GB2483495A - Rotor blade disc, eg for a turbofan engine, having blades supported by an outer ring - Google Patents

Abstract

A rotor blade disc, for example for a turbofan engine, comprises blades 20 that are supported by a high-strength outer ring 21 against rotational forces. The blades (20) are attached to a central shaft or hub (29, fig.5) in a manner that permits the transmission of torque but does not need to withstand the rotational forces, eg via retention lugs 37. Because the blades (20) are loaded in compression, a lighter design can be employed. The outer ring 21 can also serve as a containment shield for blade-off events. The outer ring 21 can form one race of a bearing, eg an air bearing, for the disc. In a turbofan engine, the rotor disc may be surrounded by a fan casing (22, fig.1) the inner wall of which defines the outer race of the air bearing. A power take-off for accessories may be taken from the outer ring 21.

Description

Mounting of rotor blades

Field of the invention

The present invention relates to the mounting of aerofoils. More particularly, but not exclusively, it relates to rotating aerofoils in high-bypass turbofan engines. But will find application in any situation where aerofoils or other bodies are required to rotate about a common axis such as in wind turbines, turbo machinery and pumping equipment.

Prior Art

Turbine, fan and impeller rotors, in particular those used in turbine engines and other turbo-machinery are generally designed with a central shaft, which runs in one or more bearings. Affixed to the shaft is a hub, to which the aerofoils are attached, the whole being placed within a duct or cowling. In the current state of the art, the aerofoils may either be formed as a part of the central hub or engaged with it by means of dovetail or fir-tree' shaped mating features formed in the root of the aerofoil (blade), and correspondingly-shaped features formed in the hub.

Due to the high rotational speed of such turbo-machinery, the blade to hub connection is subject to high stresses resulting from the centripetal forces generated. The blade roots and the hub therefore represent large concentrations of mass, which is required to retain the blades across the operational envelope of the machine.

Fig. 1 shows a turbofan aero engine constructed according to the current state of the art methods. Fan blades (1) are attached to hub (2) by means of mating slots and dovetail features (3), the gaps between the fan blades (1) are filled with fan annulus fillers (4) and an aerodynamic spinner (5) attached to the hub (2). The hub (2) is connected with the engine low pressure shaft (6), running in main bearing (7) via a curvic coupling (8), which is attached to the low pressure turbine within the engine core (9). The engine core (9) is suspended from the aircraft including the engine nacelle (10) via the fan outlet guide vanes (11) and thrust struts in some configurations, or alternatively, the engine core (9) is attached to the aircraft via struts within the bypass duct bifurcation and the nacelle (10) is mounted onto the engine core (9). A typical fan blade (1) from a turbofan engine is shown in Fig. 2, where the blade (1) has a distal end or tip (12) and a proximal end or root (13). The body of the blade (1) is of a twisted aerofoil section with a leading edge (14) and a trailing edge (15). The root (13) of the blade (1) is attached to a dovetail or fir-tree mating feature (16) which is shaped to transfer S the operational loads into the hub (2). The proximal section of the blade (17) is where most of the mass is concentrated. Such blades are generally manufactured to fine tolerances from high strength materials such titanium alloys. The engine nacelle (10) incorporates zones of sound-attenuating material (18) and containment shielding (19) as shown in Fig. 3.

It is a particular disadvantage of the current state of the art turbine designs that the means of blade attachment requires that the blade roots (13) are of suitably massive construction. The disadvantage is two-fold in that a mass penalty is not just engendered by the design of the blade roots (13) themselves, but by the engine nacelle, which is required to contain the blades and other high-energy debris in the case of a failure in the hub or blade root. The containment casing represents a parasitic (non thrust-generating) mass, which is present throughout the lifetime of the engine and apart from providing an attachment point for the nacelle nose, is only used for debris containment in the unlikely event of a blade off event. A blade-off event in engines of the current state of the art generally takes the form of a fracture at the blade root (13) in a running engine. The blade is flung outwards until the tip (12) embeds in the fan case. The root (13) of the blade will then pivot about the embedded tip (12), and is afforded further energy as the successive blades impact upon it, hammering the blade (1) upon the engine containment casing. Thus the containment system is required to absorb energy over and above that pertaining to the failed fan blade alone.

