EP4374056A1 - System for cooling oil in an aircraft turbine engine - Google Patents

System for cooling oil in an aircraft turbine engine

Info

Publication number
EP4374056A1
EP4374056A1 EP21752650.8A EP21752650A EP4374056A1 EP 4374056 A1 EP4374056 A1 EP 4374056A1 EP 21752650 A EP21752650 A EP 21752650A EP 4374056 A1 EP4374056 A1 EP 4374056A1
Authority
EP
European Patent Office
Prior art keywords
heat exchanger
turbine engine
stream
pressure compressor
aircraft
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
EP21752650.8A
Other languages
German (de)
French (fr)
Inventor
Xavier Vandenplas
Bruno Servais
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Safran Aero Boosters SA
General Electric Co
Original Assignee
Safran Aero Boosters SA
General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Safran Aero Boosters SA, General Electric Co filed Critical Safran Aero Boosters SA
Publication of EP4374056A1 publication Critical patent/EP4374056A1/en
Pending legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/14—Cooling of plants of fluids in the plant, e.g. lubricant or fuel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/213—Heat transfer, e.g. cooling by the provision of a heat exchanger within the cooling circuit
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to a system for cooling oil in an aircraft turbine engine.
  • An object of the invention is to improve the integration of a heat exchanger for cooling the oil in an aircraft turbomachine.
  • the invention proposes a system for cooling oil in an aircraft turbomachine, and comprising an intermediate support casing intended to be located between a low-pressure compressor and a high-pressure compressor of the turbomachine. aircraft, and a heat exchanger intended to cool the oil by heat exchange with air; the heat exchanger being at least partially integrated into the intermediate support casing.
  • the turbomachine can be double-flow, in which case it comprises a primary stream and a secondary stream.
  • the turbomachine can be triple flow, in which case it comprises a primary stream, a secondary stream and a tertiary stream.
  • it may comprise one or more of the following characteristics, taken in isolation or according to all possible technical combinations:
  • the heat exchanger comprises a primary stream surface configured to be in a primary stream of the aircraft turbomachine
  • the primary stream surface is configured to be between the most downstream vane of the low pressure compressor, and the most upstream vane of the high pressure compressor;
  • the heat exchanger is configured to partially obstruct the primary stream
  • the heat exchanger comprises a secondary stream surface configured to be in a secondary stream of the aircraft turbomachine
  • the heat exchanger is configured to extend radially between the primary stream and a secondary stream, and includes a secondary stream surface configured to be in a secondary stream of the aircraft turbomachine;
  • the heat exchanger is configured to partially obstruct the secondary vein
  • the heat exchanger comprises a tertiary stream surface configured to be in a tertiary stream of the aircraft turbomachine;
  • the heat exchanger is configured to partially obstruct the tertiary vein
  • the heat exchanger is annular and is configured to extend around an axis of the aircraft turbine engine
  • the heat exchanger comprises fins configured to extend radially and parallel to the axis of the aircraft turbomachine; • the heat exchanger is a part fixed to the intermediate support casing, or the heat exchanger and the intermediate support casing are of the same part.
  • the invention further provides an aircraft turbine engine comprising such a system, a low pressure compressor and a high pressure compressor, the intermediate support casing being located between the low pressure compressor and the high pressure compressor.
  • the invention further proposes an aircraft comprising such a turbomachine.
  • FIG. 1 is a sectional view, along the axis, of an aircraft turbine engine comprising an example of a system according to the invention.
  • FIG. 2 is a sectional view, along the axis, of another aircraft turbine engine comprising another example of a system according to the invention.
  • FIG. 1 illustrates an example of an aircraft turbomachine 100 comprising a system 1 for cooling oil according to one embodiment of the invention.
  • the aircraft turbomachine 100 is a dual-flow axial turbomachine comprising successively along the engine axis X, a fan 110, a low pressure compressor 120, a high pressure compressor 130, a combustion chamber 160, a high pressure turbine 140 and a low pressure turbine 150.
  • the fan 110 makes it possible to generate a primary air flow in a primary stream 106 and a secondary air flow in a secondary stream 107.
  • the aircraft turbomachine 100 comprises an inlet support casing 181 located downstream of the fan 110.
  • the inlet support casing 181 is provided with an annular sleeve defining the primary stream 106 and arms 183 structural elements that extend radially inward across primary vein 106.
  • the aircraft turbomachine 100 comprises an intermediate support casing 2 between the low 120 and high 130 pressure compressors.
  • the intermediate support casing 2 comprises an annular sleeve preferably having a gooseneck profile and delimiting the primary stream 106 between the low 120 and high 130 pressure compressors. It is also provided with structural arms 184 extending radially through the primary vein 106.
  • Figure 2 schematically illustrates the respective positions of the low 120 and high 130 pressure compressors, of the intermediate support casing 2, of the primary stream 106, and of a tertiary stream 108 in the case of a triple turbomachine flux.
  • the system 1 comprises an intermediate support casing 2, for example as shown in Figure 1 or in Figure 2, and a heat exchanger 3 at least partially integrated into the casing of intermediate support 2.
  • the heat exchanger 3 has at least one surface in the primary stream 106 and/or in the secondary stream 107 (FIG. 1) and/or in the tertiary stream 108 (FIG. 2).
  • the heat exchanger 3 comprises a fluid inlet allowing oil to enter therein and a fluid outlet allowing oil to exit therefrom.
  • the heat exchanger 3 can be annular around the axis X of the turbomachine.
  • the heat exchanger 3 can be a part attached to the intermediate support casing 2 or be integral with the intermediate support casing 2.
  • the heat exchanger 3 can be axially at the same level as the structural arms 184.
  • the heat exchanger 3 passes radially through the intermediate support casing 2 and has a primary vein surface 31, radially internal, in the primary vein 106, and a vein surface secondary 32, radially outer, in the secondary stream 107.
  • the primary stream surface 31 is downstream of the blade furthest downstream of the low-pressure compressor 120, and upstream of the blade furthest upstream of the high pressure compressor 130.
  • the heat exchanger 3 has a tertiary stream surface 33 in the tertiary stream 108. It partially obstructs the tertiary stream 108. It comprises fins 34 extending radially and parallel to the axis of the turbomachine.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

