BLADED DISC

Abstract

A bladed disc for a rotating machine comprising a central disc that rotates about a central axis, the central disc having a series of blades arranged around its periphery; the blades have dovetail roots which engage with slots on the central disc; the bladed disc being configured so that there is a pre-loading force between the blades and the central disc such that each blade is forced away from the central axis of the bladed disc; and wherein the pre-loading force is equal or greater than 40% of the maximum centrifugal force applied to the blade during a flight cycle.

Claims

1. A bladed disc for a rotating machine comprising a central disc that rotates about a central axis, the central disc having a series of blades arranged around its periphery; the blades have dovetail roots which engage with slots on the central disc; the bladed disc being configured so that there is a pre-loading force between the blades and the central disc such that each blade is forced away from the central axis of the bladed disc; and wherein the pre-loading force is equal or greater than 40% of the maximum centrifugal force applied to the blade during a flight cycle.

2. The bladed disc as claimed in claim 1, wherein the pre-loading force is from 60 to 100% of the maximum centrifugal force applied to the blade during the flight cycle.

3. The bladed disc as claimed in claim 1, wherein the pre-loading force is applied by the insertion of a shim between the blades and the disc within the slot and configured so that the shim forces the blade away from the centre of blade.

4. The bladed disc as claimed in claim 3, wherein the shim has a dry film lubricant coating.

5. The bladed disc as claimed in claim 1, wherein the pre-loading force is the result of a deformation of a shape memory alloy.

6. The bladed disc as claimed in claim 1, wherein the pre-loading force is the result of a taper being applied to at least one of the blade and the central disc.

7. The bladed disc as claimed in claim 6, wherein the taper is present on the leading edge and the trailing edge such that at least one of the centre of the blade and the central disc protrudes beyond the leading edge and the trailing edge.

8. The bladed disc as claimed in claim 1, wherein the blade is further retained by a locking plate.

9. A gas turbine engine comprising a bladed disc according to claim 1.

10. The gas turbine engine as claimed in claim 9, wherein the gas turbine engine is a geared gas turbine engine.

11. A method of reducing the low cycle fatigue of a blade within a bladed gas turbine engine, comprising; inserting a shaped blade into a corresponding slot on a disc of a gas turbine engine, and inserting a shim between the shaped blade and the disc, so as to force the blade away from the centre of the disc of the gas turbine engine.

12. The method according to claim 11, wherein the shim is inserted so that produces a force between 40% and 100% of the maximum centrifugal force applied to the blade during a flight cycle.

Description

DESCRIPTION OF THE DRAWINGS

[0048] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0049] FIG. 1 is a sectional side view of a gas turbine engine;

[0050] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0051] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0052] FIG. 4 presents an example of a typical dovetail joint;

[0053] FIG. 5a shows the force diagram and the resultant forces associated with the flight cycle for the unloaded blade in a static state, and FIG. 5b shows the same for the case of the pre-loaded blade in a static state;

[0054] FIG. 6a shows the force diagram and the resultant forces associated with the flight cycle for the unloaded blade in a rotating state, and FIG. 6b shows the same for the case of the pre-loaded blade in a rotating state;

[0055] FIG. 7a shows the stress cycle at the edge of bedding for an unloaded blade according to the prior art over a number of flights and FIG. 7b shows the stress cycle at the edge of bedding for a pre-loaded blade according to the present disclosure;

[0056] FIG. 8a shows a means of applying the force via the use of a shim, and FIG. 8b shows a means of applying force through the use of a shape memory alloy;

[0057] FIG. 9 is a graph showing the effect of preloading of the blade against the equivalent flight cycle stress reduction.

DETAILED DESCRIPTION

[0058] Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying drawings. Further aspects and embodiments will be apparent to those skilled in the art.

[0059] FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0060] In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the core exhaust nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0061] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

[0062] Note that the terms “low pressure turbine” and “low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the “low pressure turbine” and “low pressure compressor” referred to herein may alternatively be known as the “intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

[0063] The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in FIG. 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

[0064] The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

[0065] It will be appreciated that the arrangement shown in FIGS. 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the FIG. 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in FIG. 2.

[0066] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

[0067] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0068] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core exhaust nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

[0069] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

[0070] Gas turbine blades 401 need to meet several mechanical integrity requirements of which Low Cycle Fatigue (LCF) life is one. Low cycle fatigue typically results from cycles of operation i.e., from the run up and run down of the engine during its operational service life. It is, therefore, desirable to reduce the LCF. One significant area for LCF comes from the dovetails 402 style blade root design, which has a typical failure location at the edge of bedding (EoB) 403; this is located to the upper end of the contact flank as shown in FIG. 4. It is from this location where cracks 404 typically initiate and propagate as a result of LCF. The reason for the cracking in this location is because the EoB location experiences very high stresses during the engine operation owing to the local geometry and contact effects. Furthermore, the disc slot width dilates during run up due to the centrifugal forces and contracts back to its original length during run down which leads to the blade root slipping on the contact flank. This repeated relative slipping of the blade root within the disc slot causes fretting damage along the contact flank which further makes the EoB vulnerable to cracking.

