Office Action 1: Notice of Allowance (2025-06-10)

CLAIMS:
AMENDMENTS TO THE CLAIMS

claim 1, wherein the upstream suspension members are attached to the third wall

and/or to the fourth wall .

laim 1, wherein said thrust-absorbing rods extend from the upstream suspension

members or downstream of these members to connection points on the gas generator, these connection points being located inside the third wall .

laim 1, wherein at least one part of said second and third walls are connected

together by tubular reinforcements or arms, to form an assembly .

1\. (Currently amended) A triple-flow turbine engine for an aircraft, this turbine engine having a longitudinal axis and comprising:- a gas generator comprising at least one compressor , an annular combustion chamber , and at least one turbine ,
\- at least one main propeller rotated by a shaft of the gas generator ,
the gas generator comprising:
\- a first annular splitter nose for separating two annular ducts for the respective flow of an internal primary flow inside the gas generator and an external secondary flow around the gas generator ,
\- a second annular splitter nose for separating two annular ducts for the respective flow of a first internal flow and of a second external flow ,
\- a first nozzle for ejecting the first internal flow ,
\- a second nozzle for ejecting the second external flow ,
\- a first annular compartment , referred to as hot compartment, extending axially between the second splitter nose and the first nozzle , and between first and second annular walls , respectively internal and external, which extend around each other, the first wall externally defining said duct for the flow of the first internal flow , and said second wall internally defining said duc for the flow of the second external flow ,
\- a second annular compartmen, referred to as cold compartment, extending axially between the first splitter nose and the second nozzle , and between third and fourth annular walls , respectively internal and external, which extend around each other, the third wall externally defining said duct for the flow of the second external flow , and said fourth wall internally defining said duct for the flow of the secondary flow ,
\- upstream suspension members of the turbine engine , which are located in a first plane perpendicular to said axis and which are connected or attached to the gas generator ,
\- downstream suspension members of the turbine engine , which are located in a second plane perpendicular to said axis and which are connected or attached to the gas generator , and
\- thrust-absorbing rods which comprise first ends connected or attached to the gas generator and opposite second ends located in a third plane perpendicular to said axis ,
wherein said first, second and third planes are located axially between the first splitter nose and the second nozzle , and in that at least one part of said upstream and downstream suspension members are connected or attached to said third wall .
2\. (Currently amended) The turbine engine according to claim 1, wherein the downstream suspension members are attached or connected to said third wall .
3\. (Currently amended) The turbine engine according to claim 1 , wherein the downstream suspension members are located in the cold compartmen.
4\. (Currently amended) The turbine engine according to
5\. (Currently amended) The turbine engine according to claim 4, wherein the upstream suspension members are located just downstream of said first splitter nose .
6\. (Currently amended) The turbine engine according to
7\. (Currently amended) The turbine engine according to

8\. (Currently amended) The turbine engine according to claim 7, wherein the second and third walls comprise openings and removable hatches for closing these openings.

9\. (Currently amended) The turbine engine according to claim 7 , wherein said assembly forms a segment in one piece.

10\. (Currently amended) The turbine engine according to claim 7 , wherein said assembly comprises a first segment in one piece and a second segment which can be detached from the first segment and which can itself be in one piece or formed of two half- shells articulated with respect to each other about an axis substantially parallel to said longitudinal axis .

11\. (Currently amended) The turbine engine according to laim1, wherein at least one heat exchange is located in said second external flow duct , and/or at least one equipment , such as an accessory box or gearbox, is located in said cold compartment .

12\. (Currently amended) The turbine engine according to laim1, wherein said at least one main propeller is located upstream of the gas generator .

13\. (Currently amended) The turbine engine according to claim 1,wherein said at least one main propeller is located downstream of the gas generator .

14\. (Currently amended) The turbine engine according to laim1, wherein it comprises a secondary propeller driven in rotation by a shaft of the gas generator , this propeller being located in the duc for the flow of the primary flow and upstream of said second annular splitter nose .

