referring now to fig1 a turbofan engine 10 is shown to include a fan rotor 11 and a core engine rotor 12 . the fan rotor 11 includes a plurality of fan blades 13 and 14 mounted for rotation on a disc 16 . the fan rotor 11 also includes a low pressure or fan turbine 17 , which drives the fan disc 16 in a well - known manner . the core engine rotor 12 includes a compressor 18 and a power or high pressure turbine 19 which drives the compressor 18 . the core engine also includes a combustion system 21 , which combines a fuel with the air flow and ignites the mixture to inject thermal energy into the system . in operation , air enters the gas turbine engine 10 through an air inlet 22 provided by means of a suitable cowling or nacelle 23 which surrounds the fan rotor 11 . air entering the inlet 22 is compressed by means of the rotation of the fan blades 13 and 14 and thereafter is split between an annular passageway 24 defind by the nacelle 23 and an engine casing 26 , and a core engine passageway 27 having its external boundary defined by the engine casing 26 . the pressurized air which enters the core engine passageway 27 is further pressurized by means of the compressor 18 and is thereafter ignited along with high energy fuel in the combustion system 21 . this highly energized gas stream then flows through the high pressure turbine 19 to drive the compressor 18 and thereafter through the fan turbine 17 to drive the fan rotor disc 16 . the gas is then passed out the main nozzle 28 to provide propulsion forces to the engine in a manner well known in the art . additional propulsive force is gained by the exhaust - pressurized air from the annular passage 24 . it should be noted that although the present decription is limited to an aircraft gas turbine engine , the present invention may be applicable to any gas turbine engine power plant such as that used for marine and industrial applications . the description of the engine shown in fig1 is thus merely illustrative of the tupe of engine to which the present invention is applicable . referring now to fig2 the high pressure turbine portion of the engine is shown in greater detail and comprises a single - stage row of rotor blades or buckets 29 rotatably disposed in the flow path of the hot gases as shown by the arrows . the hot gasses flow from the annular combustor inner casing 31 rearwardly to a row of circumferentially spaced high pressure nozzles 32 , through the circumferentially spaced row of buckets 29 , through a circumferentially spaced stationary row of low pressure nozzles 33 to finally impinge on the circumferentially spaced row of rotatable low pressure turbine blades or buckets 34 of the fan turbine 17 and finally to exhaust out the main nozzle 28 . circumscribing the row of high pressure buckets 29 in close clearance relationship therewith is an annular shroud 36 made of a suitable abradable material for closely surrounding buckets 29 but allowing some frictional engagement and wear at particular operational moments wherein the clearance between the shroud and the blades may be temporarily lost . the shroud is preferably made of a number of annular sectors attached to the inner side of an annular band 37 by conventional means . the annular band 37 is preferably made up of a number of sectors forming a complete circle . the annular and 37 is in turn supported by an annular ring 38 having at its rearward end , a radially inwardly extending collar 39 which is attached to the annular band 37 by way of an annular bracket 41 . the forward side of the annular band 37 is attached to the annular ring 38 by way of an l - shaped annular bracket 42 and a plurality of bolts 43 . support for the annular ring 38 is derived by connection to the low pressure nozzle support 44 by bolts 45 at the rear end thereof , and connection to the turbine casing 46 along with the high pressure nozzle support 47 by way of a plurality of bolts 48 spaced circumferentially around the casing . as the turbine casing 46 extends rearwardly around the high pressure turbine portion of the engine , it is suddenly enlarged by the manifold portion 49 which forms an annular plenum 51 between the manifold and the annular ring 38 . fluidly communicating with the plenum 51 is a plurality of air bleed off conduits 50 which carry bleed off air from the intermediate stages of the compressor 18 for the purpose of turbine nozzle cooling in a manner well known in the art . referring now more specifically to the annular ring 38 , a radially outwardly extending flange 52 is formed the reon to project outwardly into the plenum 51 . axially spaced in the rearward direction an l - shaped flange 53 also projects radially outward but not to the extent of the outer diameter of the flange 52 . the ring 38 with its flanges 52 and 53 is composed of the material having a relatively low coefficient of thermal expansion . rigidly attached to flange 52 and projecting forwardly to the turbine casing 46 is a cylindrical structure 54 which surrounds a portion of the annular ring 38 to form a cavity 56 therebetween . a plurality of apertures 57 are formed around the circumference of the cylindrical structure 54 to provide fluid communication between the plenum 51 and the cavity 56 . fluid communication is further provided between the cavity 56 and that area 58 defined by the axially