# US 11,898,467 B2 · Aircraft engine struts with stiffening protrusions

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- Number: US 11,898,467 B2
- Document ID: US11898467B2
- Status: Granted
- Kind: B2
- Application number: 17/650,756
- Filed: 2022-02-11
- Published: 2023-08-17
- Granted: 2024-02-13
- Earliest claimed: 2022-02-11
- Assignees: Pratt & Whitney Canada Corp.
- Applicants: Pratt & Whitney Canada Corp.
- Inventors: Mateusz Ozog
- Web page: https://patentgrep.com/patent/US11898467B2

## CPC

- F01D 9/041: {using blades (F01D5/148 takes precedence)}
- B64D 27/10: of gas-turbine type (jet aircraft B64D27/16)
- B64D 27/402: {comprising box like supporting frames, e.g. pylons or arrangements for embracing the power plant}
- F05D 2220/323: for aircraft propulsion, e.g. jet engines
- F05D 2240/12: Fluid guiding means, e.g. vanes
- F05D 2260/941: particularly aimed at mechanical or thermal stress reduction
- F05D 2270/114: by limiting mechanical stresses
- 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)
- B64D: EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- 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

An aircraft engine, comprising: a first wall and a second wall defining a gaspath between the first wall and the second wall, the gaspath extending around a central axis, each of the first wall and the second wall having a gaspath side facing the gaspath and an opposed side facing away from the gaspath; struts circumferentially distributed around the central axis, the struts extending across the gaspath, a strut of the struts having an airfoil including a leading edge, a trailing edge, a first end secured to the first wall, and a second end secured to the second wall; and protrusions extending from a baseline surface of the opposed side of the first wall, the protrusions increasing a thickness of the first wall, a protrusion of the protrusions overlapping a location where the leading edge or the trailing edge meets the gaspath side of the first wall.

## Claims

### Claim 1

An inlet case for an aircraft engine, comprising:
a conduit annularly extending around a central axis, the conduit having an inlet oriented in a substantially radial direction and an outlet oriented in a substantially axial direction, the conduit defined between a first wall and a second wall, each of the first wall and the second wall having a gaspath side facing the conduit and an opposed side facing away from the conduit;
struts circumferentially distributed around the central axis, the struts extending across the gaspath, a strut of the struts having an airfoil including a leading edge, a trailing edge, a first end secured to the first wall, and a second end secured to the second wall, the airfoil having a chord line connecting the leading edge to the trailing edge, wherein the volume of each strut is defined in part by each of the first wall, the second wall, the leading edge, and the trailing edge; and
protrusions extending from a baseline surface of the opposed side of the first wall and external to the respective volume of each strut, the protrusions being circumferentially distributed around the central axis and being circumferentially spaced apart from one another, the protrusions increasing a thickness of the first wall, a protrusion of the protrusions overlapping a location where the leading edge or the trailing edge meets the gaspath side of the first wall, the protrusion extending from a first protrusion end to a second protrusion end in a direction of an airflow in the gaspath, the first protrusion end and the second protrusion end located on the baseline surface, the first protrusion end being upstream of the leading edge or downstream of the trailing edge relative to the airflow, the second protrusion end being downstream of the leading edge and upstream of the trailing edge, the protrusion intersected by a projection of the chord line on the opposed side of the first wall.

### Claim 2 (depends on 1)

The inlet case of claim 1, wherein the first wall, the second wall, and the struts are part of a monolithic body.

### Claim 3 (depends on 1)

The inlet case of claim 1, comprising a fillet extending around the first end of the strut, the fillet merging in the airfoil and in the gaspath side of the first wall, the location including an intersection between the fillet and the gaspath side of the first wall.

### Claim 4 (depends on 1)

The inlet case of claim 1, wherein the protrusion has a length extending from the first protrusion end to the second protrusion end in the direction parallel to the airflow flowing in the conduit, a ratio of the length to a nominal thickness of the first wall ranging from 3 to 40.

