US2712729A

US2712729A - Cooling systems of gas-turbines

Info

Publication number: US2712729A
Application number: US323418A
Authority: US
Inventors: Wigg Raymond Ernest
Original assignee: Ruston and Hornsby Ltd
Current assignee: Ruston and Hornsby Ltd
Priority date: 1951-12-06
Filing date: 1952-12-01
Publication date: 1955-07-12
Anticipated expiration: 1972-07-12

Description

July 12, 1955 R. E. WIGG COOLING SYSTEMS OF GAS-TURBINES 4 Sheets-Sheet 1 FJ' led Dec. 1, 1952 47%rney July 12, 1955 R. E. WIGG 2,712,729

COOLING SYSTEMS OF GAS-TURBINES Filed Dec. 1, 1952 4 Sheets-Sheet 2 so. p

l 45 #6 4a b 4 /nver/7:r A, K E. W/66- r Mwfi7 July 12, 1955 R. E. wise 2,712,729

COOLING SYSTEMS OF GAS-TURBINES Filed Dec. 1, 1952 4 Sheets-Sheet 3 July 12, 1955 5, W1 3 2,712,729

COOLING SYSTEMS OF GAS-TURBINES Filed Dec. 1, 1952 4 Sheets-Sheet 4 In re 3 r R. (QM/A5161 y woos/77 United States Patent 2,712,729 COOLING SYSTEMS oF GAS-TURBINES Raymond Ernest Wigg, Lincoln, England, assignor to Rnston & Homsby Limited, Lincoln, England, a British company Application December 1, 1952, Serial No. 323,418

Claims priority, application Great Britain December 6, 1951 2 Claims. (Cl. fill-39.66)

This invention relates to gas-turbine engines including a compressor, a turbine in driving connection therewith and a combustion chamber, and its primary object is to provide for the cooling of thin-walled ducts conveying gas at high temperature and pressure, and more especially a a duct conveying the working fluid from the combustion chamber to the turbine, by means of air bled from the compressor at a substantially lower pressure than that of the gas within the duct, and therefore at a temperature low enough to provide eflicient cooling, while at the same time substantially equalising the external and internal pressures to which the wall of the duct is subjected, so that said wall may be made thin enough to transmit heat readily and to avoid accumulation of thermal stresses.

The manner in which this object, and such others as may hereinafter appear, are attained will appear from the description which follows of a specific embodiment of the invention, given by way of example only and without limitation of the scope of the invention, as defined in the appended claims, and having reference to the accompanying drawings, of which,

Figure 1 is a plan view of the charging unit of a gas turbine engine, comprising a compressor, a combustion chamber and a compressor-driving turbine;

Figure 2 is an axial section of the compressor-driving turbine and turbine inlet duct, taken on the line 2-2 of Figure 3;

Figure 3 is a transverse section taken on the line 3-3 of Figure 2;

Figure 4 is a partial axial section of the compressor, part broken away; and

Figure 5 is a somewhat diagrammatic longitudinal section of the combustion chamber.

Referring to Figure 1, the charging unit comprises an air inlet casing 10, an axial flow compressor 11, an intermediate casing 12, unitary with which is a compressor outlet volute 12a, a turbine inlet duct casing 13, a compressor-driving turbine 14, an elbow 15 connected to the compressor outlet volute 12a and leading into a combustion chamber 16, and a combustion chamber outlet duct 17 connected to the turbine inlet duct enclosed in casing 13. The exhaust of the turbine 14 forms the working fluid for a separate power turbine in which it is further expanded. The power turbine does not form part of the invention and is not illustrated.

Referring now to Figure 2, the turbine inlet duct enclosed in casing 13 is a thin-walled volute 18 having an annular outlet 19 communicating with the nozzle ring 20 of the turbine 14. The latter has two stages and includes two rotor discs 21 carrying rotor blades 22, a segmental stator ring 23, carrying stator blades 24 and supported by an outer stator casing 25. The rotor discs are mounted on a shaft 26 secured to a journal 27 supported in a bearing member 28, which is secured to an inner cylindn'cal part 12b of the intermediate casing 12. An exten- "on shaft 29 secured to the journal 27 is directly coupled the shaft 30 of the compressor (Figure 4).

