Aircraft engine comprising a propeller having blades that can be steered in a propulsion or thrust reversal orientation
By using a propeller with pivotable blades that change pitch angle for thrust reversal, the aircraft engine reduces the need for traditional thrust reversers, simplifying the system and ensuring effective braking.
Patent Information
- Application Number
- FR2021009370
- Authority / Receiving Office
- FR · FR
- Patent Type
- Patents
- Current Assignee / Owner
- Filing Date
- 2021-09-07
- Publication Date
- 2025-06-20
- Estimated Expiration
- 2041-09-07
AI Technical Summary
Existing aircraft engines require large thrust reverser devices to reverse thrust for braking, which increase the engine's size and complexity.
The engine incorporates a propeller with blades that can be oriented for both propulsion and thrust reversal, where the blades pivot to change their pitch angle from positive at the root to negative at the tip during thrust reversal, eliminating the need for traditional thrust reversers.
This solution reduces the size and complexity of the thrust reverser system, ensuring effective thrust reversal while maintaining efficient air supply to the engine during the reversal phase.
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Abstract
Description
Title of the invention: Aircraft engine comprising a propeller having blades which can be oriented in a propulsion or thrust reversal orientation. Technical field
[0001] The invention relates to the thrust reversal of an aircraft engine such as a turbojet, a turboprop or the like. STATE OF THE PRIOR ART
[0002] In a turbojet, air is admitted into an inlet sleeve to pass through a fan comprising a series of rotating blades before splitting into a central primary flow and a secondary flow surrounding the primary flow.
[0003] The primary flow is then compressed in compression stages before reaching a combustion chamber, after which it is expanded through high pressure and low pressure turbines before being discharged to the rear. The secondary flow is propelled directly to the rear by the fan in a vein delimited externally by a fairing of the engine.
[0004] Such an engine comprises a low pressure body by which the fan is coupled to the low pressure turbine, and a high pressure body by which the high pressure compressor is coupled to the high pressure turbine, these two bodies being coaxial and independent in rotation.
[0005] The various components of the engine are carried by the external fairing which surrounds them, by means of radial arms, this fairing being carried by the wing of the aircraft.
[0006] This fairing is equipped with thrust reversers which, when activated, allow the secondary flow to be diverted to return it towards the front of the engine so that the overall thrust of the engine is temporarily directed from downstream to upstream so as to brake the aircraft when it lands.
[0007] These thrust reversers can be panels integrated into the hull which pivot on themselves when activated: the downstream part of the panel enters the secondary flow and the upstream part leaves the hull, so as to jointly take a part of the secondary flow and expel it out of the hull in the upstream direction.
[0008] Generally speaking, optimizing the efficiency of such an engine leads to increasing its diameter in order to increase the secondary flow mass flow, so that it is ultimately necessary to reduce the thickness of the fairing which is strongly conditioned by the presence of the thrust reversers.
[0009] The aim of the invention is to provide a solution for reducing the size of the thrust reverser device of such an engine. Statement of the invention
[0010] For this purpose, the invention relates to an aircraft engine comprising a propulsion propeller carrying blades and rotating about a longitudinal axis, each blade extending radially from a root to a tip along a span axis around which it is orientable, the engine comprising an air intake located downstream of the propeller and extending around the longitudinal axis to supply combustion members of the engine, the blades being able to occupy a propulsion orientation to generate thrust oriented from upstream to downstream of the engine, and a thrust reversal orientation to generate thrust oriented from downstream to upstream, characterized in that at the thrust reversal orientation, the blades have at their roots a pitch angle relative to a plane normal to the longitudinal axis which is of opposite sign to their pitch angle at their tips.
[0011] With this arrangement, air is validly admitted into the air intake to correctly supply the engine in the thrust reversal phase, so that the thrust reversal is entirely ensured by the propulsion propeller. It is therefore not necessary to equip the nacelle with traditional thrust reversers.
[0012] The invention also relates to an engine thus defined, of the turbojet type carried by a nacelle, in which the propulsion propeller is a fan located in the upstream part of the nacelle, and in which the nacelle is devoid of a thrust reversal device.
[0013] The invention also relates to an engine thus defined, in which at the thrust reversal orientation, the blades have at their roots a pitch angle greater than 10°.
[0014] The invention also relates to an engine thus defined, in which each blade is twisted with a pitch which evolves in a substantially linear manner from its root to its tip.
[0015] The invention also relates to an engine thus defined, in which when the blades have a propulsion orientation, the root of each blade has a pitch greater than 70°.
[0016] The invention also relates to an engine thus defined, in which when the blades have a propulsion orientation, they have a pitch angle whose sign is opposite to that of the pitch angle at the tip of the blade over an extent of between 5% and 20% of the distance separating the root from the tip starting from the root.
