Turbomachine with fan rotor and reduction gearbox driving a low-pressure decompressor shaft

Abstract

A turbomachine comprising a ducted fan, a low-pressure turbine shaft and a reduction gearbox housed in a casing between the fan and the low-pressure turbine shaft, the fan rotor supplying airflow to a primary stream and a secondary stream and comprising a hub of diameter D1, wherein—the diameter D3 of the fan rotor is greater than 82 inches (2.08 metres), —the pressure ratio of the fan is between 1.10 and 1.35, the turbomachine comprises a low-pressure compressor separate from the fan, the reduction gearbox being interposed between the fan rotor and a turbine shaft of the low-pressure compressor, and wherein the reduction gearbox casing has an outside diameter D2 greater than the diameter D1 of the hub, the pitch diameter D4 of the reduction gearbox ring being between 0.15 and 0.35 times the fan rotor diameter.

Claims

1. A turbomachine comprising: a ducted fan including a fan rotor; a low pressure turbine shaft; and a reduction gearbox comprising a ring gear and housed in a casing between the ducted fan and the low pressure turbine shaft, wherein the fan rotor is configured to supply air flow to a primary stream and a secondary stream and includes a hub of a diameter and blades each extending from a blade root at the hub to a blade tip, the blades of the fan rotor being of variable setting type, and an external surface of the hub being continuous with a leading surface of the blades at the blade roots, a diameter of the fan rotor is greater than 82 inches (2.08 meters), a pressure ratio of the ducted fan is between 1.10 and 1.35, the diameter of the hub is between 11.811 inches (300 millimeters) and 23.622 inches (600 millimeters), the turbomachine further comprises a low pressure compressor distinct from the ducted fan, the reduction gearbox is interposed between the fan rotor and the low pressure turbine shaft, the casing of the reduction gearbox has an external diameter greater than the diameter of the hub, and a pitch diameter of the ring gear of the reduction gearbox is between 0.15 and 0.35 times the diameter of the fan rotor, and the turbomachine further comprises a nacelle which is a protective fairing surrounding the fan rotor, wherein a main stream defined between the protective fairing and the hub is substantially cylindrical at the blade roots, a slope of the hub at the blade roots being null or lower than 5°.

2. The turbomachine according to claim 1, wherein the diameter of the fan rotor is between 90 inches (2.29 metres) and 150 inches (3.81 metres).

3. The turbomachine according to claim 1, wherein the reduction gearbox is of epicyclic type.

4. The turbomachine according to claim 1, wherein the diameter of the hub of the ducted fan is between 0.25 and 0.35 of a diameter of the ducted fan.

5. The turbomachine according to claim 1, wherein a reduction rate of the reduction gearbox is between 3.5 and 8.

6. The turbomachine according to claim 1, wherein a reduction rate of the reduction gearbox is between 5 and 6.

7. The turbomachine according to claim 1, wherein a ratio of the external diameter of the casing of the reduction gearbox and the diameter of the hub of the ducted fan is between 1 and 1.10.

8. The turbomachine according to claim 1, wherein the protective fairing comprises a spherical annular reinforcement wall opposing the blade tips across a clearance, the blade tips being of a general complementary form.

9. The turbomachine according to claim 8, wherein a mean clearance between the spherical annular reinforcement wall and the blade tips is less than 0.35% of a chord length of the blades when the turbomachine is at its maximum ground regime.

10. The turbomachine according to claim 8, wherein a mean clearance between the spherical annular reinforcement wall and the blade tips is less than 0.65% of a chord length of the blades in a cruise regime of the turbomachine.

