) splitting an air flow (F) into a primary flow (F1) and a secondary flow (F2) ; - a compressor (14) compressing the primary stream (F1); and - an air/oil heat exchanger (24); characterized in that the exchanger (24) is positioned in a channel (26) traversed by a tertiary flow (F3), the tertiary flow (F3) being drawn from the secondary flow (F2) upstream of the exchanger (24) and meeting downstream of the exchanger (24) at least one annular row of rotor blades (20, 22) of the compressor (14). 2. Turbomachine (2) according to claim 1, characterized in that the compressor (14) comprises an upstream part (14.1) with at least one annular row of blades (16-19) and a downstream part (14.2) with at least one at least one annular row of blades (20-22), and only the downstream part (14.2) of the compressor (14) is traversed by air from the tertiary flow (F3) coming from the exchanger (24). 3. Turbomachine (2) according to claim 2, characterized in that the blades (20-22) of the downstream part (14.2) have a radial height (H) much greater than the blades (16-19) of the upstream part ( 14.1), preferably the blades (20-22) of the downstream part (14.2) are between 1.5 and 4 times higher radially than the blades (16-19) of the upstream part (14.1) of the compressor (14). 4. Turbomachine (2) according to one of claims 1 to 3, characterized in that the downstream part (14.2) of the compressor (14) that meets the air downstream of the exchanger (24) comprises a single annular row rotor vanes (20). 5. Turbomachine (2) according to one of claims 1 to 4, characterized in that the downstream part (14.2) of the compressor (14) that the air encounters downstream of the exchanger (24) comprises one or two rows annular stator vanes (21). 6. Turbomachine (2) according to one of claims 1, 4 or 5, characterized in that the air flow (F3) leaving the exchanger (24) passes through all the annular rows of blades (18-22) compressor (14). 7. Turbomachine (2) according to one of claims 1 to 6, characterized in that an opening (28) makes it possible to suck in part of the secondary flow (F2) to form the tertiary flow (F3), the opening (28) being non-bailing. 8. Turbomachine (2) according to one of claims 1 to 7, characterized in that the channel (26) and the exchanger (24) extend circumferentially over 360° around the axis (8) of the turbomachine (2). 9. Turbomachine (2) according to one of claims 1 to 8, characterized in that a propeller (4) is arranged upstream of the nozzle (10). 10. Turbomachine (2) according to one of claims 1 to 9, characterized in that blades (34) straightening the secondary flow (F2) axially overlap the downstream part (14.2) of the compressor (14). 11. Turbomachine (2) according to one of claims 1 to 8 or 10, characterized in that two propellers in the direction of rotation opposite to each other are arranged at an axial position downstream of the exchanger (24 ). 12. Turbomachine (2) according to one of claims 1 to 11, characterized in that a bypass (36) is arranged downstream of the downstream part (14.2) of the compressor (14) to divert part of the flow leaving the compressor (14) to the secondary flow (F2). 13. Turbomachine (2) according to the preceding claim, characterized in that means are provided to isolate the primary stream (F1) from the tertiary stream (F3), the latter flowing in the channel (26) through the part downstream (14.2) of the compressor (14) then in the bypass (36). 14. Turbomachine (2) according to the preceding claim, characterized in that at least the last annular row of blades (22) of the downstream part (14.2) has an intermediate circumferential ring (44). 15. Process for cooling the oil of a turbine engine (2) with an unducted propeller (4), comprising: - the separation of an air flow (F) into a primary flow (F1) and a secondary flow (F2) by means of a separation nozzle (10), the primary flow (F1) being compressed by one or several compressor(s) (14) of the turbomachine (2) and the secondary flow (F2) being external to the compressor(s) (14); - the generation of a tertiary flow (F3) passing through an air/oil exchanger (24) by means of at least one annular row of rotor blades (20, 22) of a compressor (14), the tertiary flow ( F3) being drawn from the secondary stream (F2). 16. Method of using the turbomachine according to one of claims 1 to 14, comprising a step of rotation of the propeller during which the primary flow (F1) and the secondary flow (F2) have a Mach number of 0.5 , and the tertiary flow (F3) has a Mach number much lower than 0.3.