The present invention relates to a system (1) for cooling oil in an aircraft turbine engine (100), said system comprising an intermediate support casing (2) intended to be located between a low-pressure compressor (120) and a high-pressure compressor (130) of the aircraft turbine engine (100), and a heat exchanger (3) for cooling the oil by heat exchange with air; wherein the heat exchanger (3) is at least partially integrated into the intermediate support casing (2).

Description

Système pour refroidir de l’huile dans une turbomachine d’aéronef Domaine technique System for cooling oil in an aircraft turbomachine Technical field
[0001] La présente invention concerne un système pour refroidir de l’huile dans une turbomachine d’aéronef. The present invention relates to a system for cooling oil in an aircraft turbine engine.
Art antérieur Prior art
[0002] L’augmentation des performances des turbomachines d’aéronef conduit à une augmentation des réjections thermiques vers l’huile. Un échangeur de chaleur permettant de refroidir l'huile est connu, notamment du document BE2017/5735. [0002] The increase in the performance of aircraft turbomachines leads to an increase in thermal rejections towards the oil. A heat exchanger for cooling the oil is known, in particular from document BE2017/5735.
Résumé de l’invention Summary of the invention
[0003] Un objet de l’invention est d’améliorer l’intégration d’un échangeur de chaleur permettant de refroidir l'huile dans une turbomachine d’aéronef. [0004] A cet effet, l’invention propose un système pour refroidir de l’huile dans une turbomachine d’aéronef, et comprenant un carter de support intermédiaire destiné à être localisé entre un compresseur basse pression et un compresseur haute pression de la turbomachine d’aéronef, et un échangeur de chaleur destiné à refroidir l’huile par échange thermique avec de l’air ; l’échangeur de chaleur étant au moins partiellement intégré dans le carter de support intermédiaire. An object of the invention is to improve the integration of a heat exchanger for cooling the oil in an aircraft turbomachine. [0004] To this end, the invention proposes a system for cooling oil in an aircraft turbomachine, and comprising an intermediate support casing intended to be located between a low-pressure compressor and a high-pressure compressor of the turbomachine. aircraft, and a heat exchanger intended to cool the oil by heat exchange with air; the heat exchanger being at least partially integrated into the intermediate support casing.
[0005] L’intégration, au moins partielle, de l’échangeur de chaleur dans le carter de support intermédiaire permet d’obtenir une turbomachine plus compacte et plus légère. Elle diminue aussi le coût de production de la turbomachine. [0005] The at least partial integration of the heat exchanger in the intermediate support casing makes it possible to obtain a more compact and lighter turbomachine. It also reduces the production cost of the turbomachine.
[0006] La turbomachine peut être double-flux, auquel cas elle comprend une veine primaire et une veine secondaire. La turbomachine peut être triple flux, auquel cas elle comprend une veine primaire, une veine secondaire et une veine tertiaire. [0007] Selon des modes avantageux de l’invention, elle peut comprendre une ou plusieurs des caractéristiques suivantes, prises isolément ou selon toutes les combinaisons techniques possibles : [0006] The turbomachine can be double-flow, in which case it comprises a primary stream and a secondary stream. The turbomachine can be triple flow, in which case it comprises a primary stream, a secondary stream and a tertiary stream. [0007] According to advantageous embodiments of the invention, it may comprise one or more of the following characteristics, taken in isolation or according to all possible technical combinations:
• l’échangeur de chaleur comprend une surface de veine primaire configurée pour être dans une veine primaire de la turbomachine d’aéronef ; • the heat exchanger comprises a primary stream surface configured to be in a primary stream of the aircraft turbomachine;
• la surface de veine primaire est configurée pour être entre l’aube la plus en aval du compresseur basse pression, et l’aube la plus en amont du compresseur haute pression ; • the primary stream surface is configured to be between the most downstream vane of the low pressure compressor, and the most upstream vane of the high pressure compressor;