11. An aircraft including a turbomachine according to claim 1.

12. A turbomachine comprising: a ducted fan including a fan rotor; a low pressure turbine shaft; and a reduction gearbox comprising a ring gear and housed in a casing between the ducted fan and the low pressure turbine shaft, wherein the fan rotor is configured to supply air flow to a primary stream and a secondary stream and includes a hub of a diameter and blades each extending from a blade root at the hub to a blade tip, the blades of the fan rotor being of variable setting type, and an external surface of the hub being continuous with a leading surface of the blades at the blade roots, a diameter of the fan rotor is greater than 82 inches (2.08 meters), a pressure ratio of the ducted fan is between 1.10 and 1.35, the diameter of the hub of the ducted fan is between 0.25 and 0.35 of a diameter of the ducted fan, the turbomachine further comprises a low pressure compressor distinct from the ducted fan, the reduction gearbox is interposed between the fan rotor and the low pressure turbine shaft, the casing of the reduction gearbox has an external diameter greater than the diameter of the hub, and a pitch diameter of the ring gear of the reduction gearbox is between 0.15 and 0.35 times the diameter of the fan rotor, and the turbomachine further comprises a nacelle which is a protective fairing surrounding the fan rotor, wherein a main stream defined between the protective fairing and the hub is substantially cylindrical at the blade roots, a slope of the hub at the blade roots being null or lower than 5°.

13. A turbomachine comprising: a ducted fan including a fan rotor; a low pressure turbine shaft; and a reduction gearbox comprising a ring gear and housed in a casing between the ducted fan and the low pressure turbine shaft, wherein the fan rotor is configured to supply air flow to a primary stream and a secondary stream and includes a hub of a diameter and blades each extending from a blade root at the hub to a blade tip, the blades of the fan rotor being of variable setting type, and an external surface of the hub being continuous with a leading surface of the blades at the blade roots, a diameter of the fan rotor is greater than 82 inches (2.08 meters), a pressure ratio of the ducted fan is between 1.10 and 1.35, a ratio of the external diameter of the casing of the reduction gearbox and the diameter of the hub of the ducted fan is between 1 and 1.10, the turbomachine further comprises a low pressure compressor distinct from the ducted fan, the reduction gearbox is interposed between the fan rotor and the low pressure turbine shaft, the casing of the reduction gearbox has an external diameter greater than the diameter of the hub, and a pitch diameter of the ring gear of the reduction gearbox is between 0.15 and 0.35 times the diameter of the fan rotor, and the turbomachine further comprises a nacelle which is a protective fairing surrounding the fan rotor, wherein a main stream defined between the protective fairing and the hub is substantially cylindrical at the blade roots, a slope of the hub at the blade roots being null or lower than 5°.

14. The turbomachine according to claim 1, wherein the protective fairing comprises a spherical annular reinforcement wall opposing the blade tips across a clearance, the blade tips being of a general complementary form, wherein a mean clearance between the spherical annular reinforcement wall and the blade tips is less than 0.35% of a chord length of the blades when the turbomachine is at its maximum ground regime, wherein a mean clearance between the spherical annular reinforcement wall and the blade tips is less than 0.65% of a chord length of the blades in a cruise regime of the turbomachine.

Description

PRESENTATION OF THE FIGURES

(1) Other characteristics and advantages of the invention will become more apparent from the following description, which is purely illustrating and non limiting, and should be read in light of the accompanying figures on which:

(2) FIG. 1 is a schematic representation in a sectional view (half-view) illustrating the integration of a fan reduction gearbox in a turbomachine in accordance with a possible embodiment of the invention;

(3) FIG. 2 illustrates a configuration example with a spherical blade tip;

(4) FIGS. 3a and 3b illustrate the definition of a chord at the blade root, as well as the clearances between a blade root and the fairing wall;

(5) FIG. 4 illustrates a structure for holding the blade in a cell by means of a holding wedge.

DESCRIPTION OF ONE OR SEVERAL IMPLEMENTATION MODES AND EMBODIMENTS

(6) The turbomachine T illustrated on FIG. 1 exhibits an architecture with a ducted fan having an ultra high by-pass ratio, called UHBR.

(7) It includes a nacelle 1, a fan rotor 2, as well as a primary stream 3, defined in a casing 5.

(8) It is also represented on FIG. 1 an inter-compressor casing 8 of the turbomachine, an inter-turbine casing 9, as well as an exhaust casing 16.

(9) The nacelle 1 is compact and particularly of reduced length. In particular, it does not integrate air inlets or secondary nozzles upstream or downstream of the fan.

(10) Neither does it integrate a thrust reversal mechanism.

