GB2202907A

GB2202907A - Cooled aerofoil components

Info

Publication number: GB2202907A
Application number: GB08707300A
Authority: GB
Inventors: John Philip Dabbs Hakluytt
Original assignee: SECR DEFENCE; UK Secretary of State for Defence
Current assignee: SECR DEFENCE; UK Secretary of State for Defence
Priority date: 1987-03-26
Filing date: 1987-03-26
Publication date: 1988-10-05
Also published as: GB8707300D0

Abstract

A film cooled aerofoil 1, such as a nozzle guide vane is internally cooled by fluid flowing from a minifold 5 to vortex chambers 7,8 extending spanwise along the component. The vorticity is created as cooling fluid is fed into the vortex chambers. The high Nusselt number for flow at the periphery of the vortex chambers ensures good heat transfer to the cooling fluid which is discharged close to the leading edge through rearwardly directed slots 11,12 providing a cooling film on the surface of the aerofoil that will drain any stagnation zone that might otherwise result from entrapment at the leading edge. In one preferred form (Fig 2), fluid from manifold 5 is also conveyed via passages 21 to a manifold 22 after which it is fed via passages 23,24 to the vortex chambers 7 and 8 to supplement the flow therein. The passages 20,23,24 may be arranged in spanwise arrays and slots 11,12 may be continuous or broken. <IMAGE>

Description

COOLED AEROFOIL COMPONENTS This invention relates to cooled aerofoil components that is cooled components of aerofoil form especially, though not exclusively, nozzle guide vanes and rotor blades in the turbine section of a gas turbine engine.

It is well established that turbine rotor blades and nozzle guide vanes may be cooled in order to protect them from the full severity of the inflowing gases to an extent sufficient to achieve a satisfactory service life. The need for cooling is obviously more pressing in the first stage of the turbine, particularly for the first stage nozzle guide vanes which are the first components in the turbine stage. The conditions encountered by the first stage nozzle guide vanes may be severly exacerbated by non-uniform approach flow. If there are present any hot regions within the turbine inlet gas field and should these remain stationary or near stationary as is likely, then those nozzle guide vanes involved can experience service conditions much more severe than their neighbours.

Although other methods of turbine cooling have been proposed, turbine cooling is conventionally achieved by use of air bled from the compressor stage. This cooling air is directed to the interior of the nozzle guide vane or rotor blade ard ultimately passes into the turbine gas flow through tne aerofoil surfaces o the end surfaces of the cooled aomponerltsO The cooling air may be utilized in several well established modes. It may be utilized to extract heat from the blade in its flow within the blade by convection or by impingement upon interior surfaces. Alternatively or additionally the exit air may be utilized to sheath the aerofoil section in a film of cooler air to protect it from direct contact with the turbine gas flow.

Despite these well established modes of aerofoil component cooling for gas turbine there is still a need for improved cooling methods in order to extend component service life by reducing material temperatures or alternatively to increase tolerance temperatures in order to permit increase in engine efficiency by use of higher cycle temperatures. One problem that has been encountered in the nozzle guide vanes of the first turbine section of a diesel fuel burning gas turbine engine is manifest as a spanwise line of corrosion adjacent the leading edge of the vane mainly apparent on the pressure-side surface but encroaching on the suction side surface also. It is postulated that this problem is caused by the presence adjacent the corrosion line of a region of recirculating, virtually entrapped gases.It should be noted that it would be difficult to avoid this problem simply by improved aerofoil design for there is generally some uncertainty as to the precise flow direction of the inlet air with respect to the chord line of the aerofoil component. One factor in this uncertainty might be lack of uniformity in the inlet flow for example.

It is an object of the present invention to provide an improved cooled aerofoil component which is devoid of the above mentioned problem of leading edge recirculatory flow.

This invention is a cooled component of aerofoil form comprising, within the component, a cooling fluid inlet manifold, an elongate vortex chamber extending along the span of the component, a plurality of feed passages each connecting the inlet manifold to the vortex chamber and each terminating at a discharge orifice in the peripheral wall of the vortex chamber which discharge orifice is eccentric with respect to the vortex chamber, and a near-tangential exit passage from the vortex chamber which is in the form of a spanwise slot, the spanwise slot breaking the surface of an aerofoil surface of the component in the proximity of its leading edge and being rearwardly directed at this station.

In use of an aerofoil component as claimed, a vortex flow is established within the vortex chamber under the impetus given by the incoming flow from the feed passages. This ensures a high Nusselt number for the flow at the periphery of the vortex chamber and thus ensures a good heat transfer to the cooling fluid therewithin achieved without excessive loss of kinetic energy. Moreover in a correctly dimensioned configuration the exit flow from the spanwise slot is of sufficient magnitude to drain any leading edge region of stagnation by entrainment thus inhibiting the establishment of recirculating flow at the leading edge region. The exit slot is rearwardly directed to promote this entrainment but also to promote an attached flow downstream of the exit thereby to ensure efficient film cooling of the downstream surface of the aerofoil.The exit slot should be as close to the leading edge as is possible commensurate with practicality.