spaced flanges 52 and 53 , by a plurality of axially extending holes 59 formed in the flange 52 . to further define that area 58 between the two flanges 52 and 53 , a cylinder 61 is rigidly attached to the flange 52 by a plurality of bolts 62 , and extends axially rearwardly to surround the outer surface 63 of the flange 53 . the cylinder 61 is composed of a material having a high thermal coefficient of expansion which reacts with the ring 38 and associated flanges to control the flow of bleed off air in the plenum 51 during transient and steady - state periods of engine operation to obtain the desired state of growth characteristics for the establishment of proper clearances between the shroud 36 and the turbine buckets 29 . at the rearward end of the manifold 49 an annular support structure 64 is attached to the manifold 49 by a plurality of bolts 65 and acts to support the second stage low pressure nozzle 33 . the support structure 64 is connected to the stage one nozzle support 44 , and together they partially define a secondary plenum 66 which is downstream of and supplied by cooling air from the plenum 51 . a plurality of circumferentially spaced holes 67 provide fluid communication between the secondary plenum 66 and the nozzle cavities 68 for cooling of the nozzles in a manner well known in the art . further defining the secondary plenum 66 is the annular oblique flange 69 connected to the manifold 49 by bolts 65 extending radially inwardly to surround the l - shaped flange 53 at a radially outward spaced position so as to trap one end of the cylinder 61 therebetween . the annular flange 69 and its mechanically connected parts are composed of a material with a relatively low thermal coefficient of expansion . the interaction of the cylinder 61 with the adjacent surfaces of the flange outer surface 63 and the annular oblique flange seat surface 70 acts as a temperature responsive valve which is closed when the dotted line position and open when in the position shown in fig2 . in a typical operation of an aircraft turbine engine , assume that the aircraft engine is in the idle position . the air entering the plenum 51 is relatively cool since it hasn &# 39 ; t been compressed to any great degree , and the cylinder 61 is thus relatively contracted position as shown by the dotted lines in fig2 to present a closed valve position . the flow of air from the plenum 51 through the apertures 57 to the annular slot 59 is thus virtually shut off and the air flow pattern tends to be as shown by the dotted line arrows from the plenum 51 to the secondary plenum 66 . as the engine is accelerated , for example to maximum thrust position , the degree of compression of the air in the compressor 18 is increased , and the air being delivered to the plenum 51 is relatively hot . this hot air acts on the exposed cylinder 61 to cause it to quickly increase in size . conversely , the angular ring 38 and its associated flanges 52 and 53 are very slow to respond to this temperature change and when responding they do not expand to the degree that the cylinder 61 expands . the result is that the cylinder unrestricted end expands to the position shown by the solid lines in fig2 to open the valve . the air from the plenum 51 then passes through the apertures 57 , the holes 59 , into the space 58 and into the plenum 66 as shown by the solid arrows in fig2 . since the air flows past the flanges 52 and 53 and over the ring 38 , the hot gases tend to quickly heat up the combined structure and cause it to expand relatively fast to thereby increase the inner diameter of the supported shroud 36 . when the speed of the engine is subsequently reduced , as by a throttle chop , the air being delivered to the plenum 51 is again relatively cool and the cylinder 61 quickly responds to the thermal change to shrink back to the dotted position of fig2 and close the valve . the ring 38 and associated flanges 52 and 53 are thus virtually isolated with the hot gases and tend to cool very slowly to thereby bring the shroud 36 down to its shrunken condition at a very low rate . assuming now that the throttle is again moved to the maximum thrust position ( hot rotor burst ), the cylinder 61 is again exposed to hot gases and the valve is opened to again cause the support structure to expand relatively fast to increase the size of the shroud enclosing area . it will , of course , be understood that various other designs and configurations can be employed to achieve the objects of the present invention . for example , the thermal valve , which has been described in terms of the ring 38 , cylinder 61 and flange 69 , may comprise various other arrangements to bring about the regulation of the shroud support temperature . the &# 34 ; open &# 34 ; and &# 34 ; closed &# 34 ; positions of the valve may be interposed to route the thermal fluid in the desired direction and manner . the fluid may be derived from a location other than the compressor , and its temperature may bear a different relationship from that of being proportional to engine speed as described . further , the shroud support structure as shown and described is merely illustrative of various structures which could be thermally regulated with respect to size in order to facilitate the desired transient and steady - state radial positions of the shroud .