• l’échangeur de chaleur est configuré pour obstruer partiellement la veine primaire ; • the heat exchanger is configured to partially obstruct the primary stream;
• l’échangeur de chaleur comprend une surface de veine secondaire configurée pour être dans une veine secondaire de la turbomachine d’aéronef ; • the heat exchanger comprises a secondary stream surface configured to be in a secondary stream of the aircraft turbomachine;
• l’échangeur de chaleur est configuré pour s’étendre radialement entre la veine primaire et une veine secondaire, et comprend une surface de veine secondaire configurée pour être dans une veine secondaire de la turbomachine d’aéronef ; • the heat exchanger is configured to extend radially between the primary stream and a secondary stream, and includes a secondary stream surface configured to be in a secondary stream of the aircraft turbomachine;
• l’échangeur de chaleur est configuré pour obstruer partiellement la veine secondaire ; • the heat exchanger is configured to partially obstruct the secondary vein;
• l’échangeur de chaleur comprend une surface de veine tertiaire configurée pour être dans une veine tertiaire de la turbomachine d’aéronef ; • the heat exchanger comprises a tertiary stream surface configured to be in a tertiary stream of the aircraft turbomachine;
• l’échangeur de chaleur est configuré pour obstruer partiellement la veine tertiaire ; • the heat exchanger is configured to partially obstruct the tertiary vein;
• l’échangeur de chaleur est annulaire et est configuré pour s’étendre autour d’un axe de la turbomachine d’aéronef ; • the heat exchanger is annular and is configured to extend around an axis of the aircraft turbine engine;
• l’échangeur de chaleur comprend des ailettes configurées pour s’étendre radialement et parallèlement à l’axe de la turbomachine d’aéronef ; • l’échangeur de chaleur est une pièce fixée au carter de support intermédiaire, ou l’échangeur de chaleur et le carter de support intermédiaire sont d’une même pièce. • the heat exchanger comprises fins configured to extend radially and parallel to the axis of the aircraft turbomachine; • the heat exchanger is a part fixed to the intermediate support casing, or the heat exchanger and the intermediate support casing are of the same part.
[0008] L’invention propose en outre une turbomachine d’aéronef comprenant un tel système, un compresseur basse pression et un compresseur haute pression, le carter de support intermédiaire étant situé entre le compresseur basse pression et le compresseur haute pression. [0009] L’invention propose en outre un aéronef comprenant une telle turbomachine. The invention further provides an aircraft turbine engine comprising such a system, a low pressure compressor and a high pressure compressor, the intermediate support casing being located between the low pressure compressor and the high pressure compressor. [0009] The invention further proposes an aircraft comprising such a turbomachine.
Brève description des figures Brief description of figures
[0010] D'autres caractéristiques et avantages de l'invention apparaîtront à la lecture de la description détaillée qui suit pour la compréhension de laquelle on se reportera aux figures annexées parmi lesquelles : Other characteristics and advantages of the invention will appear on reading the detailed description which follows for the understanding of which reference will be made to the appended figures, among which:
- la figure 1 est une vue en coupe, le long de l’axe, d’une turbomachine d’aéronef comprenant un exemple de système selon l’invention ; et - Figure 1 is a sectional view, along the axis, of an aircraft turbine engine comprising an example of a system according to the invention; and
- la figure 2 est une vue en coupe, le long de l’axe, d’une autre turbomachine d’aéronef comprenant un autre exemple de système selon l’invention. - Figure 2 is a sectional view, along the axis, of another aircraft turbine engine comprising another example of a system according to the invention.
Modes de réalisation de l’invention Embodiments of the Invention
[0011] Cette partie du texte décrit en détails des modes de réalisation préférés de l’invention. Des références à des figures sont utilisées mais l’invention n’est pas limitée par celles-ci. Les dessins et/ou figures décrits ci- dessous ne sont que schématiques et ne sont pas limitants. This part of the text describes in detail preferred embodiments of the invention. References to figures are used but the invention is not limited by them. The drawings and/or figures described below are only schematic and are not limiting.
