GB2226366A - Gas turbine engine coolant temperature sensing - Google Patents

Abstract

The turbine rotor discs 17 to 21 are cooled by airflows 85 .... which flow through various restricting apertures and passages 82, 92, 100 .... and various air seals 70, 76, 118 .... acting between rotating and static components of the engine, to control admission and exhaustion of cooling air to and from chambers 78, 86, 106, 110 .... defined between the rotor discs or between the rotor discs and adjacent static structure 58, 152. A number of angularly-spaced-apart probes 300 project into an inter-turbine chamber 150 defined between the second stage high pressure turbine disc 18 and adjacent static structure 152 and sense the temperature of the air in the chamber. The thermocouple probes 300 are linked to a signal processor (316, Fig. 2) which gives a warning should a safe temperature limit be exceeded. A further set of probes may monitor the temperature in the chamber 247 inward of stator blades 29. The probes may be thermocouples or fibre-optic thermistors. <IMAGE>

Description

GAS TURBINE ENGINE The present invention relates to gas turbine engines having one or more turbine rotor stages of the bladed disc type, in which the or each disc is actively cooled by means of a cooling air system, and a means is provided for detecting hazardous failure of air seals in the cooling air system.

Modern gas turbine aeroengines have internal air systems whose airflows do not directly contribute to the engine's thrust. Such systems have several important functions to perform for the safe and efficient operation of the engine, one of these functions being to ensure that turbine rotor discs do not overheat. Excessive overheating of the turbine rotor discs could imperil the safety of the aircraft by weakening the rotor disc material so that it ruptures binder the high centrifugal loadings experienced by the rotors.

To prevent disc overheating, cooling air is caused to enter the annular chambers formed between successive rotor disc stages, and/or between a rotor disc stage and adjacent static structure. Flow of cooling air over the discs and between chambers is controlled by throttling apertures and by various air seals between static and rotating structure of the engine and, on completion of the cooling function, the air is expelled into the main turbine gas stream.

A problem arises in such cooling air systems in that excessive wear or damage to the air seals may be deleterious to the proper flow of air round the system and may lead to overheating, particularly of the rotor discs of the high pressure turbine. Due to the high stresses experienced by such rotor discs, there is a resulting possibility of hazardous failure of the deficiently cooled discs.

It is an object of the present invention to help reduce the possibility of a hazardous failure of this sort.

According to the present invention, a gas turbine engine turbine arrangement comprises: at least one turbine rotor stage comprising a rotor disc whereon are mounted turbine blades for rotation in a turbine annulus passage; static structure adjacent the disc, the disc and the static structure defining a chamber therebetween; passage means for admitting and exhausting cooling air to and from the chamber;; a plurality of air seals arranged to act between rotating and static components of the engine and having the effect of controlling said admission and exhaustion of cooling air, some of said exhaustion of cooling air from the chamber being a flow of cooling air radially outwards past the disc into the turbine annulus passage, thereby to take heat from the disc, whereby in the event of a substantial reduction in the flow of cooling air past the disc, the disc tends to overheat, and there is a corresponding rise in the temperature of the air in the chamber; a plurality of temperature transducers secured to the static structure and angularly spaced apart around the interior of the chamber, the transducers being adapted to produce signals having a value related to the temperature of the air in the chamber; and signal processing means for receiving the signals from the transducers, detecting if at least one transducer is registering a temperature above a predetermined limit and outputting an overtemperature signal if such detection occurs.

In the exemplary embodiment, the plurality of air seals includes an air seal acting between the disc and the adjacent static structure to control the flow of cooling air therepast radially outwards past the disc, the flow of cooling air past said air seal being reduced in the event of a substantial increase in air flow past others of the plurality of air seals.

The turbine arrangement may comprise at least a high pressure turbine and a low pressure turbine together constituting a plurality of turbine rotor stages, the static structure being situated between the high and low pressure turbines. Preferably, the temperature transducers are fixed to the ends of respective probe means which pass through stator means situated in the turbine annulus passage between the high and low pressure turbines and project into the chamber through the static structure, the probe means being sealingly engaged therewith.