(11) Its main functions are to ensure the turbomachine aerodynamic fairing and the retention of the fan vanes/blades and is dimensioned solely for this purpose.

(12) A rectifier assembly 4 is interposed between the nacelle 1 and the casing 5 and allows to hold said nacelle 1.

(13) In a possible embodiment, a portion of the nacelle 1 can be made jointly with an already existing surface on the aircraft, such as for example the underwing.

(14) The blades of the fan rotor 2 are blades 6 with variable settings (mechanism 6a).

(15) The blade 6 setting can in particular be controlled to drive the fan while in operation. The very low pressure ratio of the latter in fact induces variations in the cycle parameters between ground and flight conditions of unusual amplitude, particularly as regards turbine operating temperatures HP and nozzle expansion ratios.

(16) Controlling the blade setting allows to adapt to these operating condition deviations.

(17) In addition, it is used to ensure braking action of the aircraft or to contribute to it.

(18) The diameter D3 of the fan rotor 2 is important: greater than 82 inches (2.08 metres), and preferably between 90 (2.29 metres) and 150 inches (3.81 metres).

(19) The fan pressure ratio (FPR) is low: ranging between 1.10 and 1.35.

(20) With this dimensioning taken into account, the rotational speed of the rotor 2 is low.

(21) A reduction gearbox 7 is therefore provided for driving the shaft A of the low pressure turbine.

(22) This reduction gearbox 7 allows for a high low pressure turbine regime: between 3.5 and 8 times the regime of the rotor 2 and preferably between 5 and 6 times the speed regime of the latter. The reduction ratio and the torque to be transmitted define the encumbrance of the reduction gearbox. Here, the reduction gearbox 7 is of epicyclic type and hence, its reduction ratio is defined by: 1+ (the number of teeth of the ring/the number of teeth of the central sun gear). The torque to be transmitted defines the minimum size of the teeth and the minimum diameter of the central sun gear. However, here the power of the reduction gearbox should be between 10 and 40 MW. The pitch diameter of the ring D4 is hence complex to integrate for such a reduction ratio and ranges between 0.15 and 0.35 times the fan diameter.

(23) The diameter D3 of the fan is determined in a standard manner, by projection of the radial component at the fan blade 6 tip, onto a radial straight line passing by the leading edge of the blade, at its root.

(24) The hub ratio is defined as the ratio of the internal radius at the fan blade 6 root, measured at the leading edge of the blading (at its design setting, in the case where the blading has a variable setting), and the external radius of the leading edge of the blade 6 projected onto the same straight line. In order to ensure a good efficiency of the turbomachine, the hub ratio is restricted to the maximum, thereby the hub diameter ranges between 0.25 and 0.35 the diameter of the fan.

(25) In particular, the radius at the base of the fan can range between 300 and 600 mm.

(26) To integrate a reduction gearbox with a high reduction rate while maintaining the smallest possible hub ratio without affecting the aerodynamic characteristics of the primary stream the casing surrounding the reduction gearbox (casing 15) includes a limited outgrowth. Particularly, the ratio between the external diameter (diameter D2 of the casing 15 of the reduction gearbox) and the diameter D1 of the fan hub 10 is greater than 1 (D2>D1) and ranges between 1 and 1.10, and preferably lower than 1.04.

(27) Such a ratio simultaneously allows for the required aerodynamic form for the primary stream 3 and the integration of operations of the reduction gearbox (oil discharge for example) and of the fan (pitch change system), while maintaining a hub ratio that is as low as possible. The inlet casing wherein the reduction gearbox is integrated is particularly cumbersome as it has to hold the reduction gearbox, absorb the axial thrust generated by the fan by means of the ball bearing and hold the low pressure shaft.

(28) Furthermore, the inlet power of the reduction gearbox ranges between 10 and 40 MW at takeoff (@ T/O)-altitude 0, Mach ranging between 0.15 and 0.28)).

(29) The thus, constituted propulsion system meets the following objectives: maximisation of the propulsion efficiency thanks to the fan having a very low pressure ratio; competitiveness in terms of fuel consumption for classes of thrust and flight speeds for medium-haul and longer haul applications (thrust >15000 lbf in take-off condition 0 m/zero airspeed/IAS conditions; 0.65<cruise Mach<0.9.

