US2811833A - Turbine cooling - Google Patents

Description

Nov. 5, 1957 w. s. BROFFITT 2,811,833

TURBINE COOLING Filed June 5, 1953 3 Sheets-Sheet 1 lNvE'N-roR ATTORNEY NOV 5', 1957 w. s. BRoFFlTT 2,811,833

TURBINE COOLING Filed June 5, 1953 5 Sheets-Sheet 2 Nov. 5, 1957 W. S. BROFFITT TURBINE COOLING 3 Sheets-Sheet 3 Filed June 5, 1953 ff piaf United States Patent O TURBINE COOLING Wilgus S. Brott, Indianapolis, Ind., assignor to General Motors Corporation, Detroit, Mich., a corporation of Delaware Application June 5, 1953, Serial No. 359,746.

2 Claims. (Cl. 60-39.66)

This invention relates to gas turbine engines, and more particularly to an arrangement for cooling the turbine of such an engine.

An object of this invention is to provide means whereby the turbine wheel may be cooled by air from the compressor of the engine.

Another object of this invention is to provide means whereby the flow of turbine cooling air may be varied as desired.

Further objects and advantages of the present invention will be apparent from the following description, reference being had to the accompanying drawings wherein a preferred form of the present invention is clearly shown.

In the drawings:

Fig. 1 is a partial plan view, partially broken away, of a known type of turbojet engine incorporating the invention;

Fig. 2 is an enlarged fragmentary section taken substantially along the line 2-2 of Fig. 4 and also constituting an enlargement of a portion of Fig. l;

Fig. 3 is an enlarged fragmentary section taken substantially along the line 3--3 of Fig. 2;

Fig. 4 is a reduced fragmentary section taken substantially along the line 4 4 of Fig. 2; and

Fig. 5 is an enlarged fragmentary section taken substantially along the line 5 5 of Fig. 4.

Referring now to the drawings in detail, the turbojet engine includes a centrifugal compressor rotor and an axial ow turbine rotor 12 suitably supported by bearings 14 in the stationary frame and casing structure of the engine. The casing and frame structure includes an accessory housing 16 secured to the compressor housing which includes a central diffuser housing 18 and front and rear air inlet housings 20 and 22. The air inlet housings 20 and 22 support the front and rear bearings of the compressor rotor 10 and connect with the diffuser casing 18 and the compressor outlet casings 24 and 26 by struts 28 and 30 which bridge the inlets of the compressor. The air inlets are provided with screens 32 and 34 to prevent foreign object damaging of the compressor blading 36 and 38. The rear air inlet housing 22 is suitably secured to the forward end of the inner and outer shells 40 and 42 of the turbine shaft housing. A turbine nozzle casing 44 is supported by the outer shell 42 of the turbine shaft housing and provides support for the external casing 46 of the turbine. The inner tail cone 48 is supported from the exhaust portion of the turbine casing 46 by radial struts not shown. A plurality of combustors arranged around the turbine shaft housing include the flame tubes 50 concentrically supported in the outer air tubes 52.

The engine operates in the usual manner, air from the compressor rotor 10 being fed through the diifuser-turning vanes 54 and into the perforated ame tube 50 by way of the combustor air tubes 52. Fuel from the spray nozzles 55 is burned with the air in the ame tubes 50 and the motive fluid is fed by the turbine nozzle vanes 56 through the turbine rotor blading 58 to drive the compressor rotor 10 through a splined coupling 60; the exice haust of the turbine furnishing jet motive power for the aircraft.

The combustion apparatus is supplied with a large excess of air over that required for combustion to reduce the temperature of the motive fluid and thereby prevent damage to the turbine. The maximum temperature of the motive fluid may be somewhat greater and the efliciency of the turbine be thereby increased if the turbine rotor is provided with some type of cooling system. In accordance with the invention a portion of the air flow from the compressor is by-passed by the ame tubes and is led to the forward face of the turbine rotor to cool the same before entering the motive iluid stream. The invention includes an arrangement whereby the flow of cooling air to the turbine rotor may be varied to suit different operating conditions of the engine.-

