US2640319A - Cooling of gas turbines - Google Patents

Cooling of gas turbines Download PDF

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US2640319A
US2640319A US76161A US7616149A US2640319A US 2640319 A US2640319 A US 2640319A US 76161 A US76161 A US 76161A US 7616149 A US7616149 A US 7616149A US 2640319 A US2640319 A US 2640319A
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air
cooling
turbine
stream
passage
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US76161A
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George F Wislicenus
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Packard Motor Car Co
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Packard Motor Car Co
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12Cooling of plants
    • F02C7/16Cooling of plants characterised by cooling medium
    • F02C7/18Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • Y02T50/675

Description

June 1953 G. F. WISLICENUS COOLING OF GAS TURBINES 4 Sheets-Sheet 1 Filed Feb. 12, 1949 Mala/.141!!! INVENTOR. v

June 1953 G. F. WISLICENUS COOLING OF GAS TURBINES 4 Sheets-Sheet 2 Filed Feb l2 1949 IN VEN TOR.

kcerzzw BY 94w June 2, 1953 v GLF. WISLICENUS 2,640,319

' COOLING OF GAS TURBINES Filed Feb. 12, 1949 4 Sheets-Sheet 5 1.953 G. F. WISLICENUS 2,640,319

COOLING OF GAS TURBINES Filed Feb. 12, 1949 4 Sheets-Sheet 4 IN V EN TOR.

N Geozgef flzlalazema BY I QM 4w IMMW Patented .Fune 2, i953 COOLENG 0F GAS TURBENIES Application February 12, 1949, Serial N 0. "76,161

2 Claims.

The invention relates generally to engines of the gas turbine type and more particularly to the cooling thereof.

The general object of the invention is to provide a novel structure for conducting cooling air to the turbine portion of such an engine at such pressure or pressures as will permit its eventual discharge into the power stream.

Another object is to provide a novel arrangement of parts permitting the cooling air utilized in the turbine portion of the engine and discharged into the power stream, to be taken from points in the path of the air stream in the engine where the pressures are at the minimum necessary, in view of any additional head imparted thereto, to effect discharge into the power stream.

A further object is to provide an engine structure in which the cooling air is taken at a point where its pressure is at the minimum necessary to effect the desired flow thereof, so that a minimum of energy is thereby expended upon it.

Still another object is to provide an engine structure in which the cooling air utilized in different parts of the engine may be taken from the air stream at different pressure levels, determined by the pressures necesary to effect its discharge and by any additional head imparted thereto after withdrawal from the main air stream.

Other objects and advantages will become apparent from the following description taken in connection with the accompanying drawings, in which:

Figure l is a fragmentary longitudinal sectional view of the front end portion of a turbo-jet engine embodying the features of the invention.

Fig. l-A is a continuation of Fig. 1 and shows the intermediate portion of the engine.

Fig. l-B is a continuation of Fig. l-A and illustrates a still further portion of the invention.

Fig. 2 is a fragmentary transverse sectional View taken on the line of Fig. 1.

Fig. 3 is a fragmentary transverse sectional View taken on the line of Fig. l-A.

Fig. 4 is another fragmentary transverse sectional view taken on the line i i of Fig. l-A.

In engines of the gas turbine type, the cooling of various parts of the apparatus, and particularly the turbine rotor, is an important factor because of the high temperatures of the coznbustion gases which drive the turbine. The high temperatures of these gases, of course, reduce the strengths of the materials of the various parts of the rotor at which they may safely operate. The materials utilized are selected from those which permit high operating stresses even at relaill 2 1 tively high temperatures. Nevertheless, even with such materials, the safe operating stresses thereof are lower than at reduced temperatures. Consequently, it is advisable to provide adequate cooling of the parts subjected to the heat of the combustion gases to obtain a maximum safe working stress. This is particularly true in connection with rotating parts since, obviously, by providll'lg adequate cooling thereof, the mass of material may be reduced, thus reducing centrifugal stresses.

Merely by way of example, the temperature of -ie combustion gases contacting the blades of a turbine of this type may be as high as 1600 F., which temperature is so high that the metal should not be permitted to operate at such a temperature. Steel is the material commonly employed for the blades and other parts of the turbine rotor so that the actual operating temperature of the rotor and its parts should be maintained considerably below the temperature of the gases. Such cooling of the rotor parts may be accomplished by subjecting them to contact with a cooling fluid. Since air in large quantities is taken into an engine of this character to support the combustion, a portion of such air is utilized as the cooling fluid. Such air may be taken from the air stream provided to support combustion and, because of the heat absorbed thereby in cooling the rotor parts, it is preferably discharged into the power stream to minimize loss of eficiency of the engine.

