US2600235A - Gas turbine rotor cooling means - Google Patents

Description

June 10, 1952 J. A. N. GALLIOT GAS TURBINE RoToR COOLING MEANS Filed Feb. 24, 1947 5 Sheets-Sheet June 10, 1952 J. A. N. GALLIOT 2,600,235

GAS TURBINE ROTOR COOLING MEANS Filed Feb. 24, 1947 3 Sheets-Sheet 2 Inventor n /6 33 Ju LeslAnJrNorLert Gamm; BY @j m Attorneys June l0, 1952 J. A. N. GALLloT 2,600,235

GAS TURBINE ROTOR COOLING MEANS Filed Feb. 24, 1947 3 Sheets-Sheet 5 ATTORNEYS Patented June 10, 1952 UNI TED STAT E S PATE NT OFF ICE GAS TRBNE ROTOR COOLING S JlilesAndr Norbert Galliot,'London,-England `Application February 24, 1947, Serial No. 730,446 In Great Britain February 25, 1946 (Cl. (S-35.6)

13Claims. 1

This invention relates to lgas turbines such as are vemployed for *aircraft propulsion and like purposes, and comprising for example aperipherally bladed Vrotor `secured 'upon a shaft which drives a compressor feedingair to one or lmore combustion `chambers furnishing 7hot gases to drive the peripheral blades Aof the rotor.

The main object of kthe 'present invention is to provide animproved gas turbine` so arranged and constructed that its rotor affords passage for cold air flowing tothe compressor intake whereby the rotor is effectively cooled, ltogether with its blades, shaft and bearings.

A secondary vobject of the invention is to provide an vimproved gas turbine Yhaving a rotor fitted With vexternal blades driven by -hot gases and "internalvanes forcing cold air axially to a compressor feeding a combustion chamber which furnishes the hot gases, said blades and vanes being `:ln thermally conductive relation rto trans- `mit heat from said blades to said vanes vfor dissipation inthe `cold air passing ybetween said vanes.

A further object of vthe invention is to -provide an improved gas turbine having anV apertured rotor which constitutes a primary element of an auxiliary compressor forcing the air `admitted through 'the rotor apertures Yto the lintake of rthe main compressordriven by the turbineshaft.

Another object of the invention is to prov-ide an improved gas turbine 'having' an apertured rotor 'and an air compressor mounted towards 'opposite ends of a Adriving shaft, with a tubular casing around the shaft, the airadmitt'ed through the rotor apertures flowing `along the interior of 'the casing to the intake of the-compressor.

A still further `object of the invention is to provide an improved gas turbine having an apertured and externally bladed rotor and a main compressor mounted towards opposite ends of a driving shaft, with a tubular casing around :the shaft-an axial-typecompressor inside the casing, and an Aannular series of combustion chambers spaced around the casing, the air' Vadmitted through the rotor apertures being forced along rthe interior-of the casing `.through the axial-type compressor to the intake ofthe main compressor which delivers `compressed air v'to the combustion chambers, and the latter furnishing hot gases to drive the external blades of the rotor.

A specific object ofthe invention is `to lprovide an improved gas turbine for jet-propelled aircraft, with or Vwithout an airscrew driven by the turbine shaft, in which an apertured turbine rotor and its air compressor 'are mounted towards oppositeends of the turbine shaft, with anannular seriesof combustion chambers-spaced around the shaft, the air admitted through the rotor apertures flowing to the intake of the compressor which 'delivers compressed air .to wthe combustion chambers, the latter furnishing hot gases to drive external blades upon ythe turbine rotor, and the gases `exhausted from the turbine `being collected to form a propulsive jet.

