GB2149022A - Warpable guide vanes for turbomachines - Google Patents

Abstract

An air-cooled variable guide vane for effecting changes in the area of a turbine nozzle throat in a gas turbine engine. Structurally, the aerofoil 58 of the guide vane 48 comprises a load-bearing spine 70 and a thermally warpable skin 72, which includes the trailing edge and which is attached (e.g. by welds 74a,74b) to the spine 70 spanwise of each flank of the aerofoil, but is not attached to the rest of the guide vane. During operation of the engine, differential thermal expansion of the two flanks of the thermally warpable skin 72 with respect to each other relative to their cool state makes the trailing edge of the aerofoil move with respect to the leading edge to cause changes in the area of the turbine nozzle throat T, these changes being regulated by altering the pressure of cooling air using valves (65 Fig. 2, not shown) and directing it onto the interior of the thermally warpable skin 72 via holes 78 in an inner skin 60 of the aerofoil. <IMAGE>

Description

SPECIFICATION Guide vanes for turbomachines The present invention relates to guide vanes as utilised in turbomachinery, and particularly to air cooled variable guide vanes suitable for effecting variations in the area of a turbine nozzle in a gas turbine engine.

In a modern gas turbine aeroengine, it is important that the flow characteristics of the turbine be very carefully matched with those of the compressor to obtain the maximum efficiency and performance of the engine. If, for example, the turbine nozzle guide vanes allow too little flow, then a back-pressure builds up causing the compressor to surge; too much flow causes the compressor to choke. In either condition a loss of efficiency occurs very rapidly.

It will be realised that there is a wide variation in the flow conditions within an aeroengine during an operating cycle, for example between TAKE-OFF and CRUISE conditions; an even wider variation in flow conditions occurs between TAKE-OFF and HOLD conditions, "HOLD" being a low speed condition for a civil aircraft when it is being held in the "stack" by Air Traffic Control prior to landing at an airport. In order to ensure that the capacities of the compressor and turbine are accurately matched at all engine conditions, it has been previously proposed that the angles of the aerofoils of turbine nozzle guide vanes in the turbine be mechanically varied by pivoting the aerofoils about a pivot axis extending spanwise of the vane, thereby effecting variation of the turbine nozzle area to match the required flow of turbine gases therethrough.An example of such a proposal for a "variable geometry" nozzle guide vane is to be found in UK patent specification number 1,11 9,774, to which the reader is referred. It will be noted that the nozzle guide vanes disclosed therein are also supplied with cooling air.

Unfortunately, such mechanisms to vary turbine nozzle area by pivoting the nozzle guide vanes involve much additional complication and cost of the engine, and also require signficant additional maintenance. This is particularly so since the mechanisms must be designed cope with high operating temperatures caused by the turbine gases.

In contrast, the present invention provides a simpler, cheaper and more maintainable way of incorporating a variable geometry turbine nozzle guide vane in a gas turbine engine.

According to the present invention, a turbine nozzle assembly for a gas turbine comprises: (a) a plurality of air-cooled turbine nozzle guide vanes, each guide vane having a hollow aerofoil with a concave flank, a convex flank, a leading edge, and a trailing edge, the aerofoil including (i) a thermally warpable skin incorporating the trailing edge of the aerofoil plus at least a portion of each aerofoil flank contiguous with the trailing edge such that the thermally warpable skin has a concave flank and a convex flank, whereby during operation of the turbine, differential thermal expansion of the flanks of the thermally warpable skin with respect to each other relative to their cool state effects thermal warping of the thermally warpable skin, said thermal warping causing movement of the trailing edge of the aerofoil with respect to the leading edge thereof, whereby the outlet angle of the aerofoil is varied, and (ii) air cooling means adapted to differentially cool the flanks of the thermally warpable skin with respect to each other, thereby to modulate said differential thermal expansion; and (b) valve means operable to vary the amount of cooling air reaching the flanks of the thermally warpable skin and hence to regulate the outlet angle of the aerofoil.

The invention also includes the air-cooled variable guide vane of the above assembly.

Preferably..the construction and arrangement of the thermally warpable skin and the air cooling means are such that when the aerofoil is supplied with a predetermined maximum flow rate of cooling air at a predetermined high power condition of the gas turbine, said thermal warping causes a maximum outlet angle of the aerofoil to be attained, and when the aerofoil is supplied with a predetermined minimum flow rate of cooling air at a predetermined low power condition of the gas turbine, said thermal warping causes a minimum outlet angle of the aerofoil to be attained.