[0012] Dans le cadre du présent document, il est fait référence aux directions « axiale », « circonférentielle » et « radiale » correspondant de respectivement en des directions parallèles à l’axe moteur, essentiellement circulaire autour de l’axe moteur, et perpendiculaire à l’axe moteur. Des repères sur les figures illustrent ces directions (munies d’un sens) notées respectivement X, R et C. Les termes « intérieurement » et « vers l’intérieur » correspondent naturellement à un sens vers l’axe moteur X selon une direction radiale, et les termes « extérieurement » et « vers l’extérieur » au sens opposé selon cette direction. In the context of this document, reference is made to the "axial", "circumferential" and "radial" directions corresponding respectively to directions parallel to the motor shaft, essentially circular around the motor shaft, and perpendicular to the motor axis. Markers in the figures illustrate these directions (provided with a direction) denoted respectively X, R and C. The terms "internally" and "inwards" naturally correspond to a direction towards the motor axis X according to a radial direction, and the terms "externally" and "outwardly" in the opposite direction according to this direction.
[0013] La figure 1 illustre un exemple de turbomachine d’aéronef 100 comprenant un système 1 pour refroidir de l’huile selon un mode de réalisation de l’invention. La turbomachine d’aéronef 100 est une turbomachine axiale à double flux comprenant successivement le long de l’axe moteur X, une soufflante 110, un compresseur basse pression 120, un compresseur haute pression 130, une chambre de combustion 160, une turbine haute pression 140 et une turbine basse pression 150. La soufflante 110 permet de générer un flux d’air primaire dans une veine primaire 106 et un flux d’air secondaire dans une veine secondaire 107. [0013] FIG. 1 illustrates an example of an aircraft turbomachine 100 comprising a system 1 for cooling oil according to one embodiment of the invention. The aircraft turbomachine 100 is a dual-flow axial turbomachine comprising successively along the engine axis X, a fan 110, a low pressure compressor 120, a high pressure compressor 130, a combustion chamber 160, a high pressure turbine 140 and a low pressure turbine 150. The fan 110 makes it possible to generate a primary air flow in a primary stream 106 and a secondary air flow in a secondary stream 107.
[0014] La turbomachine d’aéronef 100 comprend un carter de support d’entrée 181 situé en aval de la soufflante 110. Le carter de support d’entrée 181 est muni d’une manche annulaire délimitant la veine primaire 106 et de bras 183 structuraux qui s’étendent radialement vers l’intérieur en traversant la veine primaire 106. [0014] The aircraft turbomachine 100 comprises an inlet support casing 181 located downstream of the fan 110. The inlet support casing 181 is provided with an annular sleeve defining the primary stream 106 and arms 183 structural elements that extend radially inward across primary vein 106.
[0015] La turbomachine d’aéronef 100 comprend un carter de support intermédiaire 2 entre les compresseurs basse 120 et haute 130 pression. Le carter de support intermédiaire 2 comprend une manche annulaire présentant de préférence un profil en col de cygne et délimitant la veine primaire 106 entre les compresseurs basse 120 et haute 130 pression. Il est aussi muni de bras 184 structuraux s’étendant radialement à travers la veine primaire 106. The aircraft turbomachine 100 comprises an intermediate support casing 2 between the low 120 and high 130 pressure compressors. The intermediate support casing 2 comprises an annular sleeve preferably having a gooseneck profile and delimiting the primary stream 106 between the low 120 and high 130 pressure compressors. It is also provided with structural arms 184 extending radially through the primary vein 106.
[0016] La figure 2 illustre, schématiquement, les positions respectives des compresseurs basse 120 et haute 130 pression, du carter de support intermédiaire 2, de la veine primaire 106, et d’une veine tertiaire 108 dans le cas d’une turbomachine triple flux. [0016] Figure 2 schematically illustrates the respective positions of the low 120 and high 130 pressure compressors, of the intermediate support casing 2, of the primary stream 106, and of a tertiary stream 108 in the case of a triple turbomachine flux.