A further plurality of temperature transducers may be similarly arranged around a further chamber supplied with air exhausted from the chamber defined between the disc and the static structure, the amount of air supplied to the further chamber being reduced by the occurrence of a substantial increase of air flow past the air seal means acting between the disc and the adjacent static structure, whereby the temperature of the air in said further chamber increases, the further plurality of transducers being adapted to produce signals having a value related to the temperature of the air in the further chamber and the signal processing means being adapted to receive the signals from the further plurality of transducers, detect if at least one transducer is registering a temperature above a second predetermined limit and output a second overtemperature signal if such detection occurs.

In an exemplary embodiment of the invention, the signal processing means comprises a highest wins amplifier means to which each of the first plurality of transducers are connected, and comparator means for comparing the output of the amplifier means with a predetermined limit and outputting the overtemperature signal if the limit is exceeded. Likewise if the further plurality of temperature transducers are utilised as above, the signal processing means further comprises a further highest wins amplifier means to which each of the further plurality of transducers are connected, and further comparator means for comparing the output of the further amplifier means with a further predetermined limit and outputting a further overtemperature signal if the limit is exceeded.

Alternatively, the signal processing means may comprise a multiplexer to which the signals from each transducer are passed, an analogue-to-digital converter for sampling each input to the multiplexer in sequence, and a microprocessor with memory means for comparing each successive digitised signal value with an appropriate predetermined limit and outputting an overtemperature signal as previously mentioned.

Preferably, the temperature transducers comprise thermocouples.

An exemplary embodiment of the invention will now be described with reference to the accompanying drawings, in which: Figure 1 is an axially sectioned side elevation of part of the turbine arrangement in a gas turbine engine; and Figure 2 is a diagrammatic front elevation of the arrangement of the electrical components of the invention.

Referring now to Figure 1, an axial-flow gas turbine 10 forms part of a gas turbine engine and comprises five turbine rotor stages 12 to 16, the first two being designated the high pressure turbine and the latter three being designated the low pressure turbine. The turbine rotor stages 12 to 16 comprise a rotor disc 17 to 21 respectively and a large number of aerofoil-shaped turbine rotor blades 22-26 respectively. Associated with each turbine rotor stage 12 to 16 is a preceding circumferentially extending row of stator blades 27 to 31 respectively, the first row 27 being a set of nozzle guide vanes situated just downstream of the annular nozzle 40 of an annular combustion chamber 42. As is well known, the main aerodynamic function of the stator blades 27 to 31 is to direct the combustion gases onto the turbine blades 22-26 at an efficient angle to enable the turbine blades to extract power from the gases.

As is conventional, the high pressure turbine, comprising stages 12 and 13, drives a high pressure compressor (not shown) through an outer drive shaft 44, while the low pressure turbine, comprising stages 14 to 16, drives a low pressure compressor and a large diameter fan (not shown) through an inner drive shaft 46, the two shafts 44 and 46 being of course concentric. It will be seen from Figure 1 that power from rotor disc 18 is transmitted to rotor disc 17 primarily by bolted annular flanged coupling 48, and power from both high pressure turbine discs is transmitted to outer drive shaft 44 primarily by bolted annular flanged coupling 50 joining disc 17 to the shaft.

In similar fashion, rotor discs 19 and 21 are joined to rotor disc 20 by respective couplings 52,54 and power from the low pressure turbine is transmitted to inner drive shaft 46 primarily by coupling 56 joining disc 20 to the shaft.

Various other features of the detailed design of the gas turbine are shown on the drawing, but they will not be described unless they are connected with explanation of the present invention, since such description is not essential to understanding of the invention.