Furthermore, with the current state of the art means of blade attachment, the clearance between the tip of the blade and the interior surface of the turbine casing is liable to change as the engine heats up and as the blades (1) stretch or creep in operation. It is current practice to provide some rotor stages within the engine core (9) with abradable or sacrificial linings, such that the blades (1) therein create the interior profile of the engine when it is first test-run. This method allows for the minimum overall clearance between blade tip envelope and turbine wall to be achieved, but does not provide a minimum clearance for all operating conditions due to the differential thermal expansion between the engine components at different operating conditions. The fan blade tip clearance is designed to allow the fan blades to stretch and slightly untwist, such that the minimum clearance is obtained at maximum power. This is not ideal as the engine FADEC system does not run the engine at maximum power for the majority of its operational life, which is usually spent at flight idle power levels. The fan blade tip clearance in current state of the art engine designs represents a source of inefficiency, as it introduces vortices and boundary layer detachment in the flow of the working fluid.

It is a further disadvantage that the current state of the art turbine architecture requires that power take-off shafting and other such accessories must impinge on both the bypass duct and on the turbine core in order to couple with the high pressure spool of the engine. These factors cause further flow inefficiencies and are also sources of parasitic mass in the engine design. Representative patents include: U.S. Pat. AppI. 2008/ 0038116 Al discloses a tip shroud for conventional-type gas turbine engine rotor blades whereby each blade incorporates a tip section, attached to the outer section of the aerofoil section. The shroud has beneficial aerodynamic effects by causing the gas flow to remain attached to the airfoil section. The shrouds also serve to locate the individual blades relative to one another when the engine is in a cool state, preventing shingling' and fretting of the blades. Thus the prior art shows the need to align turbine blades when the engine is in a cooled state and to maintain the boundary-layer gas flow at the aerofoil tips.

U.S. Pat AppI. 2004/00135 18 Al discloses a method for providing blade tip sealing for a gas turbine engine. The method employs portions of the casing which are fabricated from an abradable material. At first start-up, the components of the engine will expand and contract due to the action of thermal expansion and inertial forces thereby causing the blade tips to remove material from the engine casing. In this manner, the best casing profile is attained, which affords the best sealing between blade disc and casing over the operational range of the engine. Thus the prior art shows the need to provide sealing between the blade disc and the casing of a gas turbine engine.

U.S. Pat 6,543,991 discloses a means for the containment of detached rotor blades within a turbofan gas turbine engine. The prior art employs energy-absorbing structures such as honeycombs and fibrous blankets in order to contain the released debris. Further methods use thickened sections of the fan casing to act as containment rings, or a combination of the aforementioned means. Thus the prior art shows the need to provide containment for detached rotor blades and other debris in the case of an engine failure.

U.S. Pat. 6,378,293 discloses an arrangement of bearings for use within a gas turbine engine. The prior art employs a combination of conventional rolling contact bearings and electromagnetic bearings, which allows the lubrication demands of an engine constructed according to the principles of the invention, and therefore the complexity and weight to be reduced. Thus the prior art shows the need to provide bearing support within a gas turbine engine.

U.S. Pat. 3,015,524 discloses an axial-flow turbojet engine in which the turbine blades are supported on the inner periphery of an annular air-bearing. This allows for the blades to be manufactured from materials having low tensile strengths such as cermets, but which are operable at very high temperatures. Thus the prior art teaches the mounting of turbine blades such that they are loaded in compression.

U.S Pat 2,509,442 discloses an inverse rotor whereby a rotating annulus is disposed within a shroud and supported by ball bearings. Blades extend radially inwards and perform work on the fluid within the rotor when the aforesaid is rotated by means of a pinion gear enmeshed with a corresponding gear ring on the rotor. Thus the prior art allows the use of an annular rotor having inwardly-extending blades.

Statement of the Invention

The present invention provides a means whereby the aforesaid problems may be addressed in a manner allowing for a new architecture of gas turbine engine design such that the overall efficiency of such an engine may be improved, the approach noise signature reduced, and mass reductions made. Furthermore, the present invention allows for simpler blade retention in the event of a failure and allows a reduction in the number of components within an engine, whilst enabling the use of technologies that hitherto have not been employed in the aero-engine or turbine

construction fields.