It has been found that the aforemention problem of leading edge corrosion loss in first stage nozzle guide vanes has been significant on one side only of the vane - the pressure side. Thus the component of the invention need only provide slot exit flow to that one surface of the component to remidy this problem. However in other situations a different flow pattern might necessitate slot exit flow to the suction side - such as in a high angle of attack configuration. In a preferred arrangement a respective slot exit flow is provided to both the pressure surface and the suction surface in order to provide an aerofoil component tolerant of varying angles of attack.

In conventional aerofoil cooling practice the leading edge region of the aerofoil may be perforated with cooling air exit holes to promote cooling of the 'leading edge region. This practice may cause complex aggrivation of the leading edge recirculatory flow problem mentioned previously. The aerofoil component of this invention might be adequately cooled in the leading edge region by other means. In part heat is extracted from the leading edge region by the recirculatory flow of cooling fluid within the vortex chamber. The vortex chambers should be situated as close to the leading edge as is commensurate with the need for structural integrity, and for ease of manufacture.In one embodiment of the invention the or some of the feed passages pass through the leading edge region forward the vortex chambers on route thereto to provide additional cooling for the leading edge. Preferably in this arrangement there is present a spanwise leading edge chamber and there is spanwise stagger between the inflowing sections of the feed passages and the outflowing sections of the same to force spanwise flow in the leading edge chamber and hence increase heat extraction. There might still be some need for leading edge effusion cooling holes.

The invention is now described with reference to the drawings of which: Figure 1 is a transverse section of a single nozzle guide vane showing one embodiment of the invention; and Figure 2 is a transverse section of a single nozzle guide vane showing another embodiment of the invention.

The gas turbine section of an engine comprises an alternating sequence, along the axis of the engine, of stationary nozzle guide vanes and rotating rotor blades. Each nozzle section comprises an annulus of vanes each of aerofoil section. Usually bridging members link adjacent aerofoil sections at the ends. Each rotor section comprises an annulus of blades each of aerofoil section supported on a rotor disk within the rotor anntlus. The tip of each blade may or may not be linked to adjacent blade tips by a bridging member.

In Figure 1 a single nozzle guide vane is depicted in transverse section, at 1. The vane 1 has a leading edge 2, a pressure surface 3 and a section surface 4. Within the vane 1 there is an inlet manifold 5. This comprises a large passage extending from end to end of the vane. Two arrays of feed passages 6 issue from the inlet manifold 5. Each array comprises a spanwise stack of individual feed passages 6.

Adjacent the pressure surface 3 of the vane at a position near to the leading edge 2 there is a vortex chamber 7. This comprises a smooth bore spanwise chamber of near circular or elliptical section. A complementary vortex chamber 8 is present adjacent the suction surface 4 at a position near to the leading edge 2. Each of the vortex chambers 7 and 8 is linked by its respective stacked array of feed passages 6 to the inlet manifold 5.

Each of the feed passages 6 terminates at the peripheral surface of its respective vortex chamber 7 or 8, as a discharge orifice which is eccentric with respect to the vortex chamber.

Those feed passages 6 which lead to the pressure side vortex chamber 7 discharge substantially tangentially at the left side of the vortex chamber to establish a clockwise flow in the vortex chamber 7 depicted at 9. The feed passages 6 which lead to the suction side vortex chamber discharge substantially tangentially at the right side of the vortex chamber to establish an anti clockwise flow therein depicted at 10. Each of the vortex chambers 7 and 8 communicates vith a respective exit passage 11 and 12. These are each generally in the form of a slot which extends spanwise of the vane 1. They lead from the respective vortex chamber at the periphery thereof and at an angle such that they discharge preferentially the peripheral flow from the vortex chamber.The exit passages 11 and 12 break the surface of the pressure surface 8 and the suction surface 4 respectively at a position near to the leading edge 2 but sufficiently downstream thereof to be well placed to entrain flow which would otherwise develop associated vith the leading edge 2. The exit passages 11 and 12 are rearwardly directed at the exit point (with respect to the aerofoil) to promote adherence of the exit flow to the downstream aerofoil surface (this flow is shown at 13 and 14) and also to promote entrainment of the upstream air to drain the leading edge region.

The flow rate from the exit passages 11 and 12 should be sufficient to ensure adequate draining of the leading edge region and also to provide a measure of film cooling for the downstream surfaces. This flow rate is a function of the geometry of the cooling configuration as compressor delivery air pressure will not be easily variable. Consequently metered flow may be ensured by providing feed passages 6 and exit passages 11 and 12 of precise dimensions.