Looking now at the air cooling system associated with the invention, it will be seen that the nozzle guide vane row 27 is supported by annular static structure 58 which also supports two annular flanges 60,62 having lands which comprise an abradeable layer 64,65. Ribs are formed on an annular flange 68 which comprises part of rotor disc 17 and the ribs run against the abradeable layer 64 on flange 60 to form an air seal 70 between the disc 17 and static structure 58. Ribs are also formed on a further annular flange 72 which forms part of a member 74 bolted to the outer drive shaft 44 at coupling 50, and the ribs run against the abradeable layer 66 on flange 62 to form an air seal 76 between static structure 58 and outer drive shaft 44.Hence there is formed, adjacent the forward side of rotor disc 17, an annular chamber 78 into which cooling air 80 flows through apertures 82 in a panel 84 forming part of static structure 58.

Some cooling air 85 escapes from chamber 78 through seal 70 into a further annular chamber 86 situated radially outwardly of chamber 78. Chamber 86, like chamber 78, is defined between static structure 58 and rotor disc 17. Air also enters chamber 86 through an annular array of pre-swirl nozzles 88 whose function is to direct cooling air into cavities 90 at the bases of high pressure turbine blades 22 for supply to the interiors of the blades.

Excess air, not needed for cooling the blades, escapes from chamber 86 into the main turbine gas flow through a small gap 92 between blades 22 and nozzle guide vanes 27. In this way there is established a net flow of air 85 radially outwards through chamber 86 which cools the radially outer part of the forward face of the rotor disc 17.

Further cooling air 94 escapes from chamber 78 through seal 76 and is used to cool the walls 96 of the combustion chamber 42. However, much of the air 80 which enters the chamber 78 leaves it as air 98, by means of throttling passages 100 in coupling 50. Passages 100 are defined between annular flanges 102,104, formed integrally with disc 17 and shaft 44 respectively, and comprise radial grooves formed between circumferentially spaced radial lands on the forward face of flange 102.

Air 98 helps to cool the radially inner hub portion of the disc 17 as it flows towards and through passages 100, and as it passes between the inner bore of disc 17 and the outer surface of shaft 44 into an annular chamber 106 between the hub portions of adjacent discs 17 and 18.

From radially inner chamber 106, cooling air 108 is supplied to radially outer chamber 110 through apertures 112 in flange 114 which extends axially between disc 17 and mates with a complementary flange feature 116 on disc 18.

Air 108 exits from chamber 110 only through a circumferential row of holes 120 in flange 122, which extends axially from the rim of disc 17 to form part of coupling 48. Circumferentially extending ribs are formed on the radially outer surfaces of coupling flange 122 and on the corresponding coupling flange 124 extending from disc 18, and these run against an abradeable land formed on seal ring 126, thereby constituting a seal 118 which controls flow of air 108 from chamber 110 to chambers 128,130 on either side of stator support rings 132 to which seal ring 126 is bolted. From thence the air passes to the main turbine passage through gaps 134,136 between rotor blade 22, stator blade 28 and rotor blade 23. Once again, the flow of air 108 through chambers 106,110,128 and 130 takes heat away from the discs 17,18.

Returning to radially inner chamber 106, much of the air 98 which enters the chamber leaves it as air 138 by means of circumferentially spaced throttling passages 140 in a sleeve member 142 which is fixed to and surrounds shaft 44 and extends through the centre bore of rotor disc 18 to help support and locate the same. Air 138 then passes around the hub of rotor disc 18 until it encounters throttling apertures 144 which are angularly spaced around a support panel 146 whose radially outer circumference is bolted to a flange 148 extending from the hub of disc 18, and whose inner circumference is splined to the rear end of shaft 44, this arrangement again serving to support and locate the disc 18. Air 138 flows through apertures 144 into a chamber 150 adjacent the rear face of disc 18.