(30) The fairing (nacelle 1) allows for minimum drag and is not disadvantageous in mass.

(31) The configuration of blade with variable setting 6 illustrated on FIG. 2 is particularly interesting, in particular to achieve the thrust reversal function.

(32) In this configuration, the stream 11 between the fairing 1 and the hub 10 is substantially cylindrical at the root of the blade, that is to say, at the hub 10, such as to limit generating steps in the stream while avoiding complex blade forms.

(33) More particularly, the slope of the hub 10 at the blade root (discontinuous line on FIG. 2) makes a null angle or lower than 5° with respect to the axis of the shaft A of the turbomachine.

(34) In addition, the blades 6 exhibit at their tip a general spherical or substantially spherical form (slight differences in radius at the blade tips may be present between the leading edge and the trailing edge). This general spherical form is itself received in a spherical annular reinforcement formed on the fairing wall 1, thereby, contributing to defining the stream 11.

(35) FIG. 2 represents an arc of a sphere S corresponding to this general spherical form at the blade tip, and a spherical reinforcement on the fairing wall. This arc of a sphere S is centered on the intersection between the axis of the shaft A of the turbomachine and the setting axis of a blade (axis C on FIG. 2). The radius of this arc of a sphere S corresponds to the largest radius of the fan rotor at the blade tip.

(36) The spherical cutout at the blade tip prevents contact at the blade tip whatever the blade 6 setting orientation. In fact, the blade tip turns perfectly in the spherical annular reinforcement that receives it on the fairing 1 wall, without any blocking or creating of a significant clearance.

(37) This way, efficiency losses are minimised.

(38) The clearance J between a blade tip and the fairing 1 wall has been illustrated on FIG. 3a.

(39) This clearance provided between the blades 6 and the fairing 1 should enable to absorb variations as per the dimensions of the blades in operation.

(40) The mean clearance is lower than 0.35% of the chord at the tip (double arrow Co on FIGS. 3a and 3b) when the engine nears its maximum ground regime (Red Line sol). It can reach 0.65% of this chord during flight in cruise regime.

(41) It is noteworthy, that the chord corresponds to the length of the profile, that is to say, the shortest distance between the leading edge and the trailing edge (FIG. 3b). For the chord at the tip, it is the distance Co between the end at the tip of the leading edge and the end at the tip of the trailing edge (FIG. 3a).

(42) Furthermore, the mean clearance should also allow for the disassembly of the blades 6 and their exit with respect to the engine, for example once the holding wedge at the blade root is removed.

(43) In fact, it is known, that for maintaining the fan blades, it is usually provided for the latter, at their internal end, a root engaged axially in the cells of the fan disk and radially retained by the disk teeth. A wedge is interposed between each blade root and the bottom of the corresponding cell. Examples of assembly/disassembly of blades on a fan disk is for example described in application FR3034130.

(44) An example of wedge is illustrated on FIG. 4 on which is represented a blade 6, whereof the root 12 is axially engaged in a cell 13 of the rotor disk D. A wedge 14 is arranged between the bottom of the cell 13 and the blade root 12.

(45) The radially external faces of the wedges 14 of the blades espouse the blade roots 12 whereas the radially internal faces of said wedges 14 espouse the bottoms of the cells 13. These wedges 14 are relatively flat and extend over the entire length of the bottoms of the cells. A wedge is thus interposed between each blade root and the corresponding cell bottom, for the purpose of maintaining and preventing premature wear.

(46) As for the clearance J (FIG. 3a), it is provided for the assembly/disassembly of a blade 6.

(47) To this end, it should be ensured that the height difference between the radius at the leading edge at the blade tip and the maximum radius at the blade tip at the setting axis C) be lower than the sum of the clearance space at the blade tip (clearance between, on the one hand the blade tip and the fairing wall 1 which contributes in defining the stream 11) and on the other hand, the height of the wedge under the blade root.

(48) Also, the setting axis C may not be exactly perpendicular to the axis of the shaft A, but may be slightly slanted upstream or downstream.