In utilizing air from the main air stream created in the engine, consideration must be given, of course, to the pressure of such air so that it may discharge into the power stream at a sufficient pressure to effect such discharge. The pressure at which such air is withdrawn from the air stream, consequently, will depend upon whether or not any additional head is imparted thereto in cooling the parts of the turbine rotor. Thus, if no head is added to a coolin air stream when performing its cooling function, the air for such cooling stream should be withdrawn from the main air stream. at the maximum pressure thereof so that it will have sufficient pressure to enter the power stream after performing its cooling function. If, however, any pumping action is exerted on a stream of cooling air in its passage through or around the turbine rotor, then such cooling air may be withdrawn from the main air stream at a point of lower pressure with the sum of the two pressures being sufiicient to effect the eventual discharge of the cooling stream into the power stream. While, of course, the last- 3 mentioned cooling stream could be taken from the main air stream at a point of higher pressure, energy would be wasted in bringing the cooling air stream to such pressure if the ultimate pressure of the cooling stream were in excess of that needed to efiect its discharge into the power stream. For that reason, it is desirable to withdraw any stream of cooling air from the main air stream at the point of minimum pressure necessary to efiect its eventual discharge into the power stream with consideration given to any additional head that may be added to such cooling air stream.

The engine shown in the drawings and embodying the features of the present invention comprises an intake section I (see Fig. 1), in which is located air compressor means including an axial flow compressor H carried on a main drive shaft 12. The air stream created thereby flows through an annular passage l3 to a second compressor l4 of the combined axial and radial flow type, from which the air stream is discharged at its maximum pressure. From the compressor 14 the air stream passes through a collector I5 (see Fig. l-A) and thence to combustion chamber means, indicated generally at Hi. In the combustion chamber means [6, fuel is introduced into the air stream and combustion takes place to create the power stream for driving the turbine, indicated generally in Fig. 1-13 at H. After the power stream passes through the turbine II, it, of course, is discharged through a tail spout, as indicated generally at l8. The compressor 14 is also mounted and driven by the shaft l2 which, in turn, is driven by the turbine 11.

In the present construction, the inlet 10, in which the compressor H is mounted, comprises a separate annular casting bolted to an intermediate compressor casing, indicated generally at 20, comprising an outer shell 2l and an inner shell 22 interconnected by a plurality of radially extending guid vanes 23. The compressor is is enclosed by a stationary casing 24 extending to and enclosing a portion of the collector i5. To stiffen the intermediate compressor casing 20, a plurality of hollow bulged portions 25 are provided thereon which form, with the wall 2| and the stationary casing 24, a rearwardly and outwardly extending passage 26.

The collector I5 includes collector vanes 30 (see Figs. l-A and 3) which, at their forward ends, engage the rearward portion of the stationary casing 24 and are integrally formed with an outer shell portion 3! constituting a continuation of the stationary casing 24. At its front end, the outer shell 3| has a short cylindrical portion 32 and a flanged portion 33 which is bolted to a matching flange 34 on the intermediate compressor casing 20. At its rear end, the outer shell 3| is also provided with a flange 35, while the inner wall of the collector i5 extends rearwardly and tapers inwardly so that it may be rigidly secured as at 36 (see Fig. l-A) to a main frame structure 31, in which the main drive shaft 12 is journaled.

The combustion chamber means it comprises an outer annular casing 40 and an inner casing 4|. Mounted within the annular space between the inner and outer casings 40 and 4| is a plurality of combustion chamber members 42 arranged in circumferentially spaced relation to one another. The combustion chamber members 42, at their downstream ends, fit within an inner shell 43 which is spaced from the outer casing 40 to provide an annular passage 44. Outside of the outer casing 40 is an outer shell 45 spaced from the outer casing 40 to provide another annular passage 45. The inner casing 41, at its upstream end, engages the collector l5 but tapers downstream at a different angle from the collector so as to provide an annular passage therewith, indicated at 41. The passage 41 continues downstream from the collector and its inner wall is provided by an inner shell 50 radially spaced from the wall 41, as is evident in Fig. l-A.