In order to define upper and lower limits of variation of the outlet angle of the aerofoil, the guide vane is preferably provided with appropriately arranged abutments for the thermally warpable skin to bear against.

Preferably, the aerofoil comprises an inner skin and outer skin, the outer skin including the thermally warpable skin. It is convenient if the air-cooling means comprises the inner skin of the aerofoil, the inner skin having holes therethrough for directing cooling air onto the inside surface of the thermally warpable skin, the holes being distributed so as to achieve the desired differential cooling of the flanks of the thermally warpable skin.

In a convenient embodiment the vane has mutually confronting platforms and the aerofoil incorporates a spine in which joins the platforms together and forms at least the leading edge of the aerofoil, the thermally warpable skin being attached to the spine spanwise of the aerofoil but not being attached to the platforms except through the spine.

The invention will now be described, merely by way of example, with reference to the accompanying drawings, in which: Figure 1 is a diagrammatic view of a gas turbine aero-engine shown partly sectioned to reveal the location of an assembly of turbine nozzle guide vanes according to the present invention; Figure 2 is an enlarged view of part of the section shown in Fig. 1, showing the turbine nozzle assembly in more detail; Figure 3 is a view on guide vane aerofoil section line Ill-Ill of Fig. 2, structure other than the guide vane aerofoil section being omitted; Figure 4 is a block diagram of a suitable control system for the invention; and Figure 5 is a simplified model of the guide vane aerofoil section to illustrate design thereof.

Referring to Fig. 1, a gas turbine aero-engine 2 of the large turbofan type has an outer cowling 4. The main features of the engine 2 hidden by cowling 4 are indicated in a general way by dashed lines and in addition a portion of the cowling is "cut away" to reveal a section through part of the engine in a vertical plane through the engine's axis of rotation.

Engine 2 represents a known design to which the invention is applied, the engine having an engine core 6, a bypass duct 8 defined between bypass duct casing 10 and the outer casing 1 2 of engine core 6, and an exhaust system including an exhaust bullet 14 at the rear of engine core 6, a core exhaust nozzle 16, and a final propulsion nozzle 18.

The bypass duct 8 is supplied with bypass air from a front fan 20 acting as a low pressure (LP) compressor which also supplies the engine core 6. Air supplied to engine core 6 passes along the flow path outlined therein, passing through intermediate and high pressure compressor sections 22 and 24 respectively, combustion chamber 26, and high, intermediate and low pressure turbine sections 28, 30 and 32 respectively. Fan 20 is driven from low pressure (LP) turbine section 32 via a first (inner) power transmission shaft 34, intermediate pressure (IP) compressor section 22 is driven from the IP turbine 30 via a second (intermediate) power transmission shaft 36, and the high pressure (HP) compressor section 24 is driven from the HP turbine 28 via a third (outer) power transmission shaft 38. These three compressor/turbine assemblies are called the L.P., I.P. and H.P. spools respectively.

Engine core 6 is suspended within cowling 4 and bypass duct casing 10 by means of fan outlet guide vanes 39 and various other well known types of suspension struts and linkages (not shown). The compressors, turbines and shafts are held concentrically suspended within engine core 6 by, inter alia, bearings 40, which in turn are supported by a conical web 42 connected to the engine core casing 44 by means of angularly spaced-apart struts 46 extending through the hollow interiors of IP turbine nozzle guide vanes 48. Vanes 48 are supplied with cooling air as explained later.

Besides shrouding struts 46, IP turbine nozzle guide vanes 48 perform the tasks of converting some of the energy stored in the turbines gases into kinetic energy for impulse transfer to the IP turbine 30 and also turning the gas flow through a suitable angle for correct entry to the turbine 30. H.P. turbine nozzle guide vanes 50 are provided for HP turbine 28 to perform the same two tasks. However IP turbine nozzle guide vanes 48 additionally have variable geometry in accordance with the invention.