Adjacent the disc 18, but spaced rearwardly from its so as to define the chamber 150, is a static diaphragm 152 which supports seal carrier rings 154 to 157. Seal carrier ring 154 carries an abradeable pad 160 against which run ribs provided on a flange 162 extending axially from the disc support panel 146, thereby defining an air seal 164 for controlling entry of air into seal chamber 166. Seal carrier ring 155 supports seal carrier ring 156 which has an abradeable pad 168 against which run ribs provided on a seal ring 170 splined to shaft 46, thereby defining an air seal 172. Pad 168 and seal ring 170 are provided with respective circumferentially spaced apertures 174,176 situated midway of their axial extent, whereby air 178 from seal chamber 166 can pass into the interior of shaft 46 through holes 180 in the wall of the shaft.Air 178 passes through an annular space 182 defined between the wall of the shaft 46 and the external surface of a tube 184 which is concentric with the shaft and is sealed at 186 and 188 to the internal surface of the shaft. Air 178 helps to cool the shaft 46 as it passes axially through space 182 to circumferentially spaced holes 190 at the rear end of the shaft. It then passes through similar holes 192 in a sleeve 194, which helps to support and locate rotor disc 21 and is fixed to the rear end of shaft 46; it then enters a chamber 196 defined between the rear face of rotor disc 21 and static diaphragm 198.

The inner end of diaphragm 198 supports a seal carrier ring 200 and an abradeable pad 202 carried thereby cooperates with ribs on sleeve 194 to constitute seal 204.

Some of the air 178 which enters chamber 196 leaks through seal 204 into a lower pressure area (not shown), but the remainder, constituted by air 206, flows outwards past the rear face of disc 21, taking heat therefrom, and exits into the main turbine passage through gap 208 between a platform 210 of turbine blade 26 and a neighbouring stator blade platform 212.

Returning to the region of seal chamber 166, seal carrier 156 further carries abradeable pad 214, against which run sealing ribs on a flange 216 extending from support panel 146 so as to constitute a seal 218. Some air 220 passes through this seal 218 and into a lower pressure region between shafts 44 and 46, from whence it passes for utilisation in combustion chamber cooling. Besides the air lost from chamber 150 through seal 164, more air 222 is lost through seal 224 constituted by the sealing ribs on a flange 226 extending rearwardly from the rim of rotor disc 18 and by an abradeable pad 228 on static support ring 230, which supports the radially inner ends of stator blades 29.

After passing through seal 224 air 222 passes into the main turbine passage through gap 232 between rotor blades 23 and stator blades 29. However, some of the air 138 entering chamber 150 passes as air 233 through apertures 234 in static diaphragm 152, so entering a chamber 236 defined between the diaphragm 152 and the inner portion of the forward face of rotor disc 19. As mentioned previously, diaphragm 152 supports seal carrier ring 157, so as to define a seal 238 comprising abradeable pad 240 on carrier ring 157 and sealing ribs formed on a flange 242 extending from the forward face of rotor disc 19. This seal 238 divides inner chamber 236 from an outer chamber 244 defined between the radially outer parts of diaphragm 152 and the rotor disc 19.Chamber 244 is supplied with air 246 which leaks through seal 238, cools the radially outer part of the rotor disc 19, and enters a chamber 247 just inboard of the stator 29 through a gap 248 between the platform of rotor blade 24 and an adjacent static cone member 250 bolted to diaphragm 152 and support ring 230. Chamber 247 then exhausts to the turbine annulus passage through the gap 251 between adjacent stator and rotor blade platforms.

Most of the air 233 entering chamber 236 leaves it as air 250 through the gap between the hub of rotor disc 19 and the edges of the two seal carriers 155,156. There is some leakage of this air 250 past shaft seal 172 into the interior of shaft 46 due to a pressure differential, but most of it passes as air 251 through apertures 252 in sleeve member 254 which is clamped against a step on shaft 46. After passing through the annular space 256 between shaft 46 and sleeve 254, air 251 enters a chamber 258.

Some of the air 251 is exhausted into the main turbine passage as air 260,262 after passing through seal 264 associated with coupling 52 between rotor discs 19 and 20, but most of the air 251 enters a chamber 266 as air 268 which passes through the bore of rotor disc 20 and through passages 270 in coupling 56. All the air 268 entering chamber 266 exits into the main turbine passage as air 272,274 after passing through seal 276 associated with coupling 54 between rotor discs 20 and 21, there being no other exits from chamber 266. Seals 264 and 276 and couplings 52,54 between rotor discs 19 and 20, 20 and 21 are similar to seal 118 and coupling 48 between rotor discs 17 and 18.