In an exemplary embodiment of the invention, these and other objects are met by employing a construction for the rotor fan blade disc that incorporates a plurality of blade elements comprising a root section, an aerofoil section and a tip section, disposed around the inner periphery of an annular member, thereby forming a complete set of blades and an outer blade ring in one item. The rotor blades, being affixed to the ring at their respective tips are connected to a shaft by means of a hub. The hub is not required to retain the blades against centripetal forces (it may contribute towards load carrying capability), which are taken as hoop stresses in the annular member and may therefore be manufactured to a lighter weight than is possible with current state of the art methods. Also, the blades themselves are loaded in compression instead of in tension as is the case in current state of the art designs, which allows for a lighter design of blade to be employed.

In another exemplary embodiment of the invention the outer peripheral edge of the fan ring forms the inner race of a bearing, with the outer race being integral with the engine casing. Using this configuration, the mounting structure of the engine may be simplified as the main bearing may be easily accessed from the outside of the engine. Furthermore, this embodiment has the advantage of providing a seal between the rotor blade and the engine casing, which increases its efficiency. This may also be particularly useful in the field of micro and miniature turbine power generation, where the size of bearings, and therefore their heat rejection capabilities becomes increasingly important, along with overall efficiency. Suitable bearings may be of the rolling contact type, or may employ a non-contact bearing such as fluid dynamic bearings (aerodynamic or hydrodynamic) or electromagnetic bearings, either of the active compensating type or of the passive homopolar type.

An air bearing is preferred. The outlet at the downstream side of the of the air bearing may be configured to minimize detachment of the airflow from the inner wall of the fan casing.

It is a particular feature of the present invention that the functions of torque transfer and bearing support are disassociated. In all prior art, the two aforementioned functions are co-located, such as a shaft or cylinder. This dissociation is advantageous because it allows for the members employed to perform these functions in an engine to be designed independently in the most efficient manner. For example, in a fan blade disc in a turbofan engine, the connection between the hub and the blades may be designed such that it is stiff enough to transmit torque into the blades, but may be very lightweight as it does not have to retain the fan blades under imparted centripetal loading. This simplification of load paths also encourages the use of alternative materials in engine design.

It is an advantage of the present invention that the mass of material required to form the bearing surfaces at the periphery of the fan ring will serve as a sufficient containment shield. This removes the requirement to include material in the engine design specifically for the purposes of containment, which may result in a lighter, more efficient engine. The bearing structure may also be employed as a suitable mounting point to the aircraft, further simplifying the design of the engine and allowing reductions in engine weight to be made.

it is a yet further advantage of the invention that the rotor tip and casing are connected, thereby eliminating the gap between them which is common to all engines of the current state of the art. Supersonic flow in the tip-shroud interface region causes shockwaves to form, reducing the efficiency of the turbine stage and generating a distinctive, high amplitude noise signature commonly referred to as buzz-saw noise'. This noise is predominantly heard during ground manoeuvring and takeoff and will be reduced in engines constructed according to the present invention, thereby reducing noise pollution in civil aviation operations and noise signature in military applications. The elimination of rotor-tip to engine casing clearance present in current state of the art engines will also enable more efficient engines to be manufactured, with particular improvements being made in the bypass duct, where the majority of the engine's thrust is evolved. This in turn allows a smaller engine designed according to the present invention to generate the thrust of a larger, state of the art engine. By reducing the diameter of the fan, the frontal area and hence drag of the engine is reduced, resulting in an aircraft with a lower specific fuel consumption.

A still further advantage of the present invention is that power take-off may be performed by taking drive at the periphery of the rotor disc, removing the requirement for the bevel gearing and shafting employed in current turbines, which must pass through the bypass duct via the duct bifurcation. This will increase the bypass duct area available for a given fan disc diameter, decreasing the specific fuel consumption of an aircraft equipped with such an engine. Furthermore, the elimination of the power takeoff shafting and associated gear trains removes losses due to gearing efficiency, lubrication inertia and friction in addition to reducing the dry mass of the engine.

Brief description of the Drawings

Fig. 1 depicts simplified cutaway schematic of a turbofan engine fan section designed according to the current state of the art, and embodying features that may be found in most common gas turbine engines.

Fig. 2 depicts a simplified schematic of a fan blade designed according to the current state of the art.