[0071] The mounting of the blades to the discs are shown in FIGS. 5a and 5b. FIG. 5a demonstrates the case for a prior art example that is unloaded. FIG. 5b on the other hand demonstrates the case for a blade having a pre-loaded force applied to it according to the present disclosure. FIGS. 5a and 5b present the case in which the blades are stationary, and as such there is no centrifugal force acting on the blade. In FIG. 5a whilst the blade is static there is no normal reaction force Rai on the loaded flanks of the blades, and there is no centrifugal force on the blade. There is also no frictional force Q.sub.a1. In the static situation as shown in FIG. 5b, a force is applied to the base of the blade F.sub.b1 and the opposing force is applied to the disc. This force results in the normal reaction force R.sub.b11 being applied the loading flank of the blade. There is also a frictional force Q.sub.b1.

[0072] FIG. 6a presents the forces acting on a prior art blade whist it is rotating, and centrifugal force is present. When the blade is rotating a reactionary force R.sub.a2 is applied to the flank of the blade along with a frictional force Q.sub.a2 and a centrifugal load P.sub.a2. FIG. 6b presents the same for the present disclosure in which a force is pre applied to the blade. When the blade is in operation and the disc is rotating then a normal reaction force R.sub.b2 is applied to the loading flank of the blade, due to the centrifugal force P.sub.b2. There is also the presence of small frictional force Q.sub.b2. FIG. 7a presents a graph of an example of the stress at the edge of bedding for the prior art example. FIG. 7b presents the stress at the edge of bedding for the blade having a preloaded force applied to it according to the present disclosure. The stress at the EoB in the pre-loaded example changes at a much smaller rate between S.sub.b1 and S.sub.b2 than cycling between S.sub.a1 and S.sub.a2, which is the case in the unloaded example shown in FIG. 7a. The effect of the pre-loading is that although the peak stress during operation increases by a small amount over what is observed without the pre-loading the degree of stress cycling on the blade is less as the blade does not return to 0 stress but remains at its pre-loaded amount. As such, the degree of cycle stress within the blade is reduced; this consequently will reduce the low cycle fatigue experienced by the blade and the disc. This reduction is seen in both the leading and trailing edges as well as along the flanks on both blade and disc. The bladed disc of the present disclosure also has the advantage of reducing the blade root flank slips and as such also protects the blade root contact surface.

[0073] Mathematically, the situations presented in FIGS. 5a and 5b can be considered as the following: the situation can be simplified by considering the case as a 2D section of a dovetail geometry; thus, not accounting for the minor differences that are present due to the 3D effects. Although this is presented for a dovetail the same would apply to a firtree geometry. Further, by ignoring the aerodynamic loads as secondary, and therefore a minor issue, as LCF is dominated by centrifugal force the problem can be simplified. A further simplification of the physics at cause is to let the complex flight cycle, where the engine goes through several speed ranges and dwell during a typical flight, be simplified to a 0 to max RPM cycle. Therefore, we can present the example as set out in FIG. 5a as:

[0074] Under static conditions i.e., at zero RPM, the blade root does not experience any centrifugal loads and the blade does not experience any significant stress in the root EOB location.

P.sub.a1=0  Eq1

S.sub.a1=0  Eq2

[0075] Then as the engine runs up to max revolutions (MTO), the blade experiences a centrifugal force given by:—

P.sub.a2=m*r*w.sub.2.sup.2  Eq3

[0076] At equilibrium, the contact forces (reaction force and friction force) on both the flanks resist the centrifugal force as shown in FIG. 2 where:—

P.sub.a2=2(R.sub.a2*cos β+Q.sub.a2*sin β)=2R.sub.a2(cos β+μ*sin β).  Eq4 {as Q.sub.a2=R.sub.a2*μ}

[0077] The stress at the edge of bedding S.sub.a2 depends on the forces at the contact and root geometry which can be written as:

S.sub.a2=R.sub.a2*f(a,d,α,β,μ)  Eq5

[0078] In the case of the example according to the present disclosure and shown in FIG. 5b. In this case an insert is incorporated between the blade and the disc; this is designed to exert a preload onto the blade dovetail root which is similar in magnitude to the max centrifugal force.

[0079] Under static conditions, we can set the insert to be designed to exert a preload F.sub.b1 which is k times the Max centrifugal load at build. So, under static condition the insert exerts the preload given by:

F.sub.b1=k.Math.Pa.sub.2 [where 0.4<k<1]  Eq6

[0080] Due to the preload exerted by the insert, the blade root now experiences a radially outward force which is resisted by the forces at the contact flank at equilibrium.

F.sub.b1=2(R.sub.b1*cos β+Q.sub.b1*sin β=2R.sub.b1(cos β+μ*sin β)  Eq7

[0081] This results in EoB stresses even under static conditions. However, as the blade roots typically crack at the EoB and the stresses fall rapidly in magnitude as we move away from the EoB, the edge of bedding stresses determine the LCF life of the root. Therefore, the difference in the stress distribution far field from the EoB is inconsequential WRT the LCF life of the root.