Other objects and advantages of the invention will hereinafter appear from the following description, given `with reference tothe accompanying drawings,in which:

Fig. 1 is a half sectional elevation of the turbine;

Fig. 2 isa partial development of the rotor on a larger scale, showing six of the turbine blades;

Fig.3 is a partial end-view of the rotor, showing six of the external turbine blades as well as the corresponding internal vanes forming the primaryelement of an axial-type compressor;

Fig. 4 is a detail on a larger scale, showing a modified form of platform;

Fig. 5 is a partial end-view of the rotor hub mounted upon the turbine shaft, showing the method of ,assembling the roots ofthe compressor vanes;

Fig. 6 is a `half-section of the rotor hub, seen at right angles to Fig. 5 takenat the center line;

Figs. 7 and y8 are part-sectional views, showing an exhaust manifold for feeding a propulsiva-jet for aircraft;

Fig. 9 is a half sectional elevation of a modified turbine arranged to drive an airscrew for aircraft propulsion and also to deliver a jet for the same purpose; and

Fig. 10 is a half sectional elevation of another modiilcation'of turbine arranged for pure jet propulsion of an aircraft.

In the construction illustrated in Fig. l, the turbine shaft I0, driven by a rotor I I at the forward end, drives an air compressor I2 mounted towards theother end ofthe shaft; this compressor feeds a number of combustion chambers I3 spaced around the shaft, each chamber being provided with fuel-injection means I3a and delivering the hot'gases of combustion through guide vanes I4 to the external blades I5 of the turbine rotor.

The central portion of `the rotor II is apertured by providing it with spokes or webs which are shaped as helicoidal vanes I6 (see Figs. 2, 3, 5 and 6) to constitute a prima-ry element of an axial-type compressor; the cold air may bev admitted directly to the turbine rotor II by a flared inlet, opening in the direction of forward flight.

A shaft bearing, which is preferably of the selfaligning thrust-resisting type, is carried by radial arms or webs I8 inside the turbine stator, these arms being preferably formed as guide vanes to co-operate with the helicoidal vanes I6 of the rotor; in order to protect the rotor against the entrance of flying bodies, a wire net I9a may be supported by brackets or arms I3 connected at their outer ends to the exhaust manifold and at the inner end to a hub enclosing a bearing I1 for the turbine shaft.

The air compressor I2 may be of any suitable type, for example centrifugal; it is preceded by an axial-type compressor with multiple rotors 2G and stators 2I inside an induction tube or shaft casing 22, to which the cold air is fed by the primary element consisting of the internal rotor varies I6 already mentioned. The tube or casing 22, of diameter substantially equal at one end to that of the central portion of the rotor and at the other end to that of the compressor l2, surrounds the axial-type compressorl 20, 2I upon the shaft III connecting the turbine rotor to the air compressor at the rear, this tube or casing being preferably made of or lined with thermally insulating material, as indicated at 24, to prevent undue heating of the incoming air by the combustion chambers I3 outside the tube or casing.

The centrifugal compressor I2 may be provided with two non-symmetrical working faces, the front one operating as main compressor in tandem with the axial compressor which delivers to it, and the rear face operating in parallel as a separate source of compressed air, taking in air through suitable openings 213 at the rear end of the compressor casing; air may also be drawn through holes 21 in the shaft I0, adjacent to this rear face 25, the shaft being made hollow to admit air at one or both ends. The separately acting rear compressor 25 will of course be suitably dimensioned to give the desired capacity; the two deliveries of compressed air from the front and rear faces may be merged together at the periphery of the double-faced compressor I2, 25 before admission to the combustion chambers I3. As illustrated in lFig. l, the compressor I2, 25 delivers into a manifold 28 having separate branches 28a leading to the respective combustion chambers, suitable baffles or deflectors 23 being provided inside the manifold to guide the compressed air into the breech ends of tie chambers.

The one or more compressors may feed an nular combustion chamber, instead of the several separate combustion chambers I3 generally utilized, the chamber or chambers in the respective cases furnishing hot gases to drive the turbine rotor II at the forward end 0f the shaft I0.

The rear bearing 30 of the turbine shaft may be mounted in a plate 3| supported by brackets 32 from the body or casing of the compressor I2, beyond which the shaft extends to operate auxiliary equipment (not shown) such as mechanism for feeding the fuel to the combustion chamber or chambers, or any other devices that may be desired for the operation of an aircraft or for other purposes.