Referring now to Figs. 2 and 3 also, the variable geometry nozzle guide vanes 48 have conventional radially inner and outer platforms 52 and 54 respectively. Each outer platform 54 is connected both to the shroud ring 49 of the HP turbine 28 through a forward extension 51, and to the surrounding portion of casing 44 by means of front and rear lugs or flanges 53, 55 respectively, which are integral with the platforms. These flanges 53, 55 seal against their neighbouring flanges in the circumferential direction, and against casing 44, to form an annular chamber 57 between the radially outer parts of vanes 48 and the casing.

Circumferentially adjacent platforms 52 and 54 of neighbouring vanes form the inner and outer walls of the turbine nozzle passages 56, which are defined between the aerofoils 58 of the vanes as shown in Fig. 3. The inlets of the turbine nozzle passages 56 are of course defined between the leading edges of aerofoils 58, the passages being convergent in the downstream direction. The throats (which are also the exits) T of nozzle passages 56 are defined between the trailing edges and the convex flanks of adjacent aerofoils. The areas of the throats T are varied by the variable geometry feature of aerofoils 58, which feature causes the outlet angle a of the aerofoils to change within predetermined limits.

Each vane aerofoil 58 of vanes 48 is double walled, consisting of an inner wall or skin 60 and an outer wall or skin 62. The outer skin 62 is structurally joined to and supported from the inner skin 60 by twin ribs 64 and 66 which bracket the leading edge and extend over the length of the span of the aerofoil. In fact, inner skin 60, together with the leading edge region 68 of outer skin 62 and ribs 64 and 66, are integrally cast with the inner and outer vane platforms 52 and 54 and form the main structural component or spine 70 of the aerofoil 58.

The remainder of the outer skin 62 provides the variable geometry feature of aerofoils 58 and comprises a thermally warpable skin 72 - hereinafter referred to as "warp skin" 72 - which defines the trailing edge of the aerofoil together with the major parts of its concave and convex flanks.

Warp skin 72 is attached to spine 70 along weld joint lines 74a, 74b extending spanwise of each vane flank. The welds may be made by electron beam or laser beam energy. Warp skin 72 fits closely to the platforms 52 and 54 at its radially inner and outer extremities, but is not attached to them, there being small expansion gaps 76 to allow for movement of the warp skin as detailed below. To avoid stress-induced cracking in the outer skin 62 of aerofoil 58, the ribs 64 and 66 which join the outer skin 62 to inner skin 60 are spaced from weld joint lines 74a, 74b respectively, and the inner and outer corners 71 at those joint lines, where the warp skin 72 joins spine 70, are radiused as shown in the magnified inset to Fig. 2. The warp skin 72 is itself fabricated from two sheet metal flank pieces and a trailing edge portion which are welded together along weld lines 73a and 73b.

The supply of cooling air to the vanes 48 will now be described.

Referring back to Fig. 1 for a moment, high pressure (HP) cooling air is ducted away from an offtake (not shown) about half-way along the HP compressor 24 and radially outwards of it.

Such offtakes for cooling air are well known in the art. From thence it passes between the inner wall (not shown) and the outer wall 44a (Fig. 2) of casing 44 outboard of the combustion chamber 26 and enters annular chamber 59 (Fig. 2) defined between the outer wall 44a of casing 44, outer platform extension 51 of vane 48, front fixing flange 53 of vane 48, and the radially outer surface of HP turbine shroud 49, which is indirectly fastened to the rear end of the inner wall (not shown) of casing 44. Air exits chamber 59 via passages 61 through casing 44 and enters a further annular chamber 63 on the exterior of casing 44 surrounding the guide vanes 48.

The HP cooling air exits from chamber 63 via valves 65 in casing 44 to the chamber 57 between vanes 48 and casing 44, and from thence passes into the hollow interiors of the vanes.

Although valves 65 are provided in order to control the supply of cooling air in accordance with the invention, it should be realised that appreciable "leakage" of cooling air occurs direct from chamber 59 to chamber 57 past front fixing flanges 53 of vanes 48. This leakage is wanted in the design, because a continuous feed of cooling air through the interior of vanes 48 and into the interior of the engine is necessary in order to pressurise the vanes and their inner platform seal with the IP turbine 30 against unwanted ingress of turbine gases. However, this continuous feed of cooling air, derived from continuous "leakage" into chamber 57, is at a significantly lower pressure than the HP cooling air in chambers 59 and 63.Consequently, when valves 65 are partially or completely opened, the cooling air pressure within chamber 57 and vanes 48 is increased, thus increasing the cooling effect on the vanes.