Fig. 3 depicts simplified schematic section of a turbofan engine.

Fig. 4 depicts a simplified cutaway schematic of a turbofan engine fan section designed according to the present invention.

Fig. 5 depicts a simplified schematic section of a turbofan engine fan section designed according to the present invention.

Fig. 6 depicts a simplified schematic section of the means of hub attachment of a turbine blade according to the present invention.

Fig. 7 depicts a simplified schematic section of the air bearing means of a fan blade disc according to the present invention.

Fig. 8 depicts a first stage compressor fan blade disc according to the present invention.

Fig. 9 depicts a detail view of the impeller ridges on a fan blade disc according to the present invention.

Specific Description

Fig. 4, Fig. 5, Fig. 6 and Fig. 7 show a turbofan aero-engine constructed according to the present invention. The fan blades (20) are carried by the fan outer ring (21) which runs in a groove formed in the fan casing (22). The fan outer ring incorporates a reservoir pocket (23) and impeller ridges (24) which enable a layer of pressurised air to be maintained between the fan casing (22) and fan outer ring (21), thereby forming an aerostatic fluid bearing. The impeller ridges (24) are angled such that a quantity of intake air is drawn into the bearing cavity, the relative dimensions of the aforementioned features being such that the layer of pressurised air imparts sufficient force on the fan outer ring (21) to support it during the entire operational envelope of engine conditions. The air pressure may be controlled, and a suitable flow of air supplied during starting operations via air ducts (25) which communicate with manifold (26). The air pressure for starting may be supplied from a modified air-starter system or from an external source such as an auxiliary power unit or the bleed air of another turbine in the case of a multi-engine vehicle. The bearing can be supplied with air at idle speeds from the engine core bleed air, should there be insufficient pressure generated by the impeller ridges (24). The use of a fluid-dynamic bearing in this manner allows for the boundary layer at the interface between the fan outer ring (21) and the fan casing (22) to be controlled using the exhaust air from the bearing, thereby enabling the prevention of boundary layer detachment and vortex formation. The fact that the fan blade (20) does not have a tip, but is blended into the fan outer ring (21) further enhances the flow efficiency through the bypass duct (27). In case of a loss of bearing air pressure, catch bearings (28) manufactured of a low-friction material such as boron nitride, PTFE or boron-aluminium-magnesium are positioned to support the fan outer ring (21) when the engine is stationary and will also allow is the turbine to be shut down safely, and will allow the engine to windmill, such that a multi-engine aircraft equipped with such engines may be safely flown to a suitable landing or ditching site.

The hub (29) is not required to withstand the centripetal loads generated by the rotation of the blades (20), as all such forces are taken compressively in the blades (20) and as a hoop stress in the fan outer ring (21). Due to the crystal structure of metals, their compressive strength is generally greater than their tensile strength, thus the blades (20) may be manufactured lighter or from cheaper materials than is possible with state of the art designs. Furthermore, the fan outer ring (21), which must be manufactured from a high strength material in order to withstand the hoop stresses imparted by normal running conditions acts as a retention band in the case of a blade-off event of the engine.

Torque is transferred to the blades (20) and fan outer ring (21) via the hub (29) which incorporates two offset rings of castellated mounting features (30), which allow the blades (20) to be simply fastened into place during assembly using bolts or some other form of fastener (31), followed by the fan annulus fillers (32) which incorporate sealing strips of an elastomeric material such as silicone or synthetic rubber. The hub is attached to the low pressure turbine shaft (33) of the engine core (34) via a curvic coupling (35) in the usual manner and covered with an aerodynamic spinner (36).

As shown in Fig. 8 and Fig. 9 the fan blades (20) employed in an engine designed according to the invention are manufactured as a part of the fan outer ring (21), by a means such as welding, diffusion bonding or adhesive bonding in the case of a composite fan blade disc. The blades (20) have an aerofoil section which may be tapered to widen at the fan outer ring interface, in order to carry the centripetal forces generated during engine operation. This has the effect of providing a large aerodynamic surface where the blade speed is greatest, enabling greater bypass mass flow rates to be achieved for a given blade mass. Retention lugs (37) at the core end (38) of the blades (20) allow them to be releasably attached to the hub (29). It should be noted that the complex shapes and tight tolerances required in the formation of a fir-tree root of a state of the art design engine are not required.