Since the primary compressor vanes I6 are effectively cooled by the incoming air, the central turbine blades I5 to the cold vanes IB, their junctions are preferably in metallic continuity, for

example, they may be formed by segmental plates or platforms 33 integral with the blades and vanes. These platforms are assembled to constitute a complete annulus, of which a portion is shown in Figs. 2 and 3, the opposite edges of the platforms being engaged by clamping rings 34 of channel or V-section, held together by bolts 35 threaded into screwed nipples 36 of similar design to those employed on cycle wheel spokes. As seen in Fig. 1, the annulus formed by the assembled plates or platforms 33, with the clamping rings 34, fits into grooves 41 on the adjacent edges of the stator walls; further, on the side directed towards the exhaust, a thin lip 48 projects from the roots of the turbine blades I5 so as to overlap the grooved edge of the stator wall on this side. Due to the high velocity of the gases acting upon the turbine blades, leakage of the gases into the interior of the rotor, which would pollute and heat up the incoming air, is prevented by the frictionless glands provided by these means.

Each of the segmental plates or platforms 33 may be made integral with one turbine blade I5 and the corresponding vane I6, in order to secure the greatest possible conduction of heat between those parts; due to the contrary curvature of the blade and vane, as seen in Fig. 2, it may be neces sary to give the platforms a rhomboid or nonrectangular shape, as shown in that figure.

Fig. 4 illustrates a modified form of platform 33 made integral with the respective turbine blade I5 and rotor vane I6, in which the oblique sides of the rhomboidal platform are castellated or provided with alternate projections and notches to interengage with the sides of the adjacent platforms; the ends of the platform may be made with square corners, as shown, and notched for lightness.

The inner ends or roots of the rotor spokes or vanes are mounted upon the shaft I0 by forming these roots 31 to a wedge-shape, as shown in Fig. 5, the angle between the opposite faces of each wedge being equal to the angular spacing of the vanes; for example, with seventy-two Vanes, the included angle of each wedge will be five degrees. The wedges are arranged to engage at their inner ends with splines 38 upon the exterior of the shaft; as shown in Fig. 5, the splines are each of an angular extent of five degrees, the inner ends of alternate wedges resting upon the outer faces of the splines and those of the intermediate ones being made slightly longer in the radial direction so as to engage in the spaces between the splines, thus securing positive transmission of the rotary motion. The wedge-shaped roots 31 are clamped together by means of hub-plates 39; each of these plates is provided with circumferential grooves 40 on its inner face, the grooves fitting upon arcuate ribs 4I on the opposite faces of the wedges, and the plates being tightly held in engagement with the wedges, for example by pressure exerted along the shaft by a nut 42 securing the inner race of the bearing I1 (Fig. 1).

, Theinvention procures a considerable economy of space and dimensions; the tubular casing 22, into the open front end of which the air enters by ram effect, with initial compression by the primary element constituted by the vanes I3 of the apertured rotor and further compression by the axial-type compressor 2li, 2I inside the casing 22, is made convergent towards the main compressor I2 at the rear end, this main compressor being of relatively small diameter and capacity because of the reduction of Volume of the air corresponding to its increased pressure due to the initial and `secondary compressions. Theinvention also .furnishes a solution of the problem of cooling the turbine rotor (including the blades I5 exposed to the combustion gases), the driving shaft I0 and `its bearings; it follows that at least the central portion of the turbine rotor can be constructed of steelor other material chosen solely from the point of View of its mechanical qualities, without regard to the questions usually involved by high temperature-of operation. Where, however, some pre-heating of the air is desired, the thermal insulation of thecasing 22 (such as indicated at 24) may be dispensed with, wholly or in part, in order to facilitate the transfer of heat from the combustion chamber or chambers I3.

The gases exhausted from the turbine may be collected in amanifold of annular form having one or more lateral outlets for discharging the gases rearwardly; Figs. 7 and 8 illustrate a manifold 43 fitted to the stator of the turbine, having ltwo semi-annular branches of spiral shape leading to two outlets 44 which are united by a Y- pipe fitting 45 to deliver a single jet for the propulsion of the aircraft or the like. Alternatively, the exhaust gases may be led to suitable reaction devices, such devices being utilized for the propulsion of the aircraft or the like, as described in my prior U. S. Patent No. 2,224,260, dated December 10, 1940; an example of such reaction device, in which the gases are reversed through substantially 180 degrees by a deflecting surface, is hereafter described with reference to Fig. 9.