[0017] Le système 1 selon l’invention comprend un carter de support intermédiaire 2, par exemple tel qu’illustré à la figure 1 ou à la figure 2, et un échangeur de chaleur 3 au moins partiellement intégré dans le carter de support intermédiaire 2. L’échangeur de chaleur 3 a au moins une surface dans la veine primaire 106 et/ou dans la veine secondaire 107 (figure 1 ) et/ou dans la veine tertiaire 108 (figure 2). L’échangeur de chaleur 3 comprend une entrée fluidique permettant d’y faire entrer de l’huile et une sortie fluidique permettant d’en faire sortie de l’huile. The system 1 according to the invention comprises an intermediate support casing 2, for example as shown in Figure 1 or in Figure 2, and a heat exchanger 3 at least partially integrated into the casing of intermediate support 2. The heat exchanger 3 has at least one surface in the primary stream 106 and/or in the secondary stream 107 (FIG. 1) and/or in the tertiary stream 108 (FIG. 2). The heat exchanger 3 comprises a fluid inlet allowing oil to enter therein and a fluid outlet allowing oil to exit therefrom.
[0018] L’échangeur de chaleur 3 peut être annulaire autour de l’axe X de la turbomachine. The heat exchanger 3 can be annular around the axis X of the turbomachine.
[0019] L’échangeur de chaleur 3 peut être une pièce rapportée au carter de support intermédiaire 2 ou être monobloc avec le carter de support intermédiaire 2. [0019] The heat exchanger 3 can be a part attached to the intermediate support casing 2 or be integral with the intermediate support casing 2.
[0020] L’échangeur de chaleur 3 peut être axialement au même niveau que les bras 184 structuraux. The heat exchanger 3 can be axially at the same level as the structural arms 184.
[0021] Dans l’exemple illustré à la figure 1 , l’échangeur de chaleur 3 traverse radialement le carter de support intermédiaire 2 et a une surface de veine primaire 31, radialement interne, dans la veine primaire 106, et une surface de veine secondaire 32, radialement externe, dans la veine secondaire 107. [0022] La surface de veine primaire 31 est en aval de l’aube la plus en aval du compresseur basse pression 120, et en amont de l’aube la plus en amont du compresseur haute pression 130. In the example illustrated in Figure 1, the heat exchanger 3 passes radially through the intermediate support casing 2 and has a primary vein surface 31, radially internal, in the primary vein 106, and a vein surface secondary 32, radially outer, in the secondary stream 107. The primary stream surface 31 is downstream of the blade furthest downstream of the low-pressure compressor 120, and upstream of the blade furthest upstream of the high pressure compressor 130.
[0023] Dans l’exemple illustré à la figure 2, l’échangeur de chaleur 3 a une surface de veine tertiaire 33 dans la veine tertiaire 108. Il obstrue partiellement la veine tertiaire 108. Il comprend des ailettes 34 s’étendant radialement et parallèlement à l’axe de la turbomachine. In the example illustrated in Figure 2, the heat exchanger 3 has a tertiary stream surface 33 in the tertiary stream 108. It partially obstructs the tertiary stream 108. It comprises fins 34 extending radially and parallel to the axis of the turbomachine.
[0024] La présente invention a été décrite en relation avec des modes de réalisations spécifiques, qui ont une valeur purement illustrative et ne doivent pas être considérés comme limitatifs. D’une manière générale, la présente invention n’est pas limitée aux exemples illustrés et/ou décrits ci-dessus. L’usage des verbes « comprendre », « inclure », « comporter », ou toute autre variante, ainsi que leurs conjugaisons, ne peut en aucune façon exclure la présence d’éléments autres que ceux mentionnés. L’usage de l’article indéfini « un », « une », ou de l’article défini « le », « la » ou « », pour introduire un élément n’exclut pas la présence d’une pluralité de ces éléments. Les numéros de référence dans les revendications ne limitent pas leur portée. The present invention has been described in relation to specific embodiments, which are purely illustrative and should not be construed as limiting. In general, the present invention is not limited to the examples illustrated and/or described above. The use of the verbs "understand", "include", "compose", or any other variant, as well as their conjugations, can in no way exclude the presence of elements other than those mentioned. The use of the indefinite article “un”, “une”, or the definite article “le”, “la” or “”, to introduce an element does not exclude the presence of a plurality of these elements. The reference numbers in the claims do not limit their scope.