The retention lug (37) contains fewer high-tolerance features, thereby reducing the manufacturing cost associated with the blade (20).

It should be noted that the blade-off event debris energy is much lower in engines designed according to the present invention than in those according to the current state of the art. This is due to the distribution of material in the blades (20).

Instead of mass being concentrated at the blade root, the material is concentrated at the blade-outer ring interface. This causes the failure modes whereby one or more blades are released to be relatively benign compared to that of a state of the art engine. A fracture at the blade core end (38) will cause the blade to deform under aerodynamic loads, resulting in an out-of balance condition which may be safely carried by the bearings. A fracture at the fan outer ring (21) end will result in the weak core end (40) carrying the centripetal loads and also failing. The massive portion of the blade is already in contact with the fan outer ring (21), and the net result will be an out-of balance condition. This may, in extreme cases cause the fan outer ring (21) to deform and contact the catch bearing (28) in the fan casing (22), which will allow a portion of stored fan energy to dissipate as heat until the engine stabilises to a state where it can continue to rotate (windmill) freely albeit out of balance at a speed acceptable for an aircraft to fly on to destination.

Sound attenuation infill panels (38) are included in the engine nacelle (39) in line with current practice; however the reduced disturbance to the bypass airflow resulting from the elimination of blade tip-casing clearance gap will reduce the amplitude of noise emitted by the engine.

-10 -In order to provide power for accessories and for other aircraft systems a power take-off may be taken from the fan outer ring (21), enmeshed with which is a bevel gear driving a gearbox and electrical generator. This arrangement allows the extraction of energy from the engine without the need to obstruct the bypass duct with a driveshaft. Alternatively, a portion of the bleed air from the engine core (34) may be employed to drive an auxiliary free power turbine which is coupled to a generator and hydraulic pump.

The above description of the preferred embodiment has been given by way of an example. From the disclosures given, those skilled in the art will understand the invention and its advantages, and will also find apparent changes and modifications to the structures and processes disclosed. It is sought therefore to cover all such changes that lie within the scope of the invention, as defined in the appended claims and equivalents thereof.

Claims (14)

1. -11 -Claims 1. A rotor blade disc comprising: a plurality of radially extending blades disposed about the axis of the disc; a ring that supports the outer ends of the blades against the rotational forces exerted on the blades during rotation of the rotor; and a central shaft to which the inner ends of the blades are attached to permit the transmission of torque between the shaft and the blades.
2. 2. A rotor blade disc as defined in claim 1, wherein the transmission of torque is from the shaft to the blades.
3. 3. A rotor blade disc as defined in claim 1 or claim 2, wherein the ring comprises take-off means for coupling rotation of the rotor to other apparatus.
4. 4. A rotor blade disc as defined in any preceding claim, wherein the ring r serves as a shield for containing a blade-off event. C)
5. 5. A rotor blade disc as defined in any preceding claim, wherein the ring N' defines the inner race of a bearing.
6. 6. A rotor blade disc as defined in claim 5, wherein an outer race of the bearing comprises means for mounting an assembly of which the rotor blade disc forms part.
7. 7. A rotor blade disc as defined in claim 5 or claim 6, wherein the bearing is a rolling contact bearing.
8. 8. A rotor blade disc as defined in claim S or claim 6, wherein the bearing is an electromagnetic bearing.
9. 9. A rotor blade disc as defined in claim 5 or claim 6, wherein the bearing is a fluid dynamic bearing, which uses the working fluid of the rotor blade disc as a bearing medium.

   -12 -
10. 10. A rotor blade disc as defined in claim 9, further comprising catch bearings between the inner and outer races of the fluid dynamic bearing.
11. 11. A rotor blade disc as defined in claim 9 or claim 10, wherein the blades are fan blades for a turbofan engine.
12. 12. A turbofan engine comprising a rotor blade disc as defined in claim 11.
13. 13 A turbofan engine as defined in claim 12, wherein the rotor blade disc is surrounded by a fan casing, an inner wall of the fan casing defining the outer race of the fluid dynamic bearing, and the working fluid being air.
14. 14. A turbofan engine substantially as described herein with reference to any of Figures 4 to 9. r a)N