The invention may be applied to gas turbines in general, apart from their employment for jetpropelled aircraft.

Fig. 9 illustrates the arrangement of a turbine differing slightly from that shown in Fig. l, and having the forward end of the turbine shaft H0 fitted with suitable speed-reducing gear IOI for the driving of an airscrew |02, the air flowing from the screw entering with ram effect into the apertured central portion of the rotor III; the rotor is provided with internal vanes I It of helicoidal form, cooperating with stationary guide vanes IIB and constituting the primary element of an axial-type compressor with multiple rotors and stators I2I, insidea tubular casing |22, which delivers the compressed air to the intake of a main compressor I|2 towards the rear end of the turbine shaft. This compressor feeds air through a manifold |28 into the breech ends of the combustion chambers II3 which are spaced in an annular series around the casing |22; a

Vseparately acting compressor |25 formed by the rear face of the main compressor delivers a supply of air through fuel injectors |03 into the breech ends of the combustion chambers, this fuel-injection air being drawn partly from the interior of the hollow shaft IIO by way of holes |21 near its rear end and to a much larger extent from the exterior, as hereinafter described, while the fuel is supplied by the auxiliary equipment |04 driven by the rear end of the turbine shaft. It will be noted that the rear compressor |25 feeds the combustion chamber burners directly with air conveying fuel from the injectors |03, while the main compressor I|2 furnishes air through the manifold |28 at higher speed, at higher pressure and in greater quantity to maintain the names in full activity.

The hot gases. discharged from the combustion chambers I I3 through the stator guide vanes I I4 impinge upon the turbine blades ||5 at the periphery of the rotor III; these blades are preferably arranged for multi-stage expansion of the gases, intermediate guide vanes |I4a being provided between the stages, and the blades II5, II5 of the successive stages being of increasing height and external diameter to `allow for the expansion of the gases, while at the same time increasing the centrifugal effect. After passing the final-stage blades I I5a, the gases are reversed through substantially 180 degrees without shocks or eddies, but taking advantage of the outward diifusion or spread of the gases under centrifugal effect, by means of a curved deflecting vsurface |05; the latter, which serves asa reaction Ydevice to assist the propulsion of the aircraft, directs the gases rearwards over and between the combustion chambers I I3, within the progressively expanding annular space between an outer cylindrical casing |06 and a slightly convergent partition |01 through which the combustion chambers extend, so that the gases flow along the gaps between the said chambers, the rear endof this partition being suitably apertured to allow passage of the compressed air to the injectors |03 and manifold |28. The said injectors and manifold are provided with external fins |08, in addition to internal deectors |29, the whole of these fittings being preferably cast integrally in high-conductivity metal, so that a considerable and intensive heating of the compressed air is effected atthis point by the highspeed gases returning over and between the combustion chambers H3, outside the partition |01. Such a heat-exchanger, with its external surfaces bathed or swept by very .hot gases travelling at high speed, procures a much .more eflicient transfer of heat through the Walls of the injectors and manifold than would their immersion in a hot gaseous medium of low speed or even stagnant.

From the moment when the hot gases have passed the finned injectors |03 fed by the rear compressor |25, they are directed into a jet device of the well-known kind, comprising two coaxial cones |09, I09a with a free space between them leading to a jet outlet for the rearward discharge of the gases at high speed; the outer cone |09 forms a continuation of the casing |06, while the inner cone |0911 forms a `continuation of the partition |07, this inner cone or bullet enclosing the auxiliary equipment |04. The inner cone is stayed in relation to the outer cone by flared or stream-line Webs |26, suitably heat-insulated, but instead of these webs being solid or closed at the ends, they are made hollow (as seen in section in the case of the uppermost one) to connect with openings through the two cones so that the external air can pass freely and substantially unheated to the intake of the rear compressor |25. The openings through the Webs |26 may be made of sufficient size to allow access to the auxiliary equipment for the purpose of inspection and maintenance.

