Multi-flow aircraft turbine engine

Abstract

A multi-flow turbine engine (10) for an aircraft, the turbine engine having a gas generator (12), the geometry of which is such that the gas generator has a cold compartment (C2) on its periphery that is larger than a hot compartment (C1) of the gas generator.

Description

Technical Field

[0001] The present invention relates to the field of multi-flow turbine engines, in particular to the field of multi-flow turbine engines for aircraft. Background Art

[0002] The prior art includes in particular documents FR3074476A1 and FR2981686A1, which relate in particular to turbine engines comprising at least one fan in the form of an unshrouded propeller. More particularly, these documents disclose turbine engines in which certain air flow ducts are suitable for improving the thermal management of the turbine engine and in particular for cooling equipment or compartments containing equipment.

[0003] like Figure 1 As shown, an aircraft turbine engine 10 generally comprises a gas generator 12 including at least one compressor 14 , an annular combustion chamber 16 and at least one turbine 18 .

[0004] The turbine engine 10 also includes at least one propeller 20 driven by the shaft of the turbine 18 .

[0005] The gas generator 12 comprises an annular splitter nose 22 for separating two annular ducts V1 , V2 for the respective flows of an inner primary flow F1 inside the gas generator 12 and an outer secondary flow F2 around the gas generator 12 .

[0006] When the propeller 20 is located upstream of the gas generator 12, as in the example shown, the propeller is called a fan, the purpose of which is to accelerate the air taken in. A portion of the air flow leaving the fan flows around the gas generator 12 to form a secondary flow F2, which generates most of the thrust generated by the turbine engine 10. The remaining air flow leaving the fan enters the gas generator 12 to form the main flow F1. The air is compressed in the compressor or each compressor 14, mixed with fuel and burned in the combustion chamber 16. The combustion gases then expand in the turbine or each turbine 18 to drive the rotor of the turbine in rotation, thereby driving the propeller 20 in rotation.

[0007] The propeller 20 comprises blades 24 which generally have an angular position, called pitch, which is fixed about an axis Y perpendicular to the longitudinal axis X of the turbine engine 10. Even in low engine speed phases such as idle, this pitch is configured so that the flow of the secondary flow F2 produced is sufficient to supply the fluid / air exchangers 26, whether they are surface-based or not, which are installed in the ducts for the flow of the secondary flow F2. These exchangers 26 are necessary for cooling the equipment of the turbine engine.

[0008] In order to improve engine performance (higher thrust and lower fuel consumption), some engine architectures aim to increase the bypass ratio (also known as BPR) by increasing the diameter of the propeller 20. In order to limit the mass of the turbine engine 10, the shroud 28 arranged around the propeller 20 can be removed and the blades 24 of the propeller 20 can be made steerable, thus having a variable pitch, in order to control the thrust level of the turbine engine.

[0009] However, in certain phases (such as ground idling, starting or thrust reversal phases), the flow rate of the secondary flow F2 is no longer high enough for thermal management of the turbine engine which requires heat exchange between fluid and air to ensure cooling of its equipment and its optimal operation.

[0010] In order to solve this problem, it has been proposed to provide an additional bypass flow dedicated to the thermal management of the turbine engine.

[0011] like Figure 2 As shown, the turbine engine 10 comprises a secondary propeller 30 driven in rotation by the shaft of the gas generator 10. This propeller 30 is located in the duct V1 for the flow of the main flow F1 and upstream of a second annular splitter nose 32 which separates two annular ducts V11, V12 for the respective flows of the first internal flow F11 and the second external flow F12.

[0012] The exchanger 26 is housed in the duct V12 and is fed by the second external flow F12 . As mentioned above, the first internal flow F11 feeds the compressor 14 .

[0013] Figure 2 The gas generator 12 is shown to comprise two annular compartments C1 , C2 extending around the axis X.

[0014] The first annular compartment C1 is called a hot compartment, since it is closer to the duct V11 and is therefore more exposed to the heat generated by the compression, combustion chamber 16 and combustion gases. This hot compartment C1 extends axially between the second splitter nose 32 and the first nozzle 34, which is used to discharge the first internal flow F11 from one or more turbines. This hot compartment C1 is also radially delimited by a first annular wall 36 and a second annular wall 38, which are respectively an inner annular wall and an outer annular wall, which extend coaxially around each other. The first wall 36 defines externally the duct V11 for the flow of the first internal flow F11, and the second wall 38 defines internally the duct V12 for the flow of the second external flow F12.