[0017] The invention also relates to an engine thus defined, comprising control means for increasing the rotation speed of the propeller in the thrust reversal phase. Brief description of the drawings
[0018] [Fig.l] is an overall view in longitudinal section of a dual-flow turbojet engine;
[0019] [Fig.2] is a half-view in longitudinal section of a front part of a turbojet engine in the propulsion phase;
[0020] [Fig. 3] is a perspective view of a fan blade according to the invention;
[0021] [Fig.4] is a view along the span axis of a fan blade having a propulsion orientation in accordance with the invention;
[0022] [Fig. 5] is a view along the span axis of a fan blade having a thrust reversal orientation in accordance with the invention;
[0023] [Fig.6] is a half-view in longitudinal section of a front part of a turbojet engine in the thrust reversal phase;
[0024] [Fig.7] is a diagram representing the linear evolution of the blade pitch at a propulsion orientation and at a thrust reversal orientation in accordance with the invention;
[0025] [Fig.8] is a table giving ranges of pitch angle values along the blade for thrust reversal and propulsion orientations.
[0026] DETAILED DESCRIPTION OF PARTICULAR EMBODIMENTS
[0027] In the aircraft engine 1 shown in FIGS. 1 and 2, the air is admitted at the level of an inlet sleeve 2 located upstream AM to pass through a propulsion propeller 3, also called a fan, comprising a series of rotating blades before splitting into a central primary flow FP and a secondary flow FS surrounding the primary flow. These two flows circulate in the engine in its longitudinal direction AX, from upstream AM to downstream AV of this engine when it generates a propulsion thrust.
[0028] The primary flow FP is admitted into an air intake 4 located downstream of the propeller 3 and extending around the longitudinal axis AX, to then be compressed in low pressure 5 and high pressure 6 compressors before arriving in a combustion chamber 7. This primary flow FP is then expanded through a high pressure turbine 8 and a low pressure turbine 9 before being evacuated towards the rear. The secondary flow is propelled directly towards the rear by the fan in a vein delimited externally by a nacelle 11 carrying the entire engine.
[0029] Such a twin-body engine comprises a so-called low-pressure body by which the propulsion propeller 3 is coupled to the low-pressure turbine 9, and a so-called high-pressure body by which the high-pressure compressor 6 is coupled to the high-pressure turbine 8, these two bodies being coaxial and independent in rotation.
[0030] These bodies are surrounded by a set of casings 12 following one another along the longitudinal axis which they surround. These casings 12 comprise internal shells surrounded by external shells and connected to each other by radial arms, the primary flow circulating in an annular space extending between these internal and external shells. This set of casings 12 is itself surrounded by the nacelle 11 of the engine while being carried by the latter by means of other radial arms.
[0031] As visible in [Fig. 3], a blade 13 of the propeller 3 has a lower surface face and an upper surface face extending between a leading edge 14 and a trailing edge 15. Generally speaking, it is a lopsided shape extending, along a so-called span axis EV normal to the axis AX, from a root P by which it is rigidly secured to a rotor element not shown, to a vertex S which corresponds to its free end which is opposite the internal face of the nacelle 11.
[0032] For a given position along the span axis EV, the blade pitch angle is defined as the angle between the chord axis, which passes through the leading edge and the trailing edge while extending in a plane normal to the EV axis, and the fan plane PS which is normal to the AX axis. In the following, when such a pitch angle has a positive value, it corresponds to a downstream propulsion of the air flow in which the rotating blade is immersed, i.e. a propulsion direction. A negative value corresponds to an upstream propulsion of the air flow.
[0033] In the case of a fan blade like that of [Fig.3], the shape of the blade has a significant twist, which means that the pitch angle is very different depending on whether the root or the top of the blade is considered, this angle evolving gradually from the root to the top.
[0034] In practice, the pitch angle evolves in a substantially linear manner along the span axis, from its value at the root P of the blade to its value at the tip S. The pitch is thus substantially collinear with the local direction of the air flow seen from the moving blade, and this at any point of this blade along its span direction.
[0035] In [Fig.4], this blade 13 is in a propulsion orientation, typically at a so-called cruising regime: the chord axis Ap at the blade root is inclined by a positive pitch angle Cp worth approximately +85° relative to the axis AX, while the chord axis As at the blade tip is inclined by a pitch angle Cs worth approximately +30° relative to the axis AX. The twist of this blade 13 is thus approximately 55° between its root P and its tip S.
[0036] In the configuration corresponding to a propulsion orientation of [Fig.4], the pitch angle of the blade 13 is positive all along the span axis EV, so that when it is rotated in the direction marked R in the figures, it propels the primary flow FP and the secondary flow FS from upstream AM to downstream AV. The direction of rotation R corresponds to the clockwise direction when looking at the blower along the axis AX, from the front, that is to say from its upstream AM.