As in the case of the finned injectors |63 and manifold |28, the rotor III may be `cast in tegrally with its rim |33, turbine blades |I5 and compressor vanes I`I5, the whole being made of a high-conductivity metal to facilitate the transmission of heat from the blades to the vanes which are cooled by the air admitted through the rotor. The rim |33 fits into grooves |41 on the adjacent edges of the stator walls, providing a frictionless seal against leakage of the turbine gases to the interior of the rotor.

Due to the fact that the rearward flow of the exhaust gases from the deflecting surface |05, which itself produces a tractive force greater than the static thrust of the jet, takes place mainly in the gaps between the spaced combustion chambers `||3, the diameter of the outer casing |06 need be very little greater than the overall diameter of these chambers; consequentlyI the improved turbine presents the important advantage of relatively small external diameter, with resulting low weight, in addition to the advantages of superior cooling of the moving parts due to the admission of cold air through the apertured rotor, and that of fuel economy due primarily to the combination of a turbine-driven airscreW with the propulsive jet, and secondly to the preheating of the compressed air which is controlled by the design of the fins |08 and by the provision of heat-insulation at suitable points.

Fig. 10 illustrates the arrangement of a gas turbine according to the present invention as applied to the propulsion of an aircraft by a jet device 245 without the provision of an airscrew driven by the turbine shaft; in this case, the shaft 2|0 has the turbine rotor 2|| mounted towards the rear end, and the double compressor 2|2, 225 mounted towards the front end, with the guide vanes 2 |8 and the multiple rotors 220 and stators 22| of the axial-type compressor disposed along the shaft. The combustion chambers 2 |3 are arranged in an annular series around the turbine, their rear ends being tted with guide vanes 2|4 delivering the hot gases upon the turbine blades 2 5 integral with the rim 233 of the rotor; the exhaust gases are discharged to the jet device 245 by way of the free space between two coaxial cones 209, 209a.

The double compressor 2|2, 225 driven by the turbine shaft feeds the combustion chambers 2|3 in a similar manner to that described with reference to Fig. 9, the fuel being supplied by the auxiliary equipment 204 and the air for the compressor 225 being admitted partly through the hollow shaft by way of holes 221 and partly from the exterior through the intake openings 226. The main supply of air to the compressor 2|2 is admitted through the apertured rotor 2|| having vanes 2 I6 which form a primary element of the axial-type compressor; since, however, this rotor 2|'| is arranged at the rear end of the turbine, where the free admission of air is hindered by the inner cone or bullet 209a, the whole of the turbine is encased by a power egg or casing 206 having its forward end 201 left open and slightly flared so that the air enters by ram effect. The air passes over and between the injectors 203, manifold 228 and combustion chambers 2|3, all of which are suitably heat-insulated to prevent loss of heat to the incoming air; as the air travels along the interior of the casing 206, it moves clear of the combustion chambers 2|3 and flows outside the cone 209 at its junction with the turbine stator 205, beyond which the casing 206 is reduced in diameter, its rear end being suitably connected to the outer cone 209. A number of air ducts or inlets 230, of stream-line section, are carried through the cones 209, 209a, between the stator 205 and the rear end of the casing 206, the walls of these ducts acting to hold the two cones in relative position, while the passage ways through the ducts enable the air from the rear end of the casing 206 to pass inwards to the interior of the inner cone 209a and thus to reach the apertured rotor 2| with a considerable initial pressure due to the ram effect at the intake end 201 of the casing.

It will be noted that both air intakes are located at the normal or forward end of the turbine, where the auxiliary equipment 204 is installed, and that the incoming air is maintained substantially unheated until it enters the compressors, heat-insulation being provided where necessary to prevent transfer to heat. The jet outlet 245 at the rear end may be arranged for a slight expansion of the gases under the effect of centrifugal force, by giving the outlet a slightly divergent shape, as indicated at 246. As in the previous examples of construction, the turbine rotor has its annular rim 233 fitting into grooves 241 on the adjacent edges of the stator walls, thus providing a frictionless seal to prevent leakage of the high-velocity gases from the combustion chambers into the interior of the rotor 2| I.