Claims

Revendications Claims
1. Système (1) pour refroidir de l’huile dans une turbomachine d’aéronef (100), et comprenant : · un carter de support intermédiaire (2) destiné à être localisé entre un compresseur basse pression (120) et un compresseur haute pression (130) de la turbomachine d’aéronef (100), et • un échangeur de chaleur (3) destiné à refroidir l’huile par échange thermique avec de l’air ; l’échangeur de chaleur (3) étant au moins partiellement intégré dans le carter de support intermédiaire (2). 1. System (1) for cooling oil in an aircraft turbine engine (100), and comprising: an intermediate support casing (2) intended to be located between a low-pressure compressor (120) and a high-pressure compressor pressure (130) of the aircraft turbine engine (100), and • a heat exchanger (3) intended to cool the oil by heat exchange with air; the heat exchanger (3) being at least partially integrated in the intermediate support casing (2).
2. Système (1) selon la revendication précédente, dans lequel l’échangeur de chaleur (3) comprend une surface de veine primaire (31) configurée pour être dans une veine primaire (106) de la turbomachine d’aéronef (100). 2. System (1) according to the preceding claim, in which the heat exchanger (3) comprises a primary stream surface (31) configured to be in a primary stream (106) of the aircraft turbine engine (100).
3. Système (1) selon la revendication précédente, dans lequel la surface de veine primaire (31) est configurée pour être entre l’aube la plus en aval du compresseur basse pression (120), et l’aube la plus en amont du compresseur haute pression (130). 3. System (1) according to the preceding claim, in which the primary stream surface (31) is configured to be between the most downstream blade of the low pressure compressor (120), and the most upstream blade of the high pressure compressor (130).
4. Système (1) selon l’une quelconque des revendications 2 ou 3, dans lequel l’échangeur de chaleur (3) est configuré pour obstruer partiellement la veine primaire (106). 4. System (1) according to any one of claims 2 or 3, wherein the heat exchanger (3) is configured to partially obstruct the primary stream (106).
5. Système (1) selon l’une quelconque des revendications précédentes, dans lequel l’échangeur de chaleur (3) comprend une surface de veine secondaire (32) configurée pour être dans une veine secondaire (107) de la turbomachine d’aéronef (100). 5. System (1) according to any one of the preceding claims, in which the heat exchanger (3) comprises a secondary stream surface (32) configured to be in a secondary stream (107) of the aircraft turbine engine. (100).
6. Système (1 ) selon l’une quelconque des revendications 2 à 4, dans lequel l’échangeur de chaleur (3) est configuré pour s’étendre radialement entre la veine primaire (106) et une veine secondaire (107), et comprend une surface de veine secondaire (32) configurée pour être dans une veine secondaire (107) de la turbomachine d’aéronef (100). 6. System (1) according to any one of claims 2 to 4, wherein the heat exchanger (3) is configured to extend radially between the primary stream (106) and a secondary stream (107), and includes a side stream surface (32) configured to be in a side stream (107) of the aircraft turbine engine (100).
7. Système (1) selon l’une quelconque des revendications 5 ou 6, dans lequel l’échangeur de chaleur (3) est configuré pour obstruer partiellement la veine secondaire (107). 7. System (1) according to any one of claims 5 or 6, wherein the heat exchanger (3) is configured to partially obstruct the secondary vein (107).
8. Système (1) selon l’une quelconque des revendications précédentes, dans lequel l’échangeur de chaleur (3) comprend une surface de veine tertiaire (33) configurée pour être dans une veine tertiaire (108) de la turbomachine d’aéronef (100). 8. A system (1) according to any preceding claim, wherein the heat exchanger (3) comprises a tertiary stream surface (33) configured to be in a tertiary stream (108) of the aircraft turbine engine. (100).