It should be noted that there is no need for each vane 48 to have a corresponding valve 65, provided that the valves are equally spaced around the circumference of the casing 44 and that their number is adequate to supply sufficient cooling air to chamber 57 to achieve the degree of cooling of the vanes required by the invention. The valves 65 are further described later.

The invention relies for its controlled operation upon variation of the supply of cooling air by the valves 65, the air escaping from the interior of each vane through a plurality of rows of small holes 78 in inner aerofoil skin 60 and cooling the outer skin 62 by impinging as jets of cooling air on its interior surface. These holes 78 are distributed over the extents of the concave and convex flanks in such a way as to achieve an approximately uniform temperature over each individual flank of the outer skin 62, but many more holes 78 are provided in the concave flank of inner skin 60 than in its convex flank, so that, when necessary in order to thermally warp the warp skin 72 as described later, the concave flank of outer skin 62 can be cooled to a temperature below that of a convex flank.

It should be noted that the cooling air which is utilised to cool warp skin 72 escapes from the space between inner skin 60 and warp skin 72 through the row of holes 79 in the convex flank of warp skin 72 upstream of weld line 73a and also through the small expansion gaps 76. The disturbing effect of the injection of cooling air into the turbine nozzle passages 59 through expansion gaps 76 will be minimised by the fact that such injection takes place in the corners of the passages where the boundary layer is thickest.

Describing first the operation of the embodiment of Figs. 2 and 3 in general terms, the function of the inner skin 60 is to differentially cool the flanks of the warp skin with respect to each other and thereby modulate the differential thermal expansion experienced by the flanks with respect to each other relative to their cool (ambient temperature) state. The IP turbine nozzle area is varied as desired during the engine's operating cycle by using valves 65 to vary the amount of cooling air reaching the two flanks of the warp skin 72 of each aerofoil 58, such that the differential thermal expansion between those two flanks relative to their cool state is regulated to cause a predetermined amount of warping of the warp skin 72 as a whole.This warping is such that the trailing edge of each aerofoil 58 moves with respect to the leading edge to change the outlet angle a of the aerofoil, the movement being across the direction of flow in the turbine nozzle passages 56, thereby changing their throat areas.

More specifically, the aeroengine is required to operate at two widely dissimilar engine conditions, viz. at the TAKE OFF and the HOLD conditions. It is also required to operate at an intermediate CRUISE condition which is nearer to TAKE OFF than HOLD. At TAKE OFF, maximum engine power is required, so the turbine blades and vanes experience maximum gas temperatures, and the maximum amount of work is required from the IP and other turbines; therefore the throat areas of the nozzle passages 56 should ideally be at a minimum consistent with an adequate surge margin for the compressor.At HOLD, minimum engine power is required, so the turbine blades and vanes experience minimum gas temperatures, and the minimum amount of work is required from the IP turbine, not only because the engine power requirement is less, but also because it is desirable to speed up the HP spool relative to the IP spool so that the IP compressor has a greater surge margin at this "throttled back" condition; therefore IP turbine nozzle throat areas should ideally be at a maximum. At the intermediate CRUISE condition an intermediate throat area is ideally required.

Our investigations show that if for example the warp skin 72 is constructed of the nickel-base sheet material known in the industry as N263. the appropriate warp variations for achieving the above-mentioned nozzle passage throat area variations can be obtained using acceptable and easily 'attainable rates of differential cooling of the two flanks of the warp skin. Thus, considering the TAKE OFF condition, the concave flank of the aerofoil 58 experiences the greatest heating effect from the turbine gases because of their high static temperatures on that flank. However, the valves 65 are held fully open at TAKE OFF so that the temperature of the concave flank of the warp skin 72 is reduced to a valve significantly below that of the convex flank by virtue of the greater number of cooling air holes 78 on the concave side.Hence, the concave flank expands less than the convex flank relative to their cool state and the warp skin 72 as a whole is warped relative to its cool state so that its trailing edge moves towards the convex flank of the neighbouring vane and minimises the nozzle passage throat area. In order to ensure the invariable attainment of a correct desired minimum throat area at TAKE OFF, the amount of warping is such that at the engine condition in question the inside of the convex flank of the warp skin 72 abuts the trailing edge of the inner skin 60 (which is part of structural spine 70), thus providing the throats T with a "closed" position limit stop. This is shown in Fig. 3.