### Claim 5 (depends on 4)

The inlet case of claim 4, wherein the ratio of the length to the nominal thickness ranges from 8 to 12.

### Claim 6 (depends on 1)

The inlet case of claim 1, wherein a ratio of a thickness of the protrusion to a nominal thickness of the first wall ranges from 0.1 to 5.

### Claim 7 (depends on 6)

The inlet case of claim 6, wherein the ratio ranges from 0.4 to 1.

### Claim 8

An aircraft engine, comprising:
a first wall and a second wall defining a gaspath between the first wall and the second wall, the gaspath extending around a central axis, each of the first wall and the second wall having a gaspath side facing the gaspath and an opposed side facing away from the gaspath;
struts circumferentially distributed around the central axis, the struts extending across the gaspath, a strut of the struts having an airfoil including a leading edge, a trailing edge, a first end secured to the first wall, and a second end secured to the second wall, the airfoil having a chord line connecting the leading edge to the trailing edge, wherein the volume of each strut is defined in part by each of the first wall, the second wall, the leading edge, and the trailing edge; and
protrusions extending from a baseline surface of the opposed side of the first wall and external to the respective volume of each strut, the protrusions being circumferentially distributed around the central axis and being circumferentially spaced apart from one another, the protrusions increasing a thickness of the first wall, a protrusion of the protrusions overlapping a location where the leading edge or the trailing edge meets the gaspath side of the first wall, the protrusion extending from a first protrusion end to a second protrusion end in a direction of an airflow in the gaspath, the first protrusion end and the second protrusion end located on the baseline surface, the first protrusion end being upstream of the leading edge or downstream of the trailing edge relative to the airflow, the second protrusion end being downstream of the leading edge and upstream of the trailing edge, the protrusion intersected by a projection of the chord line on the opposed side of the first wall.

### Claim 9 (depends on 8)

The aircraft engine of claim 8, comprising a compressor section, the struts located upstream of the compressor section.

### Claim 10 (depends on 8)

The aircraft engine of claim 8, comprising a fillet extending around the first end of the strut, the fillet merging in the airfoil and in the gaspath side of the first wall, the location including an intersection between the fillet and the gaspath side of the first wall.

### Claim 11 (depends on 8)

The aircraft engine of claim 8, wherein the protrusion has a length extending from the first protrusion end to the second protrusion end in a direction parallel to the direction of the airflow flowing in the gaspath, a ratio of the length to a nominal thickness of the first wall ranging from 3 to 40.

### Claim 12 (depends on 11)

The aircraft engine of claim 11, wherein the ratio of the length to the nominal thickness ranges from 8 to 12.

### Claim 13 (depends on 8)

The aircraft engine of claim 8, wherein the protrusion has a width in a direction transverse to a direction of an airflow flowing in the gaspath, a ratio of the width to a nominal thickness of the first wall ranging from 3 to 40.

### Claim 14 (depends on 13)

The aircraft engine of claim 13, wherein the ratio of the width to the nominal thickness ranges from 8 to 12.

### Claim 15 (depends on 8)

The aircraft engine of claim 8, wherein an intersection between the protrusion and the opposed side of the first wall has a circular perimeter.

### Claim 16 (depends on 8)

The aircraft engine of claim 8, wherein a ratio of a thickness of the protrusion to a nominal thickness of the first wall ranges from 0.1 to 5.

### Claim 17 (depends on 16)

The aircraft engine of claim 16, wherein the ratio ranges from 0.4 to 1.

### Claim 18 (depends on 8)

The aircraft engine of claim 8, wherein the protrusion overlaps the location where the leading edge meets the gaspath side of the first wall.

### Claim 19 (depends on 18)

The aircraft engine of claim 18, wherein another protrusion of the protrusions overlaps a second location where the trailing edge meets the gaspath side of the first wall.