Adjacent the turbine IT, the inner shell 50 is secured to a nozzle ring 5! (see Fig. l-B) which supports a plurality of nozzle blades or vanes 52. The outer ends of the blades 52 are mounted in a supporting ring 53 constituting a continuation of the inner shell 43. The nozzle blades are hollow and, of course, direct the power stream from the combustion chamber means in proper relation to the blades of the turbine H.

The outer casing 40 is secured to a turbine casing member 54 which provides, with the supporting ring 53, a continuation of the passage 44. Outside of the turbine casing member is a turbine shell 55 secured to and constituting a continuation of the outer shell 45 and providing with the turbine casing 54 a continuation of the passage 46.

The turbine I] in the present instance is illustrated as a two-stage turbine comprising a rotor body so bolted to a flang 6| on the rear end of the main drive shaft I2. The rotor body BI] is provided with a pair of annular flange portions 62 which are axially spaced from one another and which respectively support turbine blades 63 constituting the first stage, and turbine blades 84 constituting the second stage. The turbine blades 63 and 64 are hollow and are open at their outer ends but are surrounded by a stationary shroud 65. Located between the two stages of turbine blades is a series of interstage guide vanes 66 supported at their outer ends by the turbine casing member 54 and extending inwardly through the shroud 65. The shroud 65, at its downstream end, is provided with a ring H! which is shaped to meet the turbine shell 55, the latter extending downstream from the ring 10 and providing the tail spout. The interstage guide vanes 66, at their inner ends, support a hollow ring structure H positioned between the annular flange portions 62 of the rotor body 50.

To provide for suitable cooling of the various parts of the turbine, it is desirable to conduct a plurality of streams of cooling air respectively arranged to cool different parts of the turbine. In the present instance, three such streams are provided, one for cooling the upstream face portion of the turbine rotor 60, another for cooling the interstage facial portions of the turbine rotor, and the third for cooling the interior of the rotor and the two stages of the blades 63 and 64. The first two cooling streams are also arranged respectively to cool the nozzle vanes 52 and the interstage guide vanes 66. All of these streams of cooling air are eventually discharged into the power stream so that the energy therein is not completely lost.

The two streams of cooling air which are respectively used for cooling different facial portions of the turbine rotor are not subjected to any pumping action while performing their cooling function. Consequently, these two streams are withdrawn from the main air stream created by the compressors H and 14 at the maximum pressure thereof so that there will be sulficient pressure in such cooling streams to effect their entrance into the power stream after performing their cooling function. The air stream utilized for cooling the interior of the turbine rotor and the two sets of blades 63 and G l is subjected to a pumping action clue to the rotation of the turbine rotor. Such pumping action, of course, Will increase the pressure of this cooling stream and, consequently, the air therefor may be withdrawn from the main air stream at a point of lower pressure than exists at the discharge of the compressor Hi.

The cooling air stream first referred to, which is adapted to cool the upstream facial portions of the turbine rotor 63, is withdrawn from the main air stream at points where such air stream passes between the combustion chamber members 42. To this end, the inner casing ll (see Fig. l-A) is provided with sets of apertures 8i] and 8! to admit air from the main air stream into the upstream end of the passage ll. The cooling air is conducted downstream by this pas sage and at the downstream end of the passage it is permitted to enter a plurality of elbowshaped tubes 32 (see Fig. 1-3) each having one end mounted in the nozzle ring 5!. Mounted within the nozzle ring 55 is a casing having a cylindrical wall portion 83 and a radially extending wall 84. form an annular chamber with the nozzle ring 5|. The elbow-shaped tubes 82 extend through the annular chamber just mentioned and have their other ends mounted in the radially extending wall 84 so that the air carried through such tubes is discharged onto the front face of the turbine rotor 60. Sheet metal nozzle members 35 are preferably formed over the ends of the tubes 82 so as to direct the air tangentially in the direction of rotation of the turbine rotor whereby a minimum resistance to air how is provided. The air thus discharged from the tubes 82 cools the front face of the turbine rotor 60 and flows outwardly into the power stream between the turbine rotor and the wall 84.

The stream of cooling air for cooling the interstage faces of the annular flange portions 62 is likewise withdrawn from the main air stream where it is substantially at its maximum pressure. Thus, air from the main air stream enters the annular passage A l adjacent the combustion chamber members 42 and flows downstream to the space between the nozzle supporting ring 53 and the turbine casing member 54. At the downstream end of this space, there is a brace in the form of a ring 86 which, at its inner edge, supports the turbine shroud 65. The ring 86 is provided with a series of perforations 37 to permit the cooling air to flow therethrough, and such air continues to flow downstream in the space between the shroud 65 and the turbine casing member 56. The outer ends of the nozzle blades 52 are closed so that such air does not enter these blades.