[0015] The second annular compartment C2 is called a cold compartment because it is less exposed to the heat and in particular because it is separated from the hot compartment by a duct V12. This cold compartment C2 extends axially between the first splitter nose 22 and the second nozzle 40 for discharging the second external flow F12. This cold compartment C1 is also radially delimited by a third annular wall 42 and a fourth annular wall 44, respectively an inner annular wall and an outer annular wall, which extend coaxially around each other. The third wall 42 externally delimits a duct V12 for the flow of the second external flow F12 and the fourth wall 44 internally delimits a duct V2 for the flow of the secondary flow F2.

[0016] Many devices of the turbine engine 10 are installed in the bulky hot compartment C1 . The cold compartment C2 is smaller and is reserved for devices that cannot withstand the temperature of the hot compartment C1 , in particular electronic devices.

[0017] This type of installation has several consequences:

[0018] The equipment in the thermal compartment C1 must be thermally protected. These thermal protections constitute additional volume and mass to be installed in an already very restricted environment, which tends to hamper accessibility for inspection and maintenance of the turbine engine.

[0019] - For some equipment (especially electronic equipment) that is more sensitive to heat and cannot be installed in the cold compartment C2 because the space is too limited, local ventilation must be added in the hot compartment C1.

[0020] - The service life of the device in the thermal compartment C1 is shortened and this restrictive environment has an impact on the design of the device: the choice of technology is limited and the mass of the device increases.

[0021] The present invention provides a solution to all or part of these problems. Summary of the invention

[0022] The invention relates to a multi-flow turbine engine for an aircraft, the turbine engine having a longitudinal axis and comprising:

[0023] - a gas generator comprising at least one compressor, an annular combustion chamber and at least one turbine,

[0024] - at least one main propeller, the at least one main propeller being driven in rotation by the shaft of the gas generator,

[0025] The gas generator includes:

[0026] a first annular splitter nose for separating two annular ducts for the respective flows of an inner primary flow inside the gas generator and an outer secondary flow around the gas generator,

[0027] a secondary propeller driven in rotation by the shaft of the gas generator, the propeller being located in the duct for the flow of the primary flow and upstream of the second annular splitter nose which separates the two annular ducts for the respective flows of the first inner flow and the second outer flow,

[0028] - a first nozzle for discharging a first internal flow,

[0029] - a second nozzle for discharging a second external flow,

[0030] a first annular compartment, referred to as the hot compartment, extending axially between the second diverter nose and the first nozzle and between a first annular wall and a second annular wall, the first annular wall and the second annular wall being respectively an inner annular wall and an outer annular wall, the first annular wall and the second annular wall extending coaxially around each other, the first wall defining externally said duct for the flow of the first inner flow and the second wall defining internally said duct for the flow of the second outer flow,

[0031] - a second annular compartment, referred to as the cold compartment, extending axially between the first diverter nose and the second nozzle and between a third annular wall and a fourth annular wall, the third annular wall and the fourth annular wall being respectively an inner annular wall and an outer annular wall, the third annular wall and the fourth annular wall extending coaxially around each other, the third wall externally defining said duct for the flow of the second outer flow, the fourth wall internally defining said duct for the flow of the secondary flow,

[0032] It is characterized in that

[0033] -R 最大壁42 Between 0.8×R 最大30 Up to 1.5×R 最大30 between,

[0034] -R 最大壁44 Greater than or equal to 1.3×R 喷嘴40 ,

[0035] -R 喷嘴40 Between 1.2×R 喷嘴34 Up to 1.5×R 喷嘴34 between, and

[0036] -V 冷隔室 Greater than or equal to 1.3×V 热隔室 ,

[0037] in:

[0038] R 最大壁42 is the maximum radius of the third wall,

[0039] R 最大30 is the maximum radius of the secondary propeller,

[0040] R 最大壁44 is the maximum radius of the fourth wall,

[0041] R 喷嘴40 is the radius of the second nozzle,

[0042] R 喷嘴34 is the radius of the first nozzle,

[0043] V 冷隔室 is the volume of the cold compartment, and

[0044] V 热隔室 is the volume of the hot compartment.

[0045] In this application, the terms "upstream" and "downstream" refer to the direction of gas flow in a turbine engine.

[0046] The radius is measured relative to the longitudinal axis of the turbine engine.

[0047] The invention therefore proposes to optimize the geometrical features of the turbine engine, in particular the geometrical features of the gas generator, to optimize the dimensions of the hot and cold compartments of the gas generator and to facilitate the integration of equipment (including heat-sensitive equipment) in these compartments. Advantageously, the cold compartment is longer and has a larger diameter than the cold compartments of the prior art. Moreover, in the event of a fire, the cold compartment is protected by a duct for the flow of a second external flow, which also protects the cold compartment from the effects of the thermal radiation associated with the combustion.