15\. (Currently amended) An assembly comprising a turbine engine according to laim1, and a pylon for hanging this turbine engine to an aircraft, the pylon having a generally elongate shape and being connected or attached respectively to the upstream and downstream suspension members of the turbine engine .

OFFICE ACTION:
Allowable Subject Matter
Claims 1-15 are allowed.
The following is an examiner’s statement of reasons for allowance: The prior art of record is silent to a triple-flow turbine engine, as claimed, and having the suspension members claimed, oriented and located in the same manner. It would not have been obvious to modify the art of record.
Any comments considered necessary by applicant must be submitted no later than the payment of the issue fee and, to avoid processing delays, should preferably accompany the issue fee.  Such submissions should be clearly labeled “Comments on Statement of Reasons for Allowance.”



CITED REFERENCE SUMMARIES: Ref US6209311B1: Claims:
A turbofan engine, comprising:
a high-pressure turbine rotated by combusted gas discharged from a combustion chamber;
a low-pressure turbine rotated by combusted gas discharged from said high-pressure turbine;
a high-pressure compressor having a first shaft common with, and driven by, said high-pressure turbine for supplying compressed air into said combustion chamber;
a low-pressure compressor having a second shaft common with said low-pressure turbine for compressing air supplied from the upstream of said engine to be discharged downstream;
an air turbine receiving an airflow discharged from said low-pressure compressor to cause rotation, at least a portion of an airflow channeled past said air turbine being directed to said high-pressure compressor;
a flow splitter supported to said air turbine rotatably around an axis of said engine; and
a fan disposed on an outer periphery side of said flow splitter for rotating integrally with said air turbine to thereby produce propulsive force.
The turbofan engine according to claim 1 , wherein
said low-pressure compressor includes a blade rotating integrally with said second shaft, and a passive blade rotating integrally with said flow splitter.
The turbofan engine according to claim 1 , further comprising:
an expansion turbine for receiving a portion of the airflow past said air turbine to rotate said flow splitter.
The turbofan engine according to claim 3 , wherein
an airflow passing through an inner part of said air turbine is directed to said high-pressure compressor, an airflow passing through an outer part thereof is directed to said expansion turbine, and an air turbine splitter is disposed at the boundary between said inner part and said outer part.
Ref US9091207B2: Claims:
An assembly of a gas turbine engine and of a nacelle in which the gas turbine engine is housed, the nacelle comprising an air inlet fairing forming an air inlet comprising: a foreign-object deflection member positioned inside the air inlet fairing, the deflection member and an interior wall of the air inlet fairing delimiting an air intake duct; a secondary deflection passage disposed downstream of the deflection member and including an inlet and an outlet opening via opening in an exterior wall of the nacelle; a main engine air supply passage disposed downstream of the deflection member and leading to an inlet of a compressor of the gas turbine engine, the main engine air supply passage being concentric with and inside of the secondary deflection passage; said air intake duct being configured to deflect at least some foreign objects that have been aspirated by the air inlet toward the secondary deflection passage, wherein an upstream edge of a surface splits the air inlet between the secondary deflection passage and the main engine air supply passage, wherein the secondary deflection passage is delimited by first and second substantially radial walls extending longitudinally between the inlet and the outlet, a radially interior cylindrical wall, and a radially exterior cylindrical wall in continuation with the interior wall of the air inlet fairing, the walls delimiting the secondary deflection passage being fixed relative to each other, and wherein an area of a cross section of the secondary deflection passage, transverse to a direction in which the air flows, progressively decreases between the inlet and the outlet of the secondary deflection passage such that a speed at which the air passing through the secondary deflection passage flows is increased from upstream to downstream.
a foreign-object deflection member positioned inside the air inlet fairing, the deflection member and an interior wall of the air inlet fairing delimiting an air intake duct;
a secondary deflection passage disposed downstream of the deflection member and including an inlet and an outlet opening via opening in an exterior wall of the nacelle;
a main engine air supply passage disposed downstream of the deflection member and leading to an inlet of a compressor of the gas turbine engine, the main engine air supply passage being concentric with and inside of the secondary deflection passage;
