Abstract:
Use of an optical fiber for the direct receipt of heat radiation for transmission to a remote pyrometer is enabled by the provision of an apertured, contaminant free compartment in the component being heated, and aligning the heat receiving end of the optical fiber with the aperture so as to receive radiated heat from within the compartment.

Description:
FIELD OF THE INVENTION 
     The present invention relates to a sensor device that is designed for use in a hot environment, and a monitor with which to measure the temperature thereof during that use. 
     BACKGROUND OF THE INVENTION 
     The present invention has particular efficacy, but by no means restrictively so, when used in the gas turbine field. 
     It is known, to measure the temperature in a gas flow through a gas turbine engine turbine section. From this, the temperature of the turbine components over which the gas flows may be assessed. An example of known art is described and illustrated in published specification GB 2 248 296, wherein an optically transparent sapphire member has a thermally emissive, metal oxide layer facing its end extremity, which layer is exposed to a flow of hot gas. Heat radiated from the layer passes through the sapphire member and a fibre optic cable, to a standard pyrometer, which translates the temperature into a useable electronic signal. 
     All the prior art known to the applicant for a patent for the present invention, have at least one common factor, this being that that surface which radiates the heat to the pyrometer, is immersed in the high speed gas flow, and consequently the optical receiver suffers gradually reducing ability to pass heat radiation thereto. This is brought about by exposure of the optical surface to the products of combustion, including carbon particles. A further drawback that other known heat monitors will experience, is that engines now being designed and built, will operate at temperatures higher than any previously achieved, their turbine structure being composed of materials capable of operating in those higher temperatures. Such temperatures will destroy known sensors. 
     SUMMARY OF THE INVENTION 
     The present invention seeks to provide an improved combination of a heatable sensing member and a temperature monitor therefor. 
     Accordingly the present invention comprises, in combination, a component operable in a hot environment and including therewithin a compartment sealed against ingress of contaminates generated in the said hot environment, and a heat monitor comprising an optic fibre located outside that space volume wherein said hot environment will occur, in spaced relationship with said component and with one end extremity aligned with an aperture in said component via which during a said hot environment operation, heat conducted into the compartment via its wall is radiated to said optic fibre end extremity for transmission thereby to a remote pyrometer. 
    
    
     DESCRIPTION OF THE DRAWINGS 
     The invention will now be described, by way of example and with reference to the accompanying drawings, in which: 
     FIG. 1 is a diagrammatic part view of a gas turbine engine with part of the associated turbine section exposed so as to show the location of a component and heat monitor in accordance with one aspect of the present invention. 
     FIG. 2 is an enlarged, lengthwise cross sectional of the component, a turbine guide vane, showing said aperture. 
     FIG. 3 is a chordal cross sectional view on line  3 — 3  of FIG. 2 . 
     FIG. 4 is an enlarged, axial cross sectional view through the heat monitor of FIG.  1 . 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Referring to FIG. 1. A gas turbine engine  10  includes a turbine section  12 , through which hot gases from combustion equipment within casing  14  are expanded in known manner. The gas flow is contaminated with combustion products, carbon particles and atmospheric dust that, inevitably, is sucked into the engine  10  by the compressor within casing  16 . 
     During operation of gas turbine engine  10 , it is important to monitor the temperature of the gases passing through the turbine section  14 , so as to assess the operating conditions of the engine. To this end, the present invention measures the temperature of one or more of guide vanes  18 , which, being soaked by the gas temperature, exhibits a closely related level of temperature thereas. The measurement is achieved by providing a compartment  20  within each guide vane  18 , at a position near its leading edge  22  (FIG.  2 ), which compartment has an aperture  24  in its radially outer end with respect to the axis of rotation of engine  10 , and fixedly positioning an optical fibre radiation transmitter  26  between a pair of casings  28  and  30  (FIG. 4) that surround the stage of guide vanes  18 , so that the end extremity of the optical fibre is facing the aperture  24 . 