Considering now the HOLD condition, the cooling air flow through valves 65 is reduced or stopped completely such that the greater heating effect of the turbine gases on the concave side raises its temperature above that of the convex side. Hence, the concave flank expands more than the convex flank relative to their cool state and the warp skin 72 as a whole is warped relative to its cool state so that its trailing edge moves away from the convex flank of the neighbouring vane and maximises the nozzle passage throat area. The spine 70 provides the throats with an "open" position limit stop at the HOLD condition in a way which is analogous to the "closed" position limit stop at the TAKE OFF condition, i.e. the amount of warping at HOLD is sufficient to cause abutment between the inside of the concave flank of warp skin 72 and the trailing edge of inner skin 60.

In order to prevent vibration of warp skin 72 against spine 70, the amount of warping at the TAKE OFF and HOLD conditions is such that there is "interference" between the warp skin 72 and spine 70 at the abutment (which is of course a line contact extending spanwise of the aerofoil); i.e. the warp skin 72 is sprung against inner skin 60 because the abutment prevents the differential expansion between the two flanks of skin 72 from warping the skin to the fullest extent possible.

Finally, at the CRUISE condition the cooling air flow to warp skin 72 could, if desired, be reduced relative to that necessary at TAKE OFF so as to reduce the temperature difference between the concave and convex flanks and hence reduce the amount of warping of the warp skin, whereby the trailing edge of the warp skin takes up a position intermediate the "throat closed" (TAKE OFF) and "throat open" (HOLD) positions, the internal surfaces of the warp skin being clear of the inner skin 60.Alternatively (and preferably, if vibrational "flap" of the warp skin 72 is a problem due to excitation by either turbulence in the turbine gas flow or pressure waves from the IP turbine blades), the cooling air flow during CRUISE may be reduced by partially closing valves 65 to take account of the reduced heating effect from the turbine gases at this condition, but is nevertheless maintained at a flow rate sufficient to ensure that the warping induced in the warp skin 72 keeps it sprung against the spine 70 in the "throat closed" position. In fact to simplify design and operation of the system to the maximum extent, it may be preferable to have valves 65 which have only two positions, namely fully open for TAKE-OFF and CRUISE and fully closed for the HOLD. In the two former cases, the "closed" throat area of the nozzle passages will be judiciously selected to give a high efficiency at CRUISE, a lower but acceptable efficiency at TAKE OFF, and an acceptable surge margin at both conditions.

Our investigations show that using the above technique, it is feasible to open the areas of the turbine nozzle throats T by at least 10% relative to the throat closed position, thereby achieving worth-while fuel savings during HOLD. Optimisation of the throat areas may also improve fuel consumption at the other engine throttle conditions.

Control of the cooling air supply valves 65 need involve only well known design expedients as indicated in Fig. 4. If it is desired only to have only fully open and closed positions of valves 65, they may be of the solenoid type activated by suitably positioned micro-switches (not shown) on the pilot's throttle lever position selector 82, this activation on being via an electronic control unit 84 linked to the solenoid valves on line 86. On the other hand, if it is desired to have continuously variable valves 65, the system will be more complex, involving, for example, rotary-type valves actuated by small stepper motors, the control unit 84 sensing throttle lever position by means of e.g., a rheostat on selector 82 and emitting corresponding control pulses on line 86 according to a schedule programmed or wired-in to the control unit.In this way the flow rate of cooling air is varied as appropriate to give the desired IP turbine nozzle area for the engine condition being experienced.

In connection with the subject of cooling air supply rates, it should be remembered that increases in cooling air consumption lead to increased fuel consumption, hence it is desirable to .minimise the amount of cooling air consumed by vanes 48. The invention helps towards this objective in that although a high pressure supply of cooling air to the vanes 48 through valves 65 is necessary at TAKE OFF, this supply can possibly be reduced at CRUISE and eliminated or at least further reduced at HOLD.

In connection with the subject of the impact of compressor bleed air offtakes on surge margins, it should be noted that although at the HOLD condition the consumption of HP compressor air for cooling of the vanes 48 is at a minimum - a fact which could be expected to reduce the surge margin - no problem will arise because at HOLD the HP turbine gas entry temperature is lower than at other conditions, the HP compressor working line thereby also being lowered; this offsets any reduction of surge margin caused by minimising usage of HP cumprG'ssor bleed air.