What I claim is:

l. A gas turbine comprising a rotor having peripheral blades, a hollow shaft driven by said rotor, a casing around said shaft, at least one combustion chamber furnishing hot gases, means for applying said gases to said blades in a direction parallel to said shaft for driving said rotor, a main compressor driven by said shaft for compressing air to feed said combustion chamber, said rotor and said main compressor being arranged towards opposite ends of said casing, and an axial-type compressor with multiple rotors and stators inside said casing, said first-mentioned rotor being apertured to admit cold air to said axial-type compressor, and said axial-type compressor forcing the air to said main compressor in a direction opposite to the ow of said gases past said blades, in combination with another compressor driven by said shaft and drawing cold air through said hollow shaft, said other compressor feeding air to at least one combustion chamber furnishing hot gases to drive said firstmentioned rotor.

2. A gas turbine comprising a rotor having peripheral blades, a hollow shaft driven by said rotor, a casing around said shaft, at least one combustion chamber furnishing hot gases, means for applying said gases to said blades in a direction parallel to said shaft for driving said rotor, a main compressor driven by said shaft for compressing air to feed said combustion chamber, said rotor and said main compressor being arranged towards opposite ends of said casing, said combustion chamber being located between said compressor and said blades, and an axial-type compressor with multiple rotors and stators inside said casing, said first-mentioned rotor being apertured to admit cold air to said axial-type compressor, and said axial-type compressor forcing the air to said main compressor in a direction opposite to the flow of said gases past said blades, in combination with another compressor driven by said shaft and drawing coldair through said hollow shaft, said other compressor being nonsymmetrical in relation to said main compressor, said axial-type and main compressors operating in series, and said other compressor operating in parallel to said series-operating axial-type and main compressors.

3. A gas turbine for the propulsion of vehicles including aircraft, comprising a rotor having peripheral blades, a shaft driven by said rotor, a plurality of combustion chambers furnishing hot gases, means for applying said gases to said blades for driving said rotor, an axial-type compressor and two centrifugal compressors, all said compressors being driven by said shaft for compressing air to feed said combustion chambers, said rotor being arranged towards the rear end of said shaft, said centrifugal compressors being arranged towards the forward end of said shaft.

9. saidrotor being apertured for passage of cold air in a' forward direction to said axial-type compressor; a manifold of annular form collecting the-exhaust'` gases flowing in a rearward direction from said rotor blades, and at least one outlet from said manifold for discharging the exhaust gases rearwardly of the vehicle.

4. A gas turbine for the propulsion of vehicles including aircraft, comprising a rotor having peripheral blades, a hollow shaft driven by said rotor; a plurality of combustion chambers furnishing hot gases, means for applying said gases tofsaidblades for driving said rotor, an axial-type compressor, ak main centrifugal compressor and another centrifugal compressor, all said compressors being driven by said shaft for compressing air to feed said combustion chambers, part f said air'being drawn through said hollow shaft, said rotor being arranged towards the rear end of said shaft, said centrifugal compressors being arranged towards the forward end of said shaft, said rotor being apertured for passage of cold air in aV forward direction to said axial-type compressor, a manifold of annular form collecting the exhaust gases fiowing in a rearward direction from said rotor blades, said manifold having two semiannular branches of spiral shape, two outlets from said branches, and means for combining the gases'discharged through said two outlets to deliver a single rearwardly directed jet for the propulsion' of the vehicle.

5. Av gas turbine for the propulsion of vehicles including aircraft, comprising a rotor having peripheral blades, a shaft driven by said rotor, a plurality of combustion chambers furnishing hot gases', means for applying said gases to said blades for driving said rotor, an axial-type compressor and two centrifugal compressors, all said compressors being driven by said shaft for compressing air to feed said combustion chambers, said rotor being arranged towards the forward end of said shaft, said centrifugal compressors being arranged towards the rear end of said shaft, said rotor being apertured for passage of cold air in a rearward direction by ram effect to the intake of said axial-type compressor, a curved defiecting surface for reversing the forwardly exhausting gases from said rotor blades and directing said gases rearwardly between said combustion chambers, and means for collecting said rearwardly directed gases to deliver a rearward jet for the propulsion of the vehicle.