The turbine nozzle and the rearward ends of the combustors are supported by an annular ange 62 secured to the turbine shaft outer housing 42 by bolts 64. An annular U-sectioned channel 66 and L-sectioned rings 68, 70 and 72 are secured to the annular ange 62 by bolts 74. The inner shroud ring 76 is welded to ring 68 and the nozzle blades 56 are welded between the inner and outer shroud rings 76 and 78. The L-sectioned rings 70 and 72 form an annular chamber A which is fed with cooling air from the annular chamber B of the U-sectioned member 66 by passages C in the channels 66, 68 and 70. The ange 62 is provided with passages D which are in registry with the passages C and which connect by conduits to the air jackets 52. A valve ring 92 is rotatably received in the interior of the channel 66 and bears against the rear face of the ange 62. The valve ring 92 is provided with passages E which may be moved in and out of registry with the passages D of the flange 62 so that cooling air from the compressor may be supplied to the chamber A in varying amounts. If lower temperature cooling air is desired, the conduits 90 may be connected to the diffuser portion of the compressor instead of the combustor air jackets 52.

The cooling air from the chamber A bathes the rim of the front face of the turbine rotor 12 to cool the same and is conducted through the hollow interiors of the turbine buckets 58 by passages 96 inthe turbine rotor, by passages 98 in the bucket roots 100, and by outlet passages 102 in the bucket blades. The flanges 70 and 72 are provided with labyrinth seals 104 and 106 to reduce radial air leakage at the front face of the turbine rotor. Circumferential slots 108 in the valve ring 92 permit suflicient rotation thereof to place the passages D and E in and out of registry without interference by the bolts 74. A sector gear 110 meshes with an internally geared portion 112 of the valve ring 92 to impart rotation thereto upon actuation of the sector gear shaft 114 by the lever 116. A peripheral slot 118 in the channel 66 allows gear engagement. The sector gear shaft 114 is journaled by a bearing 120 on the flange 62 and a bearing 122 in a bracket 124 secured on the turbine shaft inner housing 40 by bolts 126. A pair of stops 128 and 130 cooperate with a lug 132 of the actuating lever 116 to limit the rotation of the ring valve. The ring valve 92 may be positioned by manual or automatic means as desired; for example, the actuating rod 134 may be operated by the throttle of the engine, by manual action or by any suitable device that reflects the output of the engine. The flow of cooling air should increase with increases in engine output as the high motive fluid temperature encountered at higher outputs necessitates increased cooling of the turbine rotor. An oversupply of cooling air would reduce the overall eiciency of the engine and maximum eiliciency at different operating conditions is achieved by varying the supply to meet the particular operating condition. Y

While the preferred embodiment of the invention has been described fully in order to explain the principles of the invention, it is to be understood that modications in structure may bemade by the exercise'of skill ,inthe .-art within the scope of the invention, which is not toeberegarded as limited by the detailed description of thejpreferred embodiment.

I claim: Y

1. A turbojet engine comprising an engine housing, a ,turbine and compressor having interconnected rotors sup- Vported therein, said turbine having an annular inlet noz- Vzle, a plurality of combustorsconnecting said compressor to said turbine nozzle and delivering high temperature motive gases thereto, each of said combustors havinga flame `tube spaced concentrically -in an outer air jacket, labyrinth sealing structure secured radially inward of said `turbine nozzlerand forming anannularchamber proximate `the rim of the front face of said turbine rotor, conduits for conducting cooling air Vfrom each of said jackets to said i chamberfand valve means secured radially inward of'said vturbine nozzle for controlling the delivery of said lcooling vair from said conduits to said chamber including a ring adjacent said chamber and having passages therein rotatable'in and out of registry with said conduits,

2. A turbojet engine comprising an engine housing, a turbine and compressor having interconnected rotors supported therein, said turbine having an annular inlet nozzle, combustion apparatus connecting said compressor to said turbine nozzle and delivering high temperature motive gases thereto, labyrinth sealing structure secured to said housing radially inward of said nozzle and forming an annular chamber proximate the rim of the front face of said turbine rotor, a plurality of conduits for conducting cooling air from said compressor to said chamber, valve means secured radially inward of said nozzle for controlling the delivery of -said cooling air to saidcharnber including an internally geared ring rotatably supported by said housing adjacent said chamber and having a plurality ofjpassages therein for connecting and disconnecting said conduits and said chamber on rotation of said ring, and a sector gear rotatably mounted on said housing radially inward of said ring and meshed with said ring for actuation thereof.

'References Ctedin the file of this patent UNITED STATES PATENTS 2,112,391 Anxionnaz Mar. 29,` 1938 v2,445,837 McKenzie July 27, 1948 2,457,157 King Dec. 28, "1948 2,594,765 Goddard Apr. 29, 1952 2,625,793 Mierley et al Jan. 20, '1953 2,656,096 Schwarz Oct. 20, 1953

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