In order to illustrate some design aspects of the invention, some exemplary calculations will be performed for the simplified model of the invention shown in Fig. 5. There, warp skin 72 is represented as having a triangular form in chordal cross-section, being joined to a rigid nonwarping spine (not shown) at points A and B between which is defined the base dimension AB of the triangle and from which extend the two flanks AE and BE of the warp skin 72, these representing the "convex" and "concave" flanks respectively.A right triangle BCE is con structed on "concave" flank BE and contains complementary angles a and 8. For the geometry shown it is convenient to utilise angle a to assess the amount of movement at the trailing edge E due to thermal warping of the warp skin 72; a is in fact the outlet angle of the aerofoil, since CE is aligned to the axial direction of the turbine. The value of angle a will be calculated at three engine conditions namely the COOL "as assembled" condition, at TAKE OFF, and at HOLD.

Dimensions are shown in mm at the COOL condition (here taken as 0 C for simplicity) and are representative of dimensions in a large turbofan aeroengine such as the Rub211 series manufactured by Rolls-Royce Limited. The width of the nozzle throat between trailing edge E and the convex side of the neighbouring vane is assumed to be about 23 mm.

(i) COOL Let distance BC = b and distance CE = c Hence 114,32 = b2 + c2 and 130.812 = (20.32 + b)2 + c2 Substituting, 130.812 = (20.32 + b)2 + 1 14.32 - b2 Therefore, b = 130.812 114.3022O.322 =89.4156 2 X 20.32 89.4156 Hence, a = Sin ~ ' = 51.472 114.30 (ii) TAKE OFF Assumed temperatures:Side AE = 825"C Side BE = 750"C Side AB = 795"C For N263 nickel-base material, effective linear coefficients of temperature expansion for the above, sides AE, BE, AB between their respective temperatures and 0 C are as follows (the coefficient increases with temperature): Side AE = 0.00001620 Side BE = 0.00001546 Side AB = 0.00001595 Therefore the new lengths are: Side AE (0.00001620 X 825 x 130.81)+ 130.81 = 132.5575 Side BE (0.00001546 X 750 X 114.30) + 114.30 = 115.6254 Side AB (0.00001595 x 795 x 20.32) + 20.32 = 20.5776 132.55752 - 115.62542 - 20.57762 b= = 91.8185 2 X 20.5776 and 91.8185 a = Sin~ 1 = 52.595 115.6254 i.e. at TAKE-OFF the throat width is reduced relative to COLD.

From the change in a between COOL and TAKE-OFF, it is possible to calculate the resulting change in position of trailing edge E and hence the change in the width of the throat.

Thus, movement of E 11 5.6254 X (change in 0) Change in 8 # change in a = 52.595' - 51.472' = 1.123 Therefore, movement of E # 115.6254 X 1.123x 180 = 2.2663 mm This is approximately a 10% reduction in the width of the throat, or slightly less allowing for some "interference" between the warp skin and the "throat closed" stop on the inner aerofoil skin (Fig. 3).

(iii) HOLD Assumed temperatures: Side AE = 750 C Side BE = 750 C Side AB = 795 C Therefore, the new lengths are: Side AE (0.00001546 X 750 X 130.81 = 132.3267 Side BE =115.6254 Side AB = 20.5776 132.32672 - 115.62542 - 20.57762 b= = = 90.3333 2 X 20.5776 90.3333 a = Sin ~ ' = 51.3758 115.6254 Change in a (and 0) between COLD and HOLD = 51.472' - 51.3758' = 0.09620 = 0.00168 radians If at O"C the warpskin 72 is just on the "throat open" stop and if the abutment between the inner skin 60 and warp skin 72 occurs 77 mm from B along BE, then the nominal interference between inner skin 60 and warp skin 72 at HOLD is given by 77X0.00168=0.129 mm Note that a different value of interference at HOLD (or TAKE OFF) could easily be achieved by, for example: (a) altering the distance from B (or A) of the point of abutment which provides the "throat open" (or "throat closed") stop; and/or (b) modifying the cooling air flow impinging on side AE or side BE in order to change their temperatures.

It will now be apparent that appreciably more or less than 10% variation in throat area would be available as a matter of design choice.