said air intake duct being configured to deflect at least some foreign objects that have been aspirated by the air inlet toward the secondary deflection passage,
wherein an upstream edge of a surface splits the air inlet between the secondary deflection passage and the main engine air supply passage,
wherein the secondary deflection passage is delimited by first and second substantially radial walls extending longitudinally between the inlet and the outlet, a radially interior cylindrical wall, and a radially exterior cylindrical wall in continuation with the interior wall of the air inlet fairing, the walls delimiting the secondary deflection passage being fixed relative to each other, and
wherein an area of a cross section of the secondary deflection passage, transverse to a direction in which the air flows, progressively decreases between the inlet and the outlet of the secondary deflection passage such that a speed at which the air passing through the secondary deflection passage flows is increased from upstream to downstream.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 1 , wherein the foreign-object deflection member blocks any ballistic trajectory passing through the air inlet from entering the main engine air supply passage.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 1 , wherein the deflection member is in a shape of an axisymmetric bullet which forms an annular air intake passage with the air inlet fairing which is of annular shape.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 1 , the decrease of the area of the cross section of the secondary passage, transverse to the direction in which the air flows, between the inlet and the outlet of the secondary passage is achieved by azimuthal reduction.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 1 , wherein the secondary deflection passage is formed of at least two separate ducts with an annular inlet that is common and outlets with openings distributed around the periphery of the nacelle.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 1 , wherein an air discharge opening in the wall of the nacelle is configured such that the airflow is directed along the axis of the engine.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 1 , wherein the nacelle comprises a removable cowling element in the continuation of said air inlet fairing, and wherein the secondary deflection passage comprises at least one portion of the secondary deflection passage that forms a deflection scoop and is secured to said removable cowling element.
wherein the secondary deflection passage comprises at least one portion of the secondary deflection passage that forms a deflection scoop and is secured to said removable cowling element.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 7 , further comprising a removable cowling element articulated about an axis parallel to the axis of the engine so as to uncover the engine.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 1 , wherein the object deflection member is in a shape of an axisymmetric bullet which forms an annular air intake passage with the air inlet fairing which is of annular shape, the deflection member being supported at least in part by a first hub casing including an internal hub, by being engaged in said internal hub.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 9 , wherein said first hub casing is fixed to the engine by a second hub casing that forms a plane for suspending the engine from an aircraft.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 10 , further comprising a cowling element articulated about an axis parallel to the axis of the engine so as to uncover the engine, a portion of the secondary deflection passage secured to the articulated cowling element having upstream surfaces that press in a fluidtight manner against bearing surfaces forming the secondary deflection passage inside the air inlet fairing.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 11 , wherein said bearing surfaces are formed on the second hub casing.
The assembly of a gas turbine engine and of a nacelle as claimed in claim 11 , wherein fan blades of the gas turbine engine are disposed downstream of the outlet of the secondary deflection passage.
Ref US20210323690A1: Claims:
- 10 . (canceled)
Turbomachine extending along an axis, comprising a high-pressure body with a high-pressure compressor coupled in rotation to a high-pressure turbine, and a low-pressure body including a low-pressure compressor coupled in rotation to a low-pressure turbine, an upstream casing located upstream of the high-pressure compressor, and means of suspension for the turbomachine, intended for attaching the turbomachine to an aircraft pylon, the said means of suspension comprising an upstream suspension system with a beam intended to be attached to the pylon, at least two connecting rods spaced apart circumferentially from each other and each articulated at a first end to the beam and at a second end to the upstream casing, a rudder bar articulated on the beam at the level of a circumferentially central zone of said rudder bar, at least two force-recovery rods to take up the thrust forces of the turbomachine, spaced apart circumferentially from each other and each articulated at a first end on the rudder bar and at a second end on a fixed part of the turbomachine, the said at least two force-recovery rods being located upstream of a radial suspension plane defined by the two connecting rods, the means of suspension also comprising a downstream suspension system capable of connecting the pylon to a fixed downstream casing of the turbomachine.