[0037] This blade is of the variable pitch type, which means that its orientation around the axis EV is adjustable according to the operating conditions of the engine. In the situation of [Fig.4], the orientation of the blade 13 corresponds to operation of the engine in propulsion, that is to say in which the secondary flow FS is propelled by the engine from upstream AM to downstream AV.
[0038] In the thrust reversal phase, for example when the aircraft is landing and the engine is being used to brake the aircraft, the fan blades are controlled to change their orientations. They then pivot around their span axes, in order to move from the propulsion orientation corresponding to [Fig.4], to the thrust reversal orientation corresponding to the situation shown in [Fig.5].
[0039] As shown in [Fig.5], the transition from the propulsion orientation to the thrust reversal orientation corresponds to a pivot Rr of approximately +50° around the span axis EV.
[0040] Thus, at the thrust reversal orientation corresponding to [Fig.5], the apex pitch angle Cs' is negative and is approximately -20°, whereas the root pitch angle Cp' remains positive and is approximately +35°.
[0041] Thus, at the thrust reversal orientation, the positive pitch angle Cs' at the blade tip allows the fan to push back the secondary flow FS from downstream AV to upstream AM in order to brake the aircraft, but thanks to the pitch angle Cp' which remains positive at the blade root, the air of the primary flow FP continues to be propelled from upstream AM to downstream AV to be admitted into the air intake 4 in order to ensure the supply of air to the compressor.
[0042] This is obtained by means of a pitch setting Cp at the propulsion orientation which is greater than the pivot angle Rr of the blade when it passes from the propulsion orientation to the thrust reversal orientation.
[0043] More generally, the profile of the blade is arranged to have a negative setting over most of its height from the top and a positive setting from the root region and possibly beyond this root, when this blade has a thrust reversal orientation. In other words, the passage of the blade to its thrust reversal orientation closes the setting at the root, but as this is provided to be sufficiently open it remains positive to allow the admission of the primary flow into the compressor.
[0044] In practice, and as illustrated schematically in [Fig.7], when the blade is in its thrust reversal orientation, its pitch is positive over an extent corresponding to approximately 5% to 20% of the distance separating the root P from the tip S, starting from the root P. As can be understood, this extent preferably corresponds to the external diameter of the air intake 4, the pitch angle in the thrust reversal position being substantially zero at the level of the external diameter of this air intake 4. Furthermore, and as visible in [Fig.7], when the blade is in its propulsion orientation, its pitch is entirely positive, from the root P to the tip S.
[0045] As an example, the table in [Fig.8] gives possible value ranges for the pitch angles C and C' of the blade at its propulsion orientation and at its thrust reversal orientation, in a case where the resetting angle Rr to move from one orientation to the other is 65°. These angles are given for six positions along the blade span axis, starting from the root P which corresponds to the value 0% to the tip S which corresponds to the value 100%.
[0046] Furthermore, in the region of the blade root, the profile of the blade is designed to be robust to incidence, by having a significant thickness at the leading edge and an extrados slope substantially aligned with the direction of flow of the stream.
[0047] Additionally, to improve the aerodynamic operation at the blade root in thrust reversal, engine control means are provided to increase the rotation speed of the propeller in the thrust reversal phase, so as to maintain the primary flow rate at a sufficient level.
Claims
Claims
1. Aircraft engine (1) comprising a propulsion propeller (3) carrying blades (13) and rotating about a longitudinal axis (AX), each blade (13) extending radially from a root (P) to a tip (S) along a span axis (EV) about which it is orientable, the engine comprising an air intake (4) located downstream of the propeller (3) and extending about the longitudinal axis (AX) to supply combustion members of the engine (1), the blades (13) being able to occupy a propulsion orientation to generate thrust oriented from upstream (AM) to downstream (AV) of the engine (1), and a thrust reversal orientation to generate thrust oriented from downstream (AV) to upstream (AM), characterized in that at the thrust reversal orientation,the blades (13) have at their roots (P) a pitch angle relative to a plane (PS) normal to the longitudinal axis (AX) which is of opposite sign to their pitch angle (Cs') at their tips (S) over an extent of between 5% and 20% of the distance separating the root (P) from the tip (S) starting from the root (P), and in that each blade (13) is twisted having a pitch (C) which evolves in a substantially linear manner from its root (P) to its tip (S).,
2. Engine according to claim 1, of the turbojet type carried by a nacelle (11), in which the propulsion propeller (3) is a fan located in the upstream part (AM) of the nacelle (11), and in which the nacelle (11) is devoid of a thrust reverser device.
3. Engine according to claim 1 or 2, in which at the thrust reversal orientation, the blades (13) have at their roots (P) a pitch angle (Cp') greater than 10°.
4. Engine according to one of the preceding claims, in which when the blades (13) have a propulsion orientation, the root (P) of each blade (13) has a pitch greater than 70°.
5. Engine according to one of the preceding claims, comprising control means for increasing the rotation speed of the propeller (3) in the thrust reversal phase.