[0048] The turbine engine according to the invention may include one or more of the following features, considered independently of one another or in combination with one another:

[0049] -R 最小壁42 Between 0.8×R 最大30 Up to 1.5×R 最大30 between,

[0050] Among them, R 最小壁42 is the minimum radius of the third wall;

[0051] -R 喷嘴40 Between 1.95×R 最大壁42 To 1.05×R 最大壁42 between;

[0052] - said at least one main propeller is located upstream of the gas generator;

[0053] - said at least one main propeller is located downstream of the gas generator;

[0054] - the turbine engine comprises two counter-rotating main propellers located downstream of the gas generator;

[0055] - The at least one main propeller is unshrouded. BRIEF DESCRIPTION OF THE DRAWINGS

[0056] Other features and advantages will emerge from the following description of non-limiting embodiments of the invention, with reference to the accompanying drawings, in which:

[0057] [ Figure 1 ] Figure 1 is a schematic axial cross-sectional view of a twin-flow turbine engine for an aircraft,

[0058] [ Figure 2 ] Figure 2 is a schematic half view of an axial section of a twin-flow turbine engine for an aircraft, which provides an additional bypass air flow for equipment cooling,

[0059] [ Figure 3 ] Figure 3 is with Figure 2 Similar to and illustrating the prior art view of the present invention,

[0060] [ Figure 4 ] Figure 4 is with Figure 2 similar views showing embodiments of the present invention, and

[0061] [ Figure 5 ] Figure 5 is with Figure 2 Similar views illustrate alternative embodiments of the present invention. DETAILED DESCRIPTION

[0062] Figure 1 and Figure 2 Already described above.

[0063] Figure 3 An aircraft turbine engine 10 in the prior art of the present invention is shown, which is similar to Figure 2 A turbine engine 10.

[0064] In this application, we define Figure 3 The parameters are:

[0065] R 最小壁42 : the minimum radius of the third wall 42,

[0066] R最大壁42 : the maximum radius of the third wall 42,

[0067] R 最大30 : Maximum radius of the secondary propeller 30,

[0068] R 最大壁44 : the maximum radius of the fourth wall 44,

[0069] R 喷嘴40 : The radius of the second nozzle 40,

[0070] R 喷嘴34 : The radius of the first nozzle 34,

[0071] V 冷隔室 : the volume of the cold compartment C2, and

[0072] V 热隔室 : Volume of thermal compartment C1.

[0073] The radius is measured relative to the longitudinal axis X of the turbine engine 10 .

[0074] exist Figure 3 In the prior art shown, the turbine engine 10 is as follows:

[0075] -R 最大壁42 Greater than or equal to 1.6×R 最大30 ,

[0076] -R 喷嘴44 Greater than or equal to 1.6×R 喷嘴34 ,as well as

[0077] -V 冷隔室 Less than V 热隔室 .

[0078] Figure 4 An embodiment of a turbine engine 10 according to the invention is shown.

[0079] The foregoing description made with reference to the foregoing drawings applies to the following unless it is inconsistent with the following: Figure 4 .

[0080] Figure 4 The turbine engine 10 is different from the prior art in that:

[0081] -R 最大壁42 Between 0.8×R 最大30 Up to 1.5×R 最大30 between,

[0082] -R 最大壁44 Greater than or equal to 1.3×R 喷嘴40 ,

[0083] -R喷嘴40 Between 1.2×R 喷嘴34 Up to 1.5×R 喷嘴34 between, and

[0084] -V 冷隔室 Greater than or equal to 1.3×V 热隔室 .

[0085] Furthermore, preferably:

[0086] -R 最小壁42 Between 0.8×R 最大30 Up to 1.5×R 最大30 between, and / or

[0087] -R 喷嘴40 Between 1.95×R 最大壁42 To 1.05×R 最大壁42 ie the radius of the nozzle 40 is substantially equal to the maximum radius of the wall 42 (+ / - 5%).

[0088] like Figure 4 As shown, the propeller 20 may be located upstream of the gas generator 12 .

[0089] exist Figure 5 In the embodiment shown, the propeller 20 is located downstream of the gas generator 12. As shown, the turbine engine 10 may comprise a twin propeller downstream of the gas generator 12, ie, for example, two main propellers 20 that rotate in series and in opposite directions.

[0090] Figure 5 The turbine engine 10 has Figure 4 The same geometrical features (R 最大壁42 , R 最大壁44 , R 喷嘴40 、V 冷隔室 wait).