The amount of variation in throat area which is achievable for any particular overall size of vane is influenced by, for example: (a) the lengths of sides AE and BE of the warp skin.

(b) the changes in temperature of sides AE and BE between the engine conditions of interest-this is dependant upon cooling air flows; (c) the materials from which the vane is made - in this connection it is pointed out that the above-mentioned N263 material is only exemplary and that other materials with different expansion rates could be utilised. It would even be possible to make each side AE and BE of a different material with different thermal expansion coefficients, so that the possibilities for controlling and tuning the nozzle throat areas in accordance with engine condition become very wide.

Claims (9)

1. A turbine nozzle assembly for a gas turbine, comprising: (a) a plurality of air-cooled turbine nozzle guide vanes, each guide vane having a hollow aerofoil with a concave flank, a convex flank, a leading edge, and a trailing edge, the aerofoil including (i) a thermally warpable skin incorporating the trailing edge of the aerofoil plus at least a portion of each aerofoil flank contiguous with the trailing edge such that the thermally warpable skin has a concave flank and convex flank, whereby during operation of the turbine, differential thermal expansion of the flanks of the thermally warpable skin with respect to each other relative to their cool state effects thermal warping of the thermally warpable skin, said thermal warping causing movement of the trailing edge of the aerofoil with respect to the leading edge thereof, whereby the outlet angle of the aerofoil is varied, and (ii) air cooling means adapted to differentially cool the flanks of the thermally warpable skin with respect to each other, thereby to modulate said differential thermal expansion; and (b) valve means operable to vary the amount of cooling air reaching the flanks of the thermally warpable skin and hence to regulate the outlet angle of the aerofoil.

2. An air-cooled variable guide vane for a turbine nozzle of a gas turbine, the guide vane having a hollow aerofoil with a concave flank, a convex flank, a leading edge, and a trailing edge, the aerofoil including (i) a thermally warpable skin incorporating the trailing edge of the aerofoil plus at least a portion of each aerofoil flank contiguous with the trailing edge such that the thermally warpable skin has a concave flank and a convex flank, whereby during operation of the turbine, differential thermal expansion of the flanks of the thermally warpable skin with respect to each other relative to their cool state effects thermal warping of the thermally warpable skin, said thermal warping causing movement of the trailing edge of the aerofoil with respect to the leading edge thereof, whereby the outlet angle of the aerofoil is varied, and (ii) air cooling means adapted to differentially cool the flanks of the thermally warpable skin with respect to each other, thereby to modulate said differential thermal expansion.

3. An air-cooled variable guide vane according to claim 2, in which the construction and arrangement of the thermally warpable skin and the air cooling means are such that when the aerofoil is supplied with a predetermined maximum flow rate of cooling air at a predetermined high power condition of the gas turbine, said thermal warping causes a maximum outlet angle of the aerofoil to be attained, and when the aerofoil is supplied with a predetermined minimum flow rate of cooling air at a predetermined low power condition of the gas turbine, said thermal warping causes a minimum outlet angle of the aerofoil to be attained.

4. An air-cooled variable guide vane according to claim 2 or claim 3 having abutments for the thermally warpable skin to bear against, the abutments being arranged to define upper and lower limits of variation of the outlet angle of the aerofoil.

5. An air-cooled variable guide vane according to any one of claims 2 to 4, in which the aerofoil comprises an inner skin and an outer skin, the outer skin including the thermally warpable skin.

6. An air cooled variable guide vane according to claim 5, in which the air-cooling means comprises the inner skin of the aerofoil, the inner skin having holes therethrough for directing cooling air onto the inside surface of the thermally warpable skin, the holes being distributed so as to achieve said differential cooling of the flanks of the thermally warpable skin.

7. An air-cooled variable guide vane according to any one of claims 2 to 6 in which the vane has mutally confronting platforms and the aerofoil incorporates a spine which joins the platforms together and forms at least the leading edge of the aerofoil, the thermally warpable skin being attached to the spine spanwise of the aerofoil but not being attached to the platforms except through the spine.

8. A turbine nozzle assembly substantially as described herein with reference to and as illustrated by Fig. 2 of the accompanying drawings.

9. An air-cooled variable guide vane substantially as described herein with reference to and as illustrated by Figs. 2 and 3 of the accompanying drawings.