In the embodiment shown in Figure 1 heat is extracted from the leading edge region by conductive transfer to the cooling air within the vortex chambers 7 and 8. The flow within the vortex chambers is strongly recirculatory, but it also has some spanwise component consequent upon the spanwise spacing of feed passages. This flow configuration is adopted in order to secure a high Nusselt number at the chamber surface for good heat transfer.

Figure 2 shows a modified cooling configuration intended to provide enhanced internal cooling for the leading edge 2 of the vane 1. Save as identified below all features are the same as described with reference to Figure 1. In this modified cooling configuration there is present an array of leading edge feed passages 20 which lead from the inlet manifold 5. The array is depicted at 21 and comprises a spanwise stack of passages 20. Adjacent the leading edge 2 and forward of the vortex chambers 7 and 8 there is a leading edge manifold 22 which extends in a spanwise direction passages 20 lead into the manifold 20 and from this manifold 20 issue forth exit passages 23 and 24 each grouped in spanwise arrays and leading respectively to vortex chamber 7 and vortex chamber 8.

Adjacent exit passages 23 and 24 are located in common planes but there is spanwise stagger between the array of leading edge feed passages 20 and the arrays of exit passages 23 and 24 to ensure that there is spanwise flow in the leading edge manifold to secure an efficient heat transfer to the cooling air. The exit passages 23 and 24 serve a secondary feed passages for their respective vortex chambers 7 and 8 and they discharge therein in a manner which will reinforce the recirculatory flow therein. There is spanwise stagger between the feed passages 6 and passages 23 and 24 so that the secondary feed passages serve also to promote spanwise flow in the vortex chambers 8 and 9.The configuration of the embodiment shown in Figure 2 might be arranged such that the secondary feed passages can be produced by drilling with access for the drilling being available through the exit passages 11 and 12.

In the embodiments illustrated the exit passages 11 and 12 each comprise a single continuous slot. There could be advantage in adopting a broken slot configuration consisting of several axially aligned slots, in order to promote a better attachment of the discharge flow to the downstream surface of the aerofoil by interaction of adjacent discharge flows to form pairs of opposite handed vortices utilizing the Taylor-Coertler effect.

Figures 1 and 2 depict only the;rooling arrangements concerning the invention. Vanes or ator blades embodying the invention could well have additional-ronventional cooling arrangements for the rearward regions:thereof, which are not shown in the drawings. Alternatively the rearward regions could incorporate further vortex chambers and exit slots (similar to those described with reference to Figures 1 and 2) for cooling the rearward regions. 5he drawings show the invention as applied to a turbine nozlel guide vane. The invention may be applied to a rotor blade in a like or similar configuration, with no change in eFssential layout or geometry.

Claims

1. A cooled component of aerofoil form comprising, within the component, a cooling fluid inlet manifold, an elongate vortex chamber extending along the span of the component, a plurality of feed passages each connecting the inlet manifold to the vortex chamber and each terminating at a discharge orifice in the peripheral wall of the chamber which discharge orifice is eccentric with respect to the vortex chamber, and a near tangential exit passage from the vortex chamber which is in the form of a spanwise slot, the spanwise slot breaking the surface of an aerofoil surface of the component in the proximity of its leading edge and being rearwardly directed at this station.

2. A cooled component of aerofoil form as claimed in claim 1, comprising, within the component, two of said elongate vortex chambers one for the pressure surface of the aerofoil and one for the suction surface of the aerofoil, each of the vortex chambers being linked to its respective aerofoil surface by a respective exit slot.

3. A cooled component of aerofoil form as claimed in claim 1 or claim 2 comprising a leading edge manifold extending spanwise of the component and therewithin, the leading egde manifold being supply by feed passages from the inlet manifold and being drained by feed passages leading to the vortex chamber or chambers.

4. A cooled component of aerofoil form as claimed in claim 3 in which the leading edge manifold is drained by feed passages leading to the vortex chamber or chambers which feed passages are supplementary to direct feed passages linking the inlet manifold with the vortex chambers the configuration being such that discharge from these supplementary feed passages reinforces recirculatory flow in the vortex chamber.

5. A cooled component of aerofoil form as claimed in claim 4 in which there is spanwise stagger between the direct feed passages and the supplementary feed 'passages so as to generate spanwise flow components within the vortex chamber.

6. A cooled component of aerofoil form as claimed in any one of claims 3-5 in which there is spanwise stagger between the supply feed passages to the leading edge manifold and the draining feed passages so as to generate spanwise flow components within the leading edge mnaifold.

7. A cooled turbine section aerofoil component as claimed in claim 1 and substantially as hereinbefore described with reference to Figure 1 of the drawings.

8. A cooled turbine section aerofoil component as claimed in claim 1 and substantially as hereinbefore described with reference to Figure 2 of the drawings.