Turbomachine according to claim 11 , characterised in that the upstream casing is an inter-compressor casing, located axially between the high-pressure compressor and the low-pressure compressor.
Turbomachine according to claim 11 , characterised in that a first connecting rod and a second connecting rod of said connecting rods extend in a radial plane of the turbomachine, the first end and the second end of each of the first and second connecting rods being articulated about articulation axes parallel to the axis of the turbomachine.
Turbomachine according to claim 12 , characterised in that a first connecting rod and a second connecting rod of said connecting rods extend in a radial plane of the turbomachine, the first end and the second end of each of the first and second connecting rods being articulated about articulation axes parallel to the axis of the turbomachine.
Turbomachine according to claim 11 , characterised in that the upstream suspension system of the means of suspension comprises a third connecting rod articulated at a first end to the beam and articulated at a second end to the upstream casing, a functional clearance being provided on at least one articulation of the third connecting rod to the beam and/or to the upstream casing in such a way that the third connecting rod is not solicited during normal operation and is solicited in the event of rupture of one of the first and second connecting rods.
Turbomachine according to claim 12 , characterised in that the upstream suspension system of the means of suspension comprises a third connecting rod articulated at a first end to the beam and articulated at a second end to the upstream casing, a functional clearance being provided on at least one articulation of the third connecting rod to the beam and/or to the upstream casing in such a way that the third connecting rod is not solicited during normal operation and is solicited in the event of rupture of one of the first and second connecting rods.
Turbomachine according to claim 13 , characterised in that the upstream suspension system of the means of suspension comprises a third connecting rod articulated at a first end to the beam and articulated at a second end to the upstream casing, a functional clearance being provided on at least one articulation of the third connecting rod to the beam and/or to the upstream casing in such a way that the third connecting rod is not solicited during normal operation and is solicited in the event of rupture of one of the first and second connecting rods.
Turbomachine according to claim 15 , characterized in that the third connecting rod extends in an axial plane and is articulated at its first end and at its second end about articulation axes perpendicular to the axis of the turbomachine.
Turbomachine according to claim 11 , characterised in that the second ends of the connecting rods are articulated on a mounting plate attached to the upstream casing.
Turbomachine according to claim 12 , characterised in that the second ends of the connecting rods are articulated on a mounting plate attached to the upstream casing.
Turbomachine according to claim 13 , characterised in that the second ends of the connecting rods are articulated on a mounting plate attached to the upstream casing.
Turbomachine according to claim 15 , characterised in that the second ends of the connecting rods are articulated on a mounting plate attached to the upstream casing.
Turbomachine according to claim 18 , characterised in that the second ends of the connecting rods are articulated on a mounting plate attached to the upstream casing.
Turbomachine according to claim 11 , characterized in that the fixed part of the turbomachine is an intermediate casing, and the second end of each force-recovery rod is articulated on a flange of the intermediate casing, the said flange being situated radially between a primary flow vein and a secondary flow vein of the turbomachine.
Turbomachine according to claim 12 , characterized in that the fixed part of the turbomachine is an intermediate casing, and the second end of each force-recovery rod is articulated on a flange of the intermediate casing, the said flange being situated radially between a primary flow vein and a secondary flow vein of the turbomachine.
Turbomachine according to claim 13 , characterized in that the fixed part of the turbomachine is an intermediate casing, and the second end of each force-recovery rod is articulated on a flange of the intermediate casing, the said flange being situated radially between a primary flow vein and a secondary flow vein of the turbomachine.
