# US 12,595,739 B2 · Gas turbine engine blading comprising a blade and a platform which has an internal flow-intake and flow-ejection canal

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- Number: US 12,595,739 B2
- Document ID: US12595739B2
- Status: Granted
- Kind: B2
- Application number: 18/840,981
- Filed: 2023-02-21
- Published: 2025-06-12
- Granted: 2026-04-07
- Earliest claimed: 2022-02-25
- Assignees: Safran Aircraft Engines
- Applicants: Safran Aircraft Engines
- Inventors: Nour Cherkaoui; Ludovic Pintat
- Web page: https://patentgrep.com/patent/US12595739B2

## CPC

- F01D 5/147: {Construction, i.e. structural features, e.g. of weight-saving hollow blades (F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18)}
- F01D 5/14: Form or construction (selecting particular materials, measures against erosion or corrosion F01D5/28)
- F01D 5/141: {Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147)}
- F01D 5/143: {Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour}
- F01D 5/145: {Means for influencing boundary layers or secondary circulations (for compressors F04D29/68)}
- F05D 2220/323: for aircraft propulsion, e.g. jet engines
- F05D 2240/80: Platforms for stationary or moving blades
- Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION

## Abstract

The present invention relates to blading (25, 26) for a turbomachine (10), comprising: —a blade (31) having an aerodynamic profile; —a platform (32, 33) comprising a flow-path surface (321) intended to delimit a primary flow path (21A) of the turbomachine (10), which path is intended, when the turbomachine (10) is in operation, to receive a flow that splits, upstream of the blade (31), into a suction-face flow (EE) and a pressure-face flow (EI); and —an internal canal (34) which has an intake opening (35) and an ejection opening (36), these each opening onto the flow-path surface (321) of the platform (32, 33), the ejection opening (36) opening downstream of the intake opening (35) and the intake opening (35) opening toward the pressure-face flow (EI).

## Claims

### Claim 1

A gas turbine engine blade assembly intended to be mounted about a longitudinal axis, and comprising:
an airfoil extending radially with respect to the longitudinal axis and has an aerodynamic profile delimited axially upstream by a leading edge and downstream by a trailing edge, the airfoil further comprising an intrados wall and an extrados wall opposite to the intrados wall, the intrados wall and the extrados wall each connecting the leading edge to the trailing edge;
a platform comprising a path surface, the airfoil extending from the path surface, the platform being intended to delimit a primary flow path of a stream of the gas turbine engine in operation, the stream dividing, upstream of the leading edge of the airfoil during the operation of the gas turbine engine, into an extrados flow flowing on a side of the extrados wall of the airfoil and into an intrados flow flowing on a side of the intrados wall of the airfoil; and
an inner channel having a suction aperture and an ejection aperture, the suction aperture and the ejection aperture are each disposed on the side of the intrados wall of the airfoil and opened out onto the path surface of the platform, the ejection aperture opening out downstream of the suction aperture and the suction aperture opening out towards the intrados flow.

### Claim 2 (depends on 1)

The gas turbine engine blade assembly according to claim 1, wherein the suction aperture and/or the ejection
aperture have at least one of:
a circular shape,
an oblong shape,
a shape slot,
a flared shape,
a plurality of orifices.

### Claim 3 (depends on 1)

The gas turbine engine blade assembly according to claim 1, wherein the intrados flow flows globally between a point of separation located upstream of the leading edge of the airfoil and where the intrados flow and the extrados flow divide, and a point of impact located downstream of the leading edge of the airfoil and where the intrados flow comes into contact with an airfoil circumferentially adjacent to the airfoil, wherein the suction aperture extends along a main direction and corresponds to one of the following four suction apertures:
a first suction aperture having an aperture opening out onto the path surface and extending along a main direction substantially perpendicular to a direction of the intrados flow, the first suction aperture opening out towards the leading edge or directly upstream of the leading edge of the airfoil;
a second suction has aperture having an aperture opening out onto the path surface and extending along a main direction substantially perpendicular to the direction of the intrados flow, the second suction aperture opening out downstream of the leading edge of the airfoil, the second suction aperture opening out closer to the point of separation than to the point of impact;
a third suction has aperture having an aperture opening out towards the point of impact; or
a fourth suction aperture having an aperture opening out onto the path surface and extending along a main direction substantially parallel to the direction of the intrados flow, the fourth suction aperture opening out between the leading edge and the trailing edge of the airfoil along a circumferential direction.

### Claim 4 (depends on 1)

The gas turbine engine blade assembly according to claim 1, wherein the ejection aperture extends along a main direction and corresponds to one of the following three ejection aperture:
a first ejection aperture located closer to the leading edge than to the trailing edge of the airfoil, the first ejection aperture opening out towards the intrados wall of the airfoil;
a second ejection aperture located substantially between the leading edge and the trailing edge of the airfoil, the second ejection aperture opening out towards the intrados wall of the airfoil, the second ejection aperture preferably having a main direction substantially parallel to the intrados wall of the airfoil; or
a third ejection aperture located closer to the trailing edge than to the leading edge of the airfoil and the third ejection aperture opening out towards the trailing edge of the airfoil.

### Claim 5 (depends on 1)

The gas turbine engine blade assembly according to claim 1, wherein a section of the inner channel of the ejection aperture is smaller than a section of the inner channel of the suction aperture.

### Claim 6 (depends on 1)

The gas turbine engine blade assembly according to claim 1 wherein the platform is an inner platform, the path surface of the inner platform being adapted to delimit the primary flow path radially inwards.

### Claim 7 (depends on 1)

The gas turbine engine blade assembly according to claim 1, extending radially about the longitudinal axis and further comprising another airfoil circumferentially adjacent to the airfoil, wherein the circumferentially adjacent airfoil extends radially with respect to the longitudinal axis and has an aerodynamic profile delimited axially upstream by a leading edge and downstream by a trailing edge, the circumferentially adjacent airfoil further comprises an intrados wall and an extrados wall opposite to the intrados wall, the intrados wall and the extrados wall each connecting the leading edge to the trailing edge, wherein the circumferentially adjacent airfoil is adapted to extend radially from the path surface of the platform in the primary flow path so that the extrados wall of the circumferentially adjacent airfoil is located facing the intrados wall of the airfoil, wherein the intrados flow globally flows between a point of separation located upstream of the leading edge of the airfoil and where the intrados flow and the extrados flow divide, and a point of impact located downstream of the leading edge of the airfoil and where the intrados flow comes into contact with the extrados wall of the circumferentially adjacent airfoil.

### Claim 8 (depends on 1)

The gas turbine engine blade assembly according to claim 1, wherein the blade assembly is a gas turbine engine turbine distributor.

### Claim 9 (depends on 1)

A gas turbine engine comprising at least a turbine comprising at least a blade assembly according to claim 1.

### Claim 10 (depends on 4)

The gas turbine engine blade assembly according to claim 4, wherein the ejection aperture extends along a main direction and being located closer to the trailing edge than to the leading edge of the airfoil and the third ejection aperture opening out towards the trailing edge of the airfoil.