6. A gas turbine for the propulsion of vehicles including aircraft, comprising a rotor having peripheral blades, a shaft driven by said rotor, a plurality of combustion chambers furnishing hot gases,A means for applying said gases to said blades for driving said rotor, an axial-type compressor and two centrifugal compressors, all said compressors being driven by said shaft for compressing air to feed said combustion chambers, said rotor being arranged towards the forward end of said shaft, said centrifugal compressors, being arranged towards the rear end of said shaft, said rotor being apertured for passage of cold air in a rearward direction by ram effect to the intake of said axial-type compressor, a curved deflecting surface for reversing the forwardly exhausting gases from said rotor blades and directing said gases rearwardly past said combustion chambers, means for effecting heat-exchange between said rearwardly directed gases and the air fed to said combustion chambers by said centrifugal compressors, and means for collecting said rearwardly directed gases to deliver a rearward jet for the propulsion of the vehicle.

7. A gas turbine for the propulsion of4 aircraft, comprising a rotor having peripheral blades arranged for at least two stages of gas expansion, a hollow shaft driven by said rotor, a plurality of spaced combustion chambers furnishing hot gases, means for applying said gases to said blades for driving said rotor, an axial-type compressor and two centrifugal compressors driven by said shaft for compressing air to feed said combustion chambers, part of said air being drawn through said hollow shaft, said rotor being arranged towards the forward end of said shaft, said centrifugal compressors being arranged towards the rear end of said shaft, an airscrew at the forward end of said shaft, said airscrew being driven by said' shaft, said rotor being apertured for passage of cold air from said airscrew in a rearward direction to the intake of said axial-type compressor, a curved deiiecting surface receiving the forwardly exhausting gases from said rotor blades, said deflecting surface being adapted to reverse the flow of said gases by taking advantage of their outward diffusion under centrifugal eiect and to direct them rearwardly around said combustion chambers, an outer casing around said combustion chambers, said casing being adapted to enclose said rearwardly directed gases and to force them into the gaps between said spaced combustion chambers, and means for collecting the gases rearwards of said combustion chambers to deliver a single jet for the propulsion of the aircraft.

8. A gas turbine for the propulsion of aircraft, comprising a rotor having peripheral blades, a hollow shaft driven by said rotor, a plurality of combustion chambers arranged in an annular series around said shaft, means for applying hot gases from said combustion chambers to said blades in a direction parallel to said shaft for driving said rotor, means for collecting the gases in rear of said blades to deliver a jet for the propulsion of the aircraft, an axial-type compressor and two centrifugal compressors, all said compressors being driven by said shaft for compressing air to feed said combustion chambers, part of said air being drawn through said hollow shaft by one of said centrifugal compressors, said rotor being arranged towards the rear end of said shaft, said centrifugal compressors being arranged towards the forward end of said shaft, said rotor being apertured for passage of air in a forward direction to the intake of said axial-type compressor, an outer casing around said combustion chambers, the forward end of said casing being open for admission of cold air by ram effect, said air flowing along the gaps between said combustion chambers, and passages leading from the rear end of said casing to convey said air to said apertured rotor, said passages extending transversely through said collecting means.

9. A gas turbine comprising a rotary shaft, an axial-type compressor with multiple rotors and stators, the rotors of said compressor being driven by said shaft, a stator element at the inlet end of said compressor, a rotor element adjacent to said stator element, said rotor element including turbine blades upon its periphery and helicoidal spokes secured to said shaft, two centrifugal compressors driven by said shaft and located at the other end of said axial-type compressor, said helicoidal spokes co-operating with said stator element to force air into said axial-type compressor for feeding one of said centrifugal compressors, combustion chambers arranged to re-l ceive air from said centrifugal compressors and to furnish hot gases of combustion, and means for applying said hot gases to said turbine blades for driving said rotor element and shaft, the direction of said hot gases in relation to said rotor element being opposite to the direction of said air in relation thereto.