9. Système (1) selon la revendication précédente, dans lequel l’échangeur de chaleur (3) est configuré pour obstruer partiellement la veine tertiaire (108). 9. System (1) according to the preceding claim, in which the heat exchanger (3) is configured to partially obstruct the tertiary vein (108).
10. Système (1) selon l’une quelconque des revendications précédentes, dans lequel l’échangeur de chaleur (3) est annulaire et est configuré pour s’étendre autour d’un axe (X) de la turbomachine d’aéronef (100). 10. System (1) according to any one of the preceding claims, in which the heat exchanger (3) is annular and is configured to extend around an axis (X) of the aircraft turbine engine (100 ).
11. Système (1) selon l’une quelconque des revendications précédentes, dans lequel l’échangeur de chaleur (3) comprend des ailettes (34) configurées pour s’étendre radialement et parallèlement à l’axe de la turbomachine d’aéronef (100). 11. System (1) according to any one of the preceding claims, in which the heat exchanger (3) comprises fins (34) configured to extend radially and parallel to the axis of the aircraft turbomachine ( 100).
12. Système (1) selon l’une quelconque des revendications précédentes, dans lequel l’échangeur de chaleur (3) est une pièce fixée au carter de support intermédiaire (2). 12. System (1) according to any one of the preceding claims, in which the heat exchanger (3) is a part fixed to the intermediate support casing (2).
13. Système (1) selon l’une quelconque des revendications 1 à 11, dans lequel l’échangeur de chaleur (3) et le carter de support intermédiaire (2) sont d’une même pièce. 13. System (1) according to any one of claims 1 to 11, in which the heat exchanger (3) and the intermediate support casing (2) are of the same piece.
14. Turbomachine d’aéronef (100) comprenant un système (1) selon l’une quelconque des revendications précédentes, un compresseur basse pression (120) et un compresseur haute pression (130), le carter de support intermédiaire (2) étant situé entre le compresseur basse pression (120) et le compresseur haute pression (130). 14. Aircraft turbomachine (100) comprising a system (1) according to any one of the preceding claims, a low pressure compressor (120) and a high pressure compressor (130), the intermediate support casing (2) being located between the low pressure compressor (120) and the high pressure compressor (130).
15. Aéronef comprenant une turbomachine (100) selon la revendication précédente. 15. Aircraft comprising a turbomachine (100) according to the preceding claim.
EP21752650.8A 2021-07-22 2021-07-22 System for cooling oil in an aircraft turbine engine Pending EP4374056A1 (en)

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PCT/EP2021/070617 WO2023001379A1 (en) 2021-07-22 2021-07-22 System for cooling oil in an aircraft turbine engine

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EP4374056A1 true EP4374056A1 (en) 2024-05-29

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Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2075194B1 (en) * 2007-12-27 2017-08-16 Techspace Aero Air-oil heat exchanger for a turbojet, corresponding turbojet and use of said heat exchanger
FR3016956B1 (en) * 2014-01-29 2019-04-19 Safran Aircraft Engines HEAT EXCHANGER OF A TURBOMACHINE
FR3046199B1 (en) * 2015-12-23 2017-12-29 Snecma TURBOMACHINE COMPRISING A SURFACIAL AIR-OIL EXCHANGER INTEGRATED WITH AN INTER-VEIN COMPARTMENT
BE1024935B1 (en) * 2017-01-26 2018-08-27 Safran Aero Boosters S.A. COMPRESSOR WITH SEGMENTED INTERNAL VIROL FOR AXIAL TURBOMACHINE

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CN117751233A (en) 2024-03-22

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