It will be seen from the above that the flow of cooling air over the rotor discs 17 and 21 and between the chambers 78,85,106,110,150,166,196,236,244,258 and 266 is controlled by means of the various seals 70,76,118,160,172, 202,218,224,238,264 and 276 between static and rotating structure of the engine and by means of the various apertures such as 100,140 and 144 which act as flow restrictors. In particular, the flow 138 into chamber 150 is, under normal circumstances, controlled primarily by throttling apertures or passages 100,140 and 144.

It is exceedingly unlikely that any of the above-mentioned air seals would be damaged or subjected to excessive wear singly - rather, the likely cause of failure would be a degree of eccentric rotation of one or both of the low or high pressure turbines and their shafts, causing damage or excessive wear to the respective associated groups of seals 172,204,238,264 and 276 or 70,76,118,164, 218 and 284.Thus, for instance, if some or all of the seals 164,172,204,238,264 and 276 should fail or be damaged or be subject to excessive wear so as to let through large quantities of air into lower pressure areas, the air in chamber 150 at the rear of the second stage high pressure turbine disc 18 would tend to be reduced in pressure and the flow of air 222 through seal 224 would be reduced or would cease altogether, leading to overheating of the rim of the disc 18 outboard of seal 224 due to ingress of hot turbine gases and probably also of the main body of the disc 18 inboard of the seal 224, because the flow of air 222 would no longer be adequately cooling the rear face of the rotor disc and preventing the hot gases from penetrating gap 232.In addition, lower pressure in chamber 150 could also impair the throttling function of apertures 140,144 and lead to lower pressures in chambers 106,110 which again would reduce air flow 108, with consequent reduction in cooling of the radially outer parts of rotor discs 18 and 17.

It will also be realised that failure, damage, or excessive wear of seals 70,76 and 118 could also prevent sufficient cooling air from reaching chambers 150,196,236,244,258 and 266, again possibly leading to overheating of rotor discs 18 to 21 and other components in these areas.

In order to certify a civil aeroengine under current Aviation Authority rules, the manufacturer must show that if the probability of a failure in the turbine cooling air seals is greater than 1 in 109 hours, such a failure will not result in loss of coolant flow to any part sufficient to cause a hazardous failure, such as bursting of a turbine disc. If the manufacturer cannot show this, then the rules stipulate that some sort of detector must be fitted to the engine to indicate the occurrence of a problem and thus allow a safe shut-down of the engine.

Therefore, if the above-mentioned reliability figure cannot be achieved, either a detector must be fitted, or there must be a large margin of safety either in the amount of cooling air flowing through the turbine cooling system while the seals are unfailed, or in the operational life of the rotor discs as the most critical components. Both the latter two alternatives are undesireable because they involve cost penalties, the first involving reduced fuel economy and the second involving increased maintenance costs and expenditure on replacement components.

However, further difficulties arise if it is desired to utilise a detector, since such a detector must: (i) have a rapid positive response to the onset of hazardous conditions in the turbine cooling system; (ii) be sufficiently reliable, such that the probability of having the detector failed at the time when a cooling system failure occurs is 1 in 109 hours or less; (iii) be accessible for maintenance, checking or replacement without the need for disassembly of the turbine.

These difficulties are obviated in the embodiment shown in Figures 1 and 2 in that a total of nine angularly spaced thermocouples 300 are provided within chamber 150 to sense the temperature of the air therein, the thermocouples 300 being at the end of long hollow probe rods 302 which extend from housings 303 through holes 304 in circular bosses 306 on the turbine casing 308, through the hollow interiors of large stator blades 29, and through glands 310 in static support ring 230, the probe rods being cranked at 312 to accommodate the need to fit through glands 310.