Turbomachine according to claim 15 , characterized in that the fixed part of the turbomachine is an intermediate casing, and the second end of each force-recovery rod is articulated on a flange of the intermediate casing, the said flange being situated radially between a primary flow vein and a secondary flow vein of the turbomachine.
Turbomachine according to claim 11 , characterized in that the force-recovery rods extend in a plane forming an angle with the radial plane.
Turbomachine according to claim 11 , characterised in that one of the connecting rods has an L-shape adapted to take up the torque forces Mx acting on the turbomachine in a circumferential direction about the axis X.
Turbomachine according to claim 24 , with a reducer for transmission between a shaft of the low-pressure body and a fan located upstream of the intermediate casing, characterized in that the articulations of the force-recovery rods on the intermediate casing are located in a radial plane through which a downstream part of the reducer passes.
Ref US11851201B2: Claims:
Turbomachine extending along an axis, comprising a high-pressure body with a high-pressure compressor coupled in rotation to a high-pressure turbine, and a low-pressure body including a low-pressure compressor coupled in rotation to a low-pressure turbine, an upstream casing located upstream of the high-pressure compressor, and means of suspension for the turbomachine, intended for attaching the turbomachine to an aircraft pylon, the said means of suspension comprising an upstream suspension system with a beam intended to be attached to the pylon, at least two connecting rods spaced apart circumferentially from each other and each articulated at a first end to the beam and at a second end to the upstream casing, a rudder bar articulated on the beam at the level of a circumferentially central zone of said rudder bar, at least two force-recovery rods to take up the thrust forces of the turbomachine, spaced apart circumferentially from each other and each articulated at a first end on the rudder bar and at a second end on a fixed part of the turbomachine, the said at least two force-recovery rods being located upstream of a radial suspension plane defined by the two connecting rods, the means of suspension also comprising a downstream suspension system capable of connecting the pylon to a fixed downstream casing of the turbomachine.
Turbomachine according to claim 1 , characterised in that the upstream casing is an inter-compressor casing, located axially between the high-pressure compressor and the low-pressure compressor.
Turbomachine according to claim 1 , characterised in that a first connecting rod and a second connecting rod of said connecting rods extend in a radial plane of the turbomachine, the first end and the second end of each of the first and second connecting rods being articulated about articulation axes parallel to the axis of the turbomachine.
Turbomachine according to claim 2 , characterised in that a first connecting rod and a second connecting rod of said connecting rods extend in a radial plane of the turbomachine, the first end and the second end of each of the first and second connecting rods being articulated about articulation axes parallel to the axis of the turbomachine.
Turbomachine according to claim 1 , characterised in that the upstream suspension system of the means of suspension comprises a third connecting rod articulated at a first end to the beam and articulated at a second end to the upstream casing, a functional clearance being provided on at least one articulation of the third connecting rod to the beam and/or to the upstream casing in such a way that the third connecting rod is not solicited during normal operation and is solicited in the event of rupture of one of the first and second connecting rods.
Turbomachine according to claim 2 , characterised in that the upstream suspension system of the means of suspension comprises a third connecting rod articulated at a first end to the beam and articulated at a second end to the upstream casing, a functional clearance being provided on at least one articulation of the third connecting rod to the beam and/or to the upstream casing in such a way that the third connecting rod is not solicited during normal operation and is solicited in the event of rupture of one of the first and second connecting rods.
Turbomachine according to claim 3 , characterised in that the upstream suspension system of the means of suspension comprises a third connecting rod articulated at a first end to the beam and articulated at a second end to the upstream casing, a functional clearance being provided on at least one articulation of the third connecting rod to the beam and/or to the upstream casing in such a way that the third connecting rod is not solicited during normal operation and is solicited in the event of rupture of one of the first and second connecting rods.
Turbomachine according to claim 5 , characterized in that the third connecting rod extends in an axial plane and is articulated at its first end and at its second end about articulation axes perpendicular to the axis of the turbomachine.
Turbomachine according to claim 1 , characterised in that the second ends of t
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