[0091] An advantage of the invention is that a larger cold compartment C2 is obtained, so as to be able to accommodate more equipment, whatever the size of the turbine engine 10 .

[0092] Another advantage is that the cold compartment C2 is protected from fire and thermal radiation by the duct V12 for the flow F12 , which makes it possible to increase the lifetime of the equipment installed in this compartment C2 while optimizing its design without very high temperature resistance constraints.

[0093] Furthermore, from an aerodynamic point of view, the diameter of the duct V12 is relatively small compared to the prior art, and for the same channel cross section, the radial dimension of the duct is larger than that of the prior art, which facilitates the integration of the heat exchanger.

[0094] Furthermore, it is easier to control the Mach in the duct V12 (the Mach in the duct is directly linked to the channel cross section). In fact, for ducts of larger diameter, very small deviations in the radial dimension of the duct have a strong influence on the channel cross section.

Claims

1. A multi-flow turbine engine (10) for an aircraft, the turbine engine having a longitudinal axis (X) and comprising: a gas generator (12) comprising at least one compressor (14), an annular combustion chamber (16) and at least one turbine (18), - at least one main propeller (20) driven in rotation by the shaft of the gas generator (12), The gas generator (12) comprises: a first annular splitter nose (22) for separating two annular ducts (V1, V2) for the respective flows of an inner main flow (F1) inside the gas generator (12) and an outer secondary flow (F2) around the gas generator (12), a secondary propeller (30) driven in rotation by the shaft of the gas generator (12), the propeller (30) being located in the duct (V1) for the flow of the primary flow (F1) and upstream of a second annular splitter nose (32) separating two annular ducts (V11, V12) for the respective flows of a first internal flow (F11) and a second external flow (F12), - a first nozzle (34) for discharging said first internal flow (F11), - a second nozzle (40) for discharging said second external flow (F12), a first annular compartment (C1), referred to as the hot compartment, extending axially between the second splitter nose (32) and the first nozzle (34) and between a first annular wall and a second annular wall (36, 38), the first annular wall and the second annular wall being respectively an inner annular wall and an outer annular wall, the first annular wall and the second annular wall extending coaxially around each other, the first wall (36) defining externally the duct (V11) for the flow of the first inner flow (F11) and the second wall (38) defining internally the duct (V12) for the flow of the second outer flow (F12), a second annular compartment (C2), referred to as the cold compartment, extending axially between the first diverter nose (22) and the second nozzle (40) and between a third annular wall and a fourth annular wall (42, 44), the third annular wall and the fourth annular wall being respectively an inner annular wall and an outer annular wall, the third annular wall and the fourth annular wall extending coaxially around each other, the third wall (42) defining externally the duct (V12) for the flow of the second external flow (F12), the fourth wall (44) defining internally the duct (V2) for the flow of the secondary flow (F2), It is characterized in that -R 最大壁42 Between 0.8×R 最大30 Up to 1.5×R 最大30 between, -R 最大壁44 Greater than or equal to 1.3×R 喷嘴40 , -R 喷嘴40 Between 1.2×R 喷嘴34 Up to 1.5×R 喷嘴34 between, and -V 冷隔室 Greater than or equal to 1.3×V 热隔室 , in: R 最大壁42 is the maximum radius of the third wall (42), R 最大30 is the maximum radius of the secondary propeller (30), R 最大壁44 is the maximum radius of the fourth wall (44), R 喷嘴40 is the radius of the second nozzle (40), R 喷嘴34 is the radius of the first nozzle (34), V 冷隔室 is the volume of the cold compartment (C2), and V 热隔室 is the volume of the thermal compartment (C1).

2. The turbine engine (10) according to claim 1, wherein: -R 最小壁42 Between 0.8×R 最大30 Up to 1.5×R 最大30 between, in: R 最小壁42 is the minimum radius of the third wall (42).

3. The turbine engine (10) according to claim 1 or 2, wherein: -R 喷嘴40 Between 1.95×R 最大壁42 To 1.05×R 最大壁42 between.

4. The turbine engine (10) according to any one of claims 1 to 3, wherein: The at least one main propeller (20) is located upstream of the gas generator (12).

5. The turbine engine (10) according to any one of claims 1 to 3, wherein: The at least one main propeller (20) is located downstream of the gas generator (12).

6. Turbine engine (10) according to the preceding claim, comprising two counter-rotating main propellers (20) situated downstream of the gas generator (12).

7. A turbine engine (10) according to any one of the preceding claims, wherein: The at least one main propeller (20) is unshrouded.

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