10. A gas turbine comprising a rotary shaft, an axial-type compressor with multiple rotors and stators, the rotors of said compressor being driven by said shaft, a stator element at the inlet end of said compressor, a rotor element adjacent to said stator element, said rotor element including peripheral turbine blades integral with helicoidal spokes secured to said shaft, two centrifugal compressors driven by said shaft and located at the other end of said axial-type compressor, said helicoidal spokes co-operating with said stator element to force air into said axial-type compressor for feeding one of said centrifugal compressors, combustion chambers arranged to receive air from said centrifugal compressors and to furnish hot gases of combustion, and means for applying said hot gases to said turbine blades for driving said rotor element and shaft, the direction of said hot gases in relation to said rotor element being opposite to the direction of said air in relation thereto.

l1. A gas turbine comprising a rotary shaft, an axial-type compressor with multiple rotors and stators, the rotors of said compressor being driven by said shaft, a stator element at the inlet end of said compressor, a rotor element adjacent to said stator element, said rotor element including turbine blades upon its periphery and helicoidal spokes secured to said shaft, two centrifugal compressors driven by said shaft and located at the other end of said axial-type compressor, said helicoidal spokes co-operating with said stator element to force air into said axial-type compressor for feeding one of said centrifugal compressors, combustion chambers arranged to receive air from said centrifugal compressors and to furnish hot gases of combustion, and means for applying said hot gases to said turbine blades for driving said rotor element and shaft, said gas-applying means including baffles for preventing leakage of said hot gases adjacent to said rotor element into the air forced into said axial-type compressor,

and the direction of said hot gases in relation to y said rotor element being opposite to the direction of said air in relation thereto.

l2. A gas turbine comprising a hollow rotary shaft, an axial-type compressor with multiple rotors'and stators, the rotors of said compressor i being driven by said hollow shaft, a stator element at the inlet end of said compressor, a rotor element adjacent to said stator element, said rotor element including turbine blades upon its periphery and helicoidal spokes secured to said shaft, two centrifugal compressors driven by said shaft and located at the other end of said axial-type compressor, one of said centrifugal compressors drawing air through one end of said hollow shaft, the other of said centrifugal compressors receiving air from said axial-type compressor, said helicoidal spokes co-operating with said stator element to force air into said axial-type compressor, combustion chambers arranged to receive air from said centrifugal compressors and to furnish hot gases of combustion, and means for applying said hot gases to said turbine blades for driving said rotor element and shaft, the direction of said hot gases in relation to said rotor element being opposite to the direction of said air in relation thereto.

13. A gas turbine comprising a hollow rotary shaft, an axial-type compressor with multiple rotors and stators, the rotors of said compressor being driven by said hollow shaft, a stator element at the inlet end of said compressor, a rotor element adjacent to said stator element, said rotor element including turbine blades upon its periphery and helicoidal spokes secured to said shaft, a centrifugal compressor driven by said shaft and located at the other end of said axialtype compressor, said helicoidal spokes co-operating with said stator element to force air into said axial-type compressor for feeding said centrifugal compressor, another centrifugal compressor driven by said shaft and located back-to-back of said first-mentioned centrifugal compressor, said other centrifugal compressor drawing air through one end of said hollow shaft, at least one combustion chamber arranged to receive air from said centrifugal compressors and to furnish hot gases of combustion, and means for applying said hot gases to said turbine blades for driving said rotor element and shaft, the direction of said hot gases in relation to said rotor element being opposite to the direction of said air in relation thereto.

JULES ANDR NORBERT GALLIOT.

REFERENCES CITED The following references are of record in the le of this patent:

UNITED STATES PATENTS Number Name Date 2,080,425 Lysholm May 18, 1937 2,224,260 Galliot Dec. 10, 1940 2,326,072 Seippel Aug. 3, 1943 2,396,068 Youngash Mar. 5, 1946 2,405,164 Pavlecka Aug. 6, 1946 2,410,804 Baumann Nov. 12, 1946 2,423,183 Forsyth July l, 1947 2,428,330 Heppner Sept. 30, 1947 2,455,458 Whittle Dec. 7, 1948 2,477,798 Griffith Aug. 2, 1949

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