US20050089401A1

US20050089401A1 - Turbine blade tip clearance system

Info

Publication number: US20050089401A1
Application number: US10/914,077
Authority: US
Inventors: Anthony Phipps
Original assignee: Individual
Current assignee: Individual
Priority date: 2003-08-15
Filing date: 2004-08-10
Publication date: 2005-04-28
Also published as: GB2404953A; GB0319180D0

Abstract

A stage of gas turbine engine turbine blades (22) is surrounded by segments (26) that in turn, are suspended from a flexible casing (24). The inner surface of casing (24) is exposed to compressor flow. On engine thrust being increased, compressor output also increases and rapidly flexes casing (24) radially outwards, thus equally rapidly moving the segments (24) away from the tips of the turbine blades (22).

Description

The present invention relates to gas turbine engine blade tip clearance systems, wherein such an engine, the outer wall of the gas annulus that surrounds a stage of turbine blades comprises segments that are moveable relative to the blades in directions radially of the engine axis. The arrangement enables a reduction in blade tip rub on the gas annulus wall when the blades and associated turbine disk grow under the influence of heat and centrifugal forces. Also, on slowing and cooling of the assembly, the arrangement enables the spacing of the segments from the blade tips by a distance that reduces performance losses.
Known art utilises active devices i.e. complicated sensing devices for sensing the relative movement between blades and segments, which devices, on sensing a change in gap magnitude, develop signals that are passed to actuating mechanisms. These, in turn actuate segment moving means, and thereby move the segments in an appropriate direction.
Devices of the kind generally described hereinbefore have drawbacks over and above their complicated system of operation. They are heavy, which generates weight penalties where the associated engine is utilised in an aircraft. They are expensive to manufacture, and further, they cannot react with appropriate efficiency so as to cater for both the rapid growth of the disk and blade assembly during engine acceleration, and its much slower reduction in size on deceleration and cooling.
The present invention seeks to provide an improved turbine blade tip clearance system.
According to the present invention a turbine blade tip clearance system comprises a rigid outer casing and an inner casing having induced flexing capability supported by the outer casing in radially spaced relationship therewith so as to define an annular space therebetween, and wherein said inner casing supports a ring of segments within it in fixed radial relationship therewith, such that on placing of the whole around a stage of disk mounted turbine blades in co-axial relationship therewith, said segments will lie in radially close spaced relationship with the tips of said turbine blades.
The present invention will now be described, by way of example and with reference to the accompanying drawings in which:
FIG. 1 is a diagrammatic sketch of a gas turbine engine including a turbine blade tip clearance system in accordance with the present invention.
FIG. 2 is an enlarged cross sectional part view of the turbine blade tip clearance system of FIG. 1.
Referring to FIG. 1. A gas turbine engine indicated generally by the numeral 10, comprises a compressor 12, combustion equipment 14, a turbine section 16, and an exhaust duct 18. Turbine section 16 includes a rigid outer casing 20 that surrounds a stage of turbine blades 22. Casing 20 supports an inner casing 24, and inner casing 24, in turn, supports a ring of arcuate segments 26, the radially inner surfaces 28 of which lie closely adjacent the tips of turbine blades 22.
Referring now to FIG. 2. In the example, casing 24 is conical in form and is supported via its respective beaded ends 30 and 32 in grooves formed in internal flanges 34 and 36 on casing 20. Pairs of bosses 38 and 40, only one pair of which is shown, are angularly spaced around inner casing 24, each pair being arranged in axial alignment relative to the axis of engine 10. Bosses 38 and 40 are drilled and tapped so as to accept respective screw threaded bolts 42 and 44, the heads 46 and 48 of which are shaped so as to enable their location in forked ends of brackets 50 and 52 that extend from the radially outer surface of each segment 26. Segments 26 will each have the same number of axially aligned brackets 50 and 52, as there are bosses. Stopper bars 54 are fixed to flanges 34 and 36, each side of bosses 38 and 40, so as to prevent excessive radially outward flexing of inner casing 24 during operation of engine 10, as is explained later in this specification.
Assembly of the whole can be achieved by first screwing bolts 42 and 44 through respective bosses 38 and 40, in a direction radially outwardly of casing 24, followed by inserting the grooved bolt heads 46 and 48 in the respective forked ends of segment brackets 50 and 52. Nuts 56 and 58 are then screwed on to the extremities of the respective projecting ends of bolts 42 and 44, but not tightened against their respective bosses. The now loosely juxta positioned segments 26 are slid onto the land of a disk shaped jig 60, the diameter of which corresponds to the diameter of the stage of turbine blades 22, plus a cold clearance margin i.e. the required clearance 62 between the tips of blades 22 (shown by a dashed line) and the adjacent inner surfaces of segments 26, when associated engine 10 is inoperative.
It is preferable that jig 60 is supported for rotation about its axis, so that the loosely fitted segments at the bottom of the assembly can, in turn, be brought to the top to ease the positioning of the respective parts.
The positioning of the loosely assembled parts is achieved as follows: Spacers (not shown) of appropriate length are placed between the undersides of the bolts heads and the opposing inner end faces of their respective bosses. The bolts are then screwed further through bosses 38 and 40 until the spacers are lightly trapped between respective bolt heads and bosses inner faces. Each spacer (not shown) is then removed. Nuts 56 and 58 are then screwed along their respective bolts so as to engage the outer ends of their respective bosses 38 and 40. These steps are repeated all around the assembly and results in the clamping of all of the segments 26 in co-axial relationship with inner casing 24 and, nominally, in a desired spaced relationship with the tips of the stage of turbine blades 22 around which the assembly is to be fitted.
The assembly as described so far is now removed from jig 60, and fitted into casing 20 by inserting the beaded edge of what will be the upstream end of casing 24 with respect to the flow of gases through engine 10 (FIG. 1) in the annular groove in flange 34. The downstream beaded end of casing 24 is then inserted in the groove in flange 36, which is then offered up to a small flange 64 within casing 20, and bolted thereto. The completed assembly is then fitted over the pre-assembled stage of turbine blades. Should further adjustment prove necessary, due to manufacturing tolerances of the various parts, this can be achieved before fitting casing 20 to the casing of combustion equipment 14 (FIG. 1) upstream thereof, by accessing nuts 56 and 58 through normally capped apertures 66 and 68 in casing 20.
Gas turbine engine 10 (FIG. 1) is of the kind utilised to power aircraft. The power regime of engine 10 embraces aircraft taxiing, take off, cruise and landing, all of which require different engine power outputs. Thus, gas and air temperatures and pressures, and speed of revolution of rotary parts change in concert. During taxiing of the associated aircraft (not shown), no adverse temperature is experienced by the turbine system. However, when the throttle is opened to obtain full power so as to enable takeoff, there occurs an almost instantly considerable rise in compressor output pressure, combustion gas temperature, and speed of revolution of the turbine section.
The combination of increased temperature and speed of revolutions of the turbine stage causes the latter to increase its diameter through centrifugal forces and increased heat. The magnitude of the increase is such that, if the ring of segments 26 (a common inclusion in gas turbine engine turbine systems) remains in its cold position, or moves too slowly radially outwards with respect to the turbine stage, severe rubbing of the tips of the turbine blades on the segments would occur, with consequent loss of material from blade tips and segments. The resultant permanently increased annular gap therebetween would cause severe performance losses over the entire operating regime of the engine. In the present arrangement however, the increased pressure output from the compressor 12 (FIG. 1) enables a portion of that output to be immediately diverted to the radially inner surface of inner casing 24 by any known ducting (not shown) so as to induce flexing of casing 24 radially outwards against the stopper bars 54, and by this means, rapidly lift segments 26 away from the tips of the turbine blades. The time lag between the start of turbine growth and flexing of casing 24 is so small as to minimise any rubbing action that may occur, and consequently reduces efficiency losses.
When the engine 10 is throttled back e.g. when the associated aircraft reaches its cruise altitude, the turbine stage 22 contracts more slowly than it expands. Compressor output pressure also reduces and consequently reduces the force exerted on casing 24, which then could return too quickly to its non flexed shape and so cause rubbing between segments 26 and the tips of blades 22. Again, the present arrangement provides means to avoid rubbing through contraction, by making casing 20 from a material, the magnitude and rate of expansion and contraction of which can be controlled by heating and cooling. The nickel alloy marketed under the registered trade mark “Waspaloy” is one such material.
On throttling back of engine 10, with consequent reduction in pressure on casing 24, hot compressor air could be ducted (not shown) onto casing 20 so as to rapidly heat it and cause it to expand in a radially outwards direction. The movement is transmitted via flanges 34 and 36, to the sub assembly of casing 24 and segments 26, and has the effect of slowing its rate of radially inward movement to a rate more compatible with that of the turbine stage. Rubbing of segments 26 on the blade tips during contraction is thus minimised.
Reference to FIG. 2 shows that casings 20 and 24, and flanges 34 and 36 define an annular chamber 70, which will fill with engine leakage air. It is important that the leakage air pressure in chamber 70 is prevented from reaching a magnitude such that casing 24 is flexed radially inwards, or is prevented from being flexed radially outwards at appropriate speed as and when required. To this end, flange 36 has at least one vent hole 72 through its thickness.