     Compartment  20  contains a thin bridge  32 , which spans the width of aperture  24  and is aligned with optical fibre  26 . Passages  34  extend through bridge  32 , from the pressure side  36  of the guide vane  18  to its suction side  38 , as can be seen in FIG.  3 . Hot gas can thus flow across vane  18 , heating bridge  32  as it does so. The dimensional proportions of bridge  32  are sufficiently small, as to ensure that bridge  32  will easily attain a temperature equal to that of the gas flowing through the passages  34  therein. 
     It will be appreciated, however, that it is not essential that the thin bridge  32  is present. It is only necessary that the optical fibre  26  is aligned with a portion of the compartment  20  that attains an appropriate temperature. 
     Referring now to FIG.  4 . The optical fibre radiation transmitter  26  is constructed from an optical fibre  40  which includes a thermally emissive coated member  41  and lens  43  in known manner, fitted within a body in the form of a jacket  42 . Optical fibre  40  and jacket  42  extend towards aperture  24  in guide vane  18 , but their ends stop short thereof so as to allow provision of a nozzle  46  which itself, is part of a further jacket  50  and extends to a position very close to aperture  24 , for reasons explained later in this specification. 
     Jacket  42  includes an annular compartment  48 , which is filled with a cooling fluid e.g. water, for the purpose of cooling the optical fibre  40  when on engine shut down, cooling airflow stops, but engine temperature temporarily rises. Jacket  50  surrounds jacket  42  and is retained thereon by a screw threaded connection  52 . Jackets  42  and  50 , between them, define a further annular compartment  54 . A small cooling air supply (not shown) is connectable, via a conduit  56 , to compartment  54 , which also serves the purpose of preventing the entry of combustion gases to the optical path by leakage. Thereafter, the cooling air exits the optical fibre radiation transmitter via nozzle  46 . This cooling function is augmented by heat extracting fins  57  formed on the upper end of jacket  42 , as viewed in FIG.  4 . 
     The lower end  58  of jacket  42  and that part of jacket  50  that overlaps it are identically tapered and the tapered portion of jacket  42  has swirl vanes  60  formed thereon, so as to impart a swirling motion to the cooling air as it flows towards the outlet of nozzle  46 . The swirling motion, combined with the curved shape of the interior wall surface  6  of nozzle  46 , causes the airflow to adhere to wall  62 , thus avoiding interference with heat radiating up the central portion of nozzle  46 , onto the end face  44  of optical fibre  40 , which, if it occurred, could degrade the radiation intensity, and send a false signal to a standard radiation pyrometer  64  (FIG. 1) located in a cool position remote from the engine  10 . 
     On reaching the outlet of nozzle  46 , the curved surface thereof allows the cooling air to escape in directions radially away from the nozzle axis, initially entraining any air leaving aperture  24  and thereafter, any contaminated air leaked from the gas path in which vane  20  resides. That air, contaminated or not, will thus be prevented from contacting the end face  44  of optical fibre  40 , and is dispersed in an annular space  66 , which is defined by the turbine casing  30 , and the outer platforms  70  of the guide vanes  18 , the inner surfaces  72  of which, along with other known cylindrical structures (not shown)  15  define the outer boundary of the gas annulus in known manner. 
     The skilled man, having read this specification, will realise that the present invention protects the optical path from exposure to hot high velocity gas comprising combustion products, carbon particles and airborne dirt. He will further appreciate that that the present invention obviates the need for the provision of a dedicated airflow across the face of the radiation receiver, and regular maintenance to clean it. Moreover, whilst the present invention is described and illustrated in connection with a gas turbine engine, the skilled man will appreciate that the radiation transmitter  26  can be used in any component wherein the provision of a suitable compartment corresponding to compartment  20  is possible, whether it be an operating powerplant, or a manufacturing process for e.g. a heat treatment process, or a metal melting or alloy forming process.