Referring to Figure 4, the multi-stage axial flow compressor 11 comprises an outer casing 31 blading 32 and a built-up drum rotor 33 blading 34 and mounted on shaft 30. An annular diffuser 35 at the outlet end of the compressor delivers into the volute 12a shown in Figure 1, the inner wall of the diffuser being constituted by a stationary inner casing member 35a shown in elevation in Figure 4. An annular channel 36 is formed in the outer casing 31 and communicates through openings 37 with the interior of the compressor annulus at the fourth stage of the stator blading; and a pipe 38 communicating with the channel 36 bleeds off cooling air at a relatively low pressure and temperature.

Referring to Figure 5, the combustion chamber 16 comprises an outer casing 39, within which is arranged a flame tube 40, and baffle 41. The right hand end (as seen in the figure) of the combustion chamber is connected to the outlet duct 17 which has an outer pressure-resistant skin 42 and an inner-thin skin 43 forming a continuation of the flame tube. Air from the compressor outlet elbow 15 (see Figure 1) enters the space 44 at the right hand end of the combustion chamber, and the flow of air and combustion products is indicated by arrows as follows:

supporting stator supporting rotor tailless arrow one-tailed arrows two-tailed arrows three-tailed arrows four-tailed arrows ucts) five-tailed arrows The secondary combustion air and dilution air enters the flame tube through holes (not shown) in its wall and enters the continuation of the flame tube formed by the inner skin 43 of the outlet duct 17 through a gap between the end of the flame tube 40 and the inner skin 43. Disregarding the small pressure-drop that occurs in the combustion zone, all the air and gas represented by the arrows is at the same pressure, viz. the pressure at the outlet of the diffuser 35, i. e. the maximum pressure of the cycle. The cooling air is at the temperature of exit from the diffuser except in so far as it is heated by heat transfer through skin 43, the hot gas within skin 43 being at the temperature produced by combustion, i. e. the maximum temperature of the cycle. The outlet end of the tube formed by the inner skin 43 registers with the inlet end of the turbine inlet volute 18, but there is a small gap between them through which the cooling air in the space between skins 42 and 43 enters the volute 18 as dilution air.

Referring again to Figure 2 (and Figure 3), the volute 18 is enclosed in an intermediate casing formed by a skin 45, the space 47 between this skin and the wall of the volute 18, which is filled with porous lagging, being completely enclosed, except where it communicates with the space between the outer and inner skins 42, 43 of the combustion chamber outlet duct 17 (see arrows, Figure 2. and Figure 5), the outer skin 42 being connected by a flange joint with the turbine inlet duct casing 13.

The intermediate casing is surrounded by an outer casing, formed partly by the structural casing 13 and partly by a skin 46, between which outer casing and the intermediate casing 45 is a narrow space 48, which communicates through a gap 49 with the space 50 between the outer part of casing 12 and its inner cylindrical portion 12b. The narrow space 48 also communicates with passages 51 formed in the outer stator casing 24 and having an external outlet.

The pipe 38 (see also Figure 4) communicates with the space 50, as shown in Figure 3.

Cooling air entering space 50 from pipe 38 at low pressure and temperature, enters the narrow space 48 through '2 a gap 49 and after flowing through'this narrow space leaves it by the passages 51.

It will be seen that the space 47 contains stagnant air substantially at the exit pressure of the diffuser 35 which difiers very little from the pressure of the hot gas within the doctor volute 18, so that the wall of the duct 18 is subjected to no severe pressure difference and can therefore be of light gauge to relieve thermal stresses.

The pressure in space 48 on the other hand is relatively low, being that of the fourth stage of the compres 1 sor which has in all thirteen stages. Consequently, the intermediate casing 45 is subjected to a severe pressure difierence and must be correspondingly robust and is therefore relatively thick. The pressure diiference across the skin 4:: is only that between atmospheric pressure and the fourth stage of compression and skin 46 can consequently be relatively thin.

I claim:

1. In a gas-turbine engine including a compressor, a turbine in driving connection therewith, a combustion chamber and a thin-walled duct adapted to convey gas at elevated temperature and pressure from the combustion chamber to the turbine, an intermediate pressure-resistant casing surrounding said duct and separated therefrom by a first space, an outer casing surrounding said pressureresistant casing and separated therefrom by a second space, said first space communicating with the compressor outlet, but having itself no outlet so that the internal pressure in said first space approximates to that of the gas within the duct and no flow of air can take place through said first space, a bleed from a part of the compressor, at which the internal pressure is substantially less than that of the gas within the duct, means connecting said bleed with said second space, and an outlet from said second space.

2. A gasturbine engine as claimed in claim 1, in which the first-mentioned space is filled with porous lagging material.

References Cited in the tile of this patent UNiTED STATES PATENTS