Claims

1. A turbine blade tip clearance system comprising a rigid outer casing and an inner casing having induced flexing capability supported by the outer casing in radially spaced relationship therewith, so as to define an annular space therebetween, and wherein said inner casing supports a ring of arcuate segments within it in fixed radial relationship therewith, such that on placing the whole around a stage of disk mounted turbine blades in coaxial relationship therewith, said segments will lie in radial close spaced relationship with the tips of said turbine blades.

2. A turbine blade tip clearance system as claimed in claim 1 wherein said arcuate segments are suspended from said inner casing by screw threaded fastening means so as to enable radial positional adjustment of said arcuate segments with respect to a selected datum.

3. A turbine blade tip clearance system as claimed in claim 1 wherein said outer casing is manufactured from a temperature sensitive material.

4. A turbine blade tip clearance system as claimed in claim 3 wherein said material comprises nickel.

5. A turbine blade tip clearance system as claimed in claim 1 when installed in a gas turbine engine.

6. A gas turbine engine including an installed turbine blade tip clearance system as claimed in claim 5 and including ducting connecting a compressor of said gas turbine engine in flow series with the inner surface of said inner casing so as to cause said inner casing to flex outwardly of the axis of said gas turbine engine and thereby lift said segments away from said blade tips during an appropriate stage in the engine operating regime.

7. A gas turbine engine as claimed in claim 6 including ducting connecting a compressor of said gas turbine engine in flow series with said outer casing so as to heat said outer casing and cause it to expand and thereby slow the rate of return of said segments to their original positions, during an appropriate stage in the engine operating regime.