The thermocouples 300 generate signals having a voltage value which is dependent upon the difference between the temperatures of their hot and cold junctions.

Their hot junctions are at the positions 300 but their cold junctions are in a junction box 311 forming the 9-channel input to an amplifier 312 incorporating "highest wins" circuitry, such circuitry being well known to specialists in electronics. Each hot junction is connected to a corresponding cold junction through the usual twin thermocouple wires 313,314 which are routed up through the probe rods 302, around the engine casings, and to the junction box and amplifier 311,312. The latter two components are located outside the engine casing in the nacelle (not shown), where a stream of air at near ambient temperature keeps them cool, with all the junctions maintained at substantially the same temperature as each other in order to provide a common reference temperature for the thermocouples.Amplifier 312 receives the 9-channel analogue input and selects the signal having the highest voltage value, amplifies it, and outputs it as signal 315 to analogue comparator module 316 comprising, for instance, diode trigger circuitry as known, the circuitry being set to output an overtemperature signal 317 if signal 315 exceeds a predetermined voltage level, indicating an overtemperature in the chamber 150.

Overtemperature signal 317 is preferably digitised in an analogue-to-digital converter (not shown) for onward transmission to aircraft systems. Temperature-related signal 315 may also be taken and digitised for combination with digitised overtemperature signal 317.

Because the temperature of the junction box 311, and hence the temperature of the cold junctions, is dependent upon the variable temperature of the air stream flowing past it, it is necessary to sense the cold junction temperature in absolute terms and apply an appropriate correction in amplifier 312 to the selected amplified voltage signal 315. This is easily achieved by utilising a resistance thermometer (not shown) to sense the temperature of the interior of junction box 311, the variable potential drop across the thermometer being used to apply a suitable bias to the amplifier output, the whole system being, of course, calibrated before delivery to the customer.

It should be noted in connection with the above signal processing that even if several or all but one of the thermocouples fail, any overtemperature in chamber 150 will still be detected. Furthermore, if system reliability criteria demand it, the comparator module 316 can be duplicated or triplicated and failure of earlier parts of the system detected by the absence or an unduly low level of the voltage signal 315, causing triggering of a system failure alarm.

For reliability, the digital data processing components, being simple, may be duplicated or triplicated.

The combined digitised overtemperature signal is preferably a multi-bit digital word, comprising an alarm flag bit for automatic actuation of a warning alarm in the flight station, and, say, seven other bits representing the actual temperature being recorded in chamber 150 for visual display in the flight station. Shutdown of the engine is then a matter for the judgement of the pilot in the light of all the circumstances.

The thermocouples 300 are of assured reliability, being similar to existing known thermocouples 324 (Figure 1 only) already used for measuring turbine gas temperature and in the illustrated turbine being situated within the hollow stator blades 29, the thermocouples 300 therefore being mounted on extensions 302B of existing probe rods 302A for easy and inexpensive modification of existing engines.

The proposed detector arrangement will, if properly implemented, give rapid positive response to a cooling air system seal failure because chamber 150 is situated between the high pressure and low pressure turbines, and is that part of the cooling air system which receives all the cooling air excess to the requirements of components supplied by chambers 78/86 and 106/110, and which supplies all the cooling air to subsequent chambers and to the interiors of shafts 44,46; its pressure will therefore fall and its temperature will rise if there occurs damage, or excessive wear of air seals controlling exit of air from those places to areas of significantly lower pressure.

Furthermore, the rim of the second rotor disc 18 of the high pressure turbine, rather than any of the other rotor discs, is the one which will most easily be damaged due to excessive heat build-up if its necessary flow of cooling air 222 is reduced. This is because the rim of rotor disc 17 is flood-cooled by means of the large amount of air entering chamber 86 through nozzles 88, and because other discs 19 to 21 operate at lower temperatures and experience smaller centrifugal stresses due to rotation at lower speed.