[0082] The stress at the edge of bedding S.sub.b1 depends on the forces at the contact and root geometry which can be written as:

S.sub.b1=R.sub.b1*f(a,d,α,β,μ)  Eq8

S.sub.b1=k*S.sub.a2 (from Eq 8,7,6)  Eq9

[0083] Mathematically, the situations presented in FIGS. 6a and 6b can be considered as the following: in the second condition of MTO the engine has run up to maximum speed and the centrifugal force on the blade has increased. As a consequence, the preload exerted at the bottom face of the blade root decreases. At max speed the preload is negligible and the total load on the blade is equal to the centrifugal load.

F.sub.b2=0  Eq10

P.sub.b2=P.sub.a2  Eq11

R.sub.b2=R.sub.a2  Eq12

S.sub.b2=S.sub.a2  Eq13

[0084] The reason that the shim is beneficial is that as the engine runs up and runs down during every flight cycle, the EoB stress cycles between S.sub.a1 & S.sub.a2 i.e., between 0 to Sa2 which induces low cycle fatigue are shown in the lower images of FIGS. 5a and 5b. In Case of 5b, the EoB stress, cycles between S.sub.b1 and S.sub.b2; this is based on Eq 9 and 13 and can be represented as cycling between ‘k.Math.S.sub.a2’ to ‘S.sub.a2’. In order to allow a back-to-back comparison with the cases as set out in FIGS. 5a and 5b lets convert this cycle to R-ratio of 0 i.e. 0-max equivalent. Using walker stress correction as bellow:

[00001]$\begin{array}{cc}{S}_{b\left(0-\max \right)}=\left(S_{a2}-k*S_{a2}\right)*{\left(1-\frac{k*S_{a2}}{S_{\mathrm{a2}}}\right)}^{0.5-1}=S_{a2}*\sqrt{1-k}& \mathrm{Eq14}\end{array}$

[0085] As the LCF life depends on the Stress amplitude and the effective stress amplitude in the example of the present disclosure is significantly lower than that of the prior art. Therefore, the example as presented in FIG. 5b has a significantly higher LCF life than that of FIG. 5a.

[0086] The force may be applied to the blade by the insertion of a wedge or shim into the gap between the blade and the disc, as shown in FIG. 8a. The shims can be made of steel or any other suitable material. The shims may be coated with a dry film lubricant (DFL). The insertion of the shims into the slot can be done in a number of different ways as would be apparent to the person skilled in the art. For example, the wedges may be hydraulically pressed into the slot. The shim may be applied to either side. Alternatively, the shims can be applied both sides of the blade.

[0087] Another means of applying the pre-loading is to use a shape memory alloy (SMA) material, as shown in FIG. 8b. In order to achieve this, the SMA is cold pressed below its transition temperature point. The SMA remains having its thinner profile and can be forced into the gap between the blade and the disc. When the SMA reaches its transition temperature it expands back to its original shape and creates the pre-loading force. The use of a low temperature SMA means that there will be no changes in the shape of the alloy during flight conditions. Any suitable SMA materials may be used as would be apparent to the person skilled in the art.

[0088] An alternative to this is to shape the dovetail or the slot within the disc. This would involve a taper in the profile of one or both of the firtree and/or the slot. This taper would increase the loading on the blade as it is forced into the slot. This taper may be a single taper extending across the width of the slot or the fir tree or the taper mat be on both sides, such that the centre of the slot or the firtree protrudes more than the edges.

[0089] The desired pre-load force to be applied to the blade is proportional to maximum centrifugal force during operation of the engine. Any preloading has been found to have a beneficial effect on the low cycle fatigue stress undergone by the component. Nevertheless, desirable results can be achieved for a pre-loading of at least 40% of the maximum centrifugal force applied to the blade during flight cycle. The maximum centrifugal force is determined using equation 3 above. IN this, the maximum revolutions are at take-off (MTO). Increasing the pre-load force will have the desirable effect of further reducing the life cycle fatigue stress on both the blade and the disc. Preferably the load is between 60-100%; this is because as you get closer to 100% the less movement there is of the blade. Pre-loading the blade is also beneficial as it seduces the effect of slippage of the blade and this lack of movement is beneficial to the lifespan of the blade.

[0090] The force applied to blade can be varied to differing effects as shown in FIG. 9 with varying benefit. Here the dashed line represents the benefit of the loading at 40% of the centrifugal load applied to the blade. As can be seen there is a greater benefit as the loading is further increased.

[0091] The concept may be applied to any blades used within a rotating machine, such as a turbine engine or turbo machinery. However, it is particularly suited to compressor and turbine blades in a turbine engine. In this case once the blade has been inserted into the disc and the pre-loading has been applied the locking plates or any other further retention devices can be added. It may be necessary to modify the base of the blade or the top of the slot so that the shim or the SMA insert may be inserted between the blade and the disc.

[0092] It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.