Engine assembly for an aircraft comprising attachments for engines offset below on the fan frame
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02C-007/20
B64D-027/26
F01D-025/28
출원번호
US-0918500
(2009-02-27)
등록번호
US-9032740
(2015-05-19)
우선권정보
FR-08 51280 (2008-02-28)
국제출원번호
PCT/FR2009/050320
(2009-02-27)
§371/§102 date
20110316
(20110316)
국제공개번호
WO2009/112777
(2009-09-17)
발명자
/ 주소
Journade, Frédéric
Lafont, Laurent
Jalbert, Delphine
출원인 / 주소
AIRBUS OPERATIONS S.A.S.
대리인 / 주소
Oblon, McClelland, Maier & Neustadt, L.L.P.
인용정보
피인용 횟수 :
1인용 특허 :
11
초록▼
An aircraft engine assembly including an attachment mechanism of a turbojet onto a rigid structure of a mounting strut, including a first, second, and third forward engine attachment for taking up thrust loads brought onto the fan frame, and arranged such that the third attachment passes through a f
An aircraft engine assembly including an attachment mechanism of a turbojet onto a rigid structure of a mounting strut, including a first, second, and third forward engine attachment for taking up thrust loads brought onto the fan frame, and arranged such that the third attachment passes through a first diametrical plane of the turbojet, the first and second attachments being disposed on one side and another of the first plane. The first and second forward engine attachments are respectively brought to the fan frame at two points, situated beyond the second diametrical plane of the turbojet, orthogonal to the first diametrical plane, with respect to the third attachment.
대표청구항▼
1. An aircraft engine assembly comprising: a turbojet comprising a fan frame, an intermediate casing radially situated to an interior with respect to the fan frame and linked to the fan frame by a plurality of structural arms, and a central casing extending the intermediate casing towards a rear; an
1. An aircraft engine assembly comprising: a turbojet comprising a fan frame, an intermediate casing radially situated to an interior with respect to the fan frame and linked to the fan frame by a plurality of structural arms, and a central casing extending the intermediate casing towards a rear; anda mounting strut comprising a rigid structure comprising a longitudinal box and two lateral boxes integrated with the longitudinal box and arranged on two sides of the longitudinal box,the mounting strut further comprising a mount to mount the turbojet onto the rigid structure, the mount comprising first, second, and third forward engine attachments to take up thrust loads brought to the fan frame, and arranged such that the third forward engine attachment passes through a first diametrical plane of the turbojet, the first and second forward engine attachments, respectively carried on the two lateral boxes, being arranged on one side and another in the first diametrical plane, whereinthe first and second forward engine attachments are respectively brought to the fan frame at only two points, which are situated beyond a second diametrical plane of the turbojet, orthogonal to the first diametrical plane, with respect to the third forward engine attachment, such that the two lateral boxes include ends that carry the first and second forward engine attachments with the ends extending beyond the second diametrical plane of the turbojet, andthe lateral boxes each include a structural portion including a curved contour and a structural portion including a straight contour, and the structural portion including the straight contour carries one of the first and second forward engine attachments and extends beyond the second diametrical plane of the turbojet. 2. An assembly for an aircraft according to claim 1, wherein the first and second forward engine attachments that take up the thrust loads are situated symmetrically with respect to the first diametrical plane defined by a longitudinal axis of the turbojet parallel to a longitudinal direction of the turbojet, and a first direction of the turbojet orthogonal to the longitudinal direction. 3. An assembly for an aircraft according to claim 2, wherein in a head-on view along the longitudinal axis of the turbojet, an angle in a plane perpendicular to both the first and second diametrical planes and whose center is located along the longitudinal axis of the turbojet, between anchoring points of the third and first engine attachment, is strictly greater than 90° and less than or equal to 120°, or is between, but not including, 90° and 110°, and an angle in the plane perpendicular to both the first and second diametrical planes and whose center is located along the longitudinal axis of the turbojet, between anchoring points of the third and second engine attachment, is greater or equal to 240° and strictly less than 270°, or between, but not including, 250° and 270°. 4. An assembly for an aircraft according to claim 2, wherein the first and second forward engine attachments are each configured to take up loads exerted along the longitudinal direction and along the first direction of the turbojet, and the third forward engine attachment is configured to take up the loads exerted along the longitudinal direction and along a second direction of the turbojet, orthogonal to the first direction and to the longitudinal direction. 5. An assembly for an aircraft according to claim 4, wherein the first direction of the turbojet corresponds to a vertical direction of the turbojet, and the second direction of the turbojet corresponds to a transverse direction of the turbojet. 6. An assembly for an aircraft according to claim 1, wherein the mount is only constituted by the first, second, and third forward engine attachments. 7. An assembly for an aircraft according to claim 1, further comprising: an annular structure to transfer loads around the central casing and linked mechanically to the central casing via a mounting, andreinforcing structures, each of the first and second forward engine attachments is associated with one of the reinforcing structures, and each of the reinforcing structures forming a shearing plane and being rigidly attached: at a level of the annular structure in a first anchoring point;at a level of the fan frame and in a second anchoring point; andat a level of the structural arm or of the intermediate casing in a third anchoring point,each of the reinforcing structures lying along an imaginary plane, which is parallel to a longitudinal axis of the turbojet or passes through the longitudinal axis of the turbojet, and equally passing through the second anchoring point of the forward engine attachment on the fan frame. 8. An assembly for an aircraft according to claim 1, further comprising: an annular structure to take up loads surrounding the central casing and linked mechanically to the latter by a mounting comprising a plurality of attaching struts, the annular structure being attached to a plurality of structures arranged externally with respect to the annular structure and radially stressing respectively in a plurality of load introduction points spread circumferentially on the latter, and wherein at least one of the attaching struts is associated with each of the load introduction points, the at least one of the attaching struts being, in a head-on view along a longitudinal axis of the turbojet, disposed tangentially with respect to the central casing, and comprising an inner end attached to the central casing, and an outer end attached to the annular structure so as to be crossed by an imaginary radial plane passing through the longitudinal axis of the turbojet, and also through one of the load introduction points. 9. An aircraft comprising at least one engine assembly according to claim 1, assembled on a wing or on a part of an aft fuselage of the aircraft. 10. An assembly for an aircraft according to claim 1, wherein the first, second, and third forward engine attachments are each aligned with one of the plurality of structural arms. 11. An assembly for an aircraft according to claim 1, wherein the first, second, and third forward engine attachments form an isostatic take-up system of the thrust loads. 12. An assembly for an aircraft according to claim 1, wherein an external closing skin of each of the lateral boxes constitutes a part on an external aerodynamic surface of a nacelle. 13. An assembly for an aircraft according to claim 1, wherein an internal closing skin of each of the lateral boxes is part of an external radial delimitation of an annular bypass channel. 14. An assembly for an aircraft according to claim 1, wherein the lateral boxes are entirely situated above the fan frame.
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이 특허에 인용된 특허 (11)
Freid Wilbert B. (Bridgton ME), Aircraft gas turbine engine thrust mount.
Udall Kenneth F. (Derbyshire GB2) Wright Eric (Nottinghamshire GB2) Drew David D. K. (Derby GB2), Mounting for coupling a turbofan gas turbine engine to an aircraft structure.
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