Of course, the detection system as so far described would not detect overheating of discs 19 to 21 caused by failure of seal 224 only and consequent large increase in flow of cooling air 222, leading to lowered air pressure in chambers 196,244,258 and 266. However, this is not a serious drawback because the failure of a single seal on its own is unlikely and because of the less highly stressed nature of the discs 19 to 21 as previously discussed; furthermore, if necessary, such a failure would be detectable by a simple modification to the invention involving the addition of a further set of thermocouples 326 monitoring the temperature of the chamber 247 inboard of the stator blade 29, which would rise due to a reduction in cooling air flow 246.The thermocouples 326 would be connected back through similar signal processing circuitry to that shown and described in relation to Figure 2.

Although the only temperature transducers mentioned above are thermocouples, other sorts of temperature transducer could also be used, such as thermistors of temperature sensitive fibre-optics, appropriate changes in the signal processing being necessary.

Totally digital signal processing is also achievable in connection with the present invention, the junction box 311 being utilised as the input to a multiplexer through which the voltage value of the signal from each thermocouple is sampled in turn by an ADC. After digitisation, a microprocessor with appropriate ROM and RAM performs a simple algorithm which detects null returns from failed thermocouples and stores appropriate engineering maintenance actions for retrieval, adjusts the signal values with respect to absolute temperature, and compares the adjusted signal values to an appropriate limit, an overtemperature signal being generated if the limit is exceeded.

Claims (7)

Claims:

1. A gas turbine engine arrangement comprising: at least one turbine rotor stage comprising a rotor disc whereon are mounted turbine blades for rotation in a turbine annulus passage; static structure adjacent the disc, the disc and the static structure defining a chamber therebetween; passage means for admitting and exhausting cooling air to and from the chamber;; a plurality of air seals arranged to act between rotating and static components of the engine and having the effect of controlling said admission and exhaustion of cooling air, some of said exhaustion of cooling air from the chamber being a flow of cooling air radially outwards past the disc into the turbine annulus passage, thereby to take heat from the disc, whereby in the event of a substantial reduction in the flow of cooling air past the disc, the disc tends to overheat and there is a corresponding rise in the temperature of the air in the chamber; a plurality of temperature transducers secured to the static structure and angularly spaced apart around the interior of the chamber, the transducers being adapted to produce signals having a value related to the temperature of the air in the chamber; and signal processing means for receiving the signals from the transducers, detecting if at least one transducer is registering a temperature above a predetermined limit and outputting an overtemperature signal if such detection occurs.

2. A turbine arrangement according to claim 1 in which the plurality of air seals includes air seal means acting between the disc and the adjacent static structure to control the flow of cooling air therepast radially outwards past the disc, the flow of cooling air past said air seal means being reduced in the event of a substantial increase in air flow past others of the plurality of air seals.

3. A turbine arrangement according to claim 1 or claim 2 in which at least a high pressure turbine and a low pressure turbine together constitute a plurality of turbine rotor stages, the static structure being situated between the high and low pressure turbines.

4. A turbine arrangement according to any one of claims 1 to 3 in which the temperature transducers are fixed to the ends of respective probe rods which pass through stator means situated in the turbine annulus passage and project into the chamber through the static structure, the probe rods being sealingly engaged therewith.

5. A turbine arrangement according to any one of claims 1 to 4 in which a further plurality of temperature transducers also secured to static structure are angularly spaced apart around the interior of a further chamber supplied with air exhausted from the chamber defined between the disc and the static structure, the amount of air supplied to the further chamber being reduced by the occurrence of a substantial increase of air flow past the air seal means acting between the disc and the adjacent static structure, whereby the temperature of the air in said further chamber increases, the further plurality of transducers being adapted to produce signals having a value related to the temperature of the air in the further chamber and the signal processing means being adapted to receive the signals from the further plurality of transducers, detect if at least one transducer is registering a temperature above a second predetermined limit and output a second overtemperature signal if such detection occurs.

6. A turbine arrangement according to any one of claims 1 to 5 in which the temperature transducers comprise thermocouples.

7. A turbine arrangement substantially as described in this specification with reference to the accompanying drawings.