As this cooling stream moves downstream between the shroud 65 and the turbine casing member 54, it passes around the outer end portions of the interstage guide vanes 66 and enters the space between the ring it and the outer shell 55. The cooling stream then reverses its flow to pass between the turbine casing member 54 and the outer shell 55 so that it may enter the outer ends of the hollow interstage guide vanes 66. The cooling air thus flows inwardly through these interstage guide vanes to maintain them The walls 83 and St thus 6 at a desirably low operating temperature. At the inner ends of the vanes 66, the cooling air enters the hollow ring structurev ll supporting the vanes 66. At its inner periphery, the hollow ring structure is provided with two series of apertures opening toward the adjacent faces of the respective flange portions 62 of the rotor 60, and the sure or" these streams, however, is .suiiicient to' cause their discharge into the power stream after effecting the desired cooling.

The third stream of cooling air utilized in the present construction is arranged to cool the interior of the rotorttl as well as the two stages of turbine blades 63 and (it. However, as this cooling air passes through the turbine rotor and outwardly through the turbine blades, a pumping action is exerted thereon which increases the pressure thereof. Consequently, the pressure at which this stream of cooling air is received from the air compressor may be lower than the maximum thereof to substantially the extent that the pumping action of the .rotor increases the pressure. In the present instance, it is found that the pressure of the main air stream at a point between the compressor H and the compressor it is sufficient, with the pumping action of the rotor, to effect the discharge of the air into the power stream.

To provide for this last-mentioned stream, the outer wall 2! of the intermediate compressor casing is (see Figs. 1 and 3) is provided with a plurality of apertures ti opening into the passages 26 formed by the bul es 25 in the intermcdiate compressor casing. At the downstream end of the are a series. of openings at (see Figs. l-A and 3) in the cylindrical portion 32 of the outer shell iii of the collector E5. The flanges and 35 of the collector E5 in line with the bulges 25 are provided with sheet-metal covers or plates forming, with the outer shell 35 of the collector l5, continuations of the passages 26% so that the air taken from the passage iii in the intermediate compressor casing 26 thus will flow through the spaces between the covers 93 and the collector 65. The flanges 35 of the collector are similarly provided with a series of apertures he (see Figs. l-A and 4) to discharge this air into the passage 458 formed by the outer casing id and the outer shell 55 of the combustion chamber means.

At the downstream end of the passage 45 is a ring 95 which prevents the air from the passage ts from entering the space between the turbine casing member Si l and the outer'shell 55. The nozzle blades or vanes 52, however, are provided with expansion supports 98 (see Fig. 1-B) which are hollow, so that the air from the passage Mi may enter the interiors of the nozzle vanes 52. Cooling air is thus introduced into the nozzle vanes to maintain them at a sufficiently low operating temperature. It is to be noted that the expansion supports 95 are located adjacent the front edge of the nozzle blades 52, so that the maximum cooling effect 7 on the nozzle blades is produced at the front edge thereof, where the greatest amount of heat is received from the power stream emerging from the combustion chamber means.

The cooling air flowing inwardly through the nozzle vanes 52 enters the annular chamber formed by the nozzle ring 5I and the casing walls 83 and 84. The casing wall 83 is provided with a plurality of apertures 91 which permit this cooling air to enter the space immediately in front of the flange El which is formed on the downstream end of the main drive shaft I2 and supports the turbine rotor 60. From the space in front of the flange 6|, the air is permitted to flow through a series of apertures I provided in the flange GI to enter a central bore IDI in the turbine rotor I00. The downstream end of this bore is closed by a flange member I02 abutting the turbine rotor 60 and carried on an extension I03 secured to the main drive shaft I2. The extension I03 is substantially smaller than the diameter of the central bore IUI so that an annular space is thus provided within the turbine rotor to receive the cooling air entering through the apertures I00.

The interior of the turbine rotor 60 is provided with a pair of annular grooves IM extending from the central bore HM, and from the grooves I04 extend in each of the flange portions 62 a series of radial passages Hi5. There are the same number of passages I05 as there are blades 63 and 64 and the passages are aligned with such blades and communicate with the interiors of the blades so that air from the central bore IDI through the radial passages I04 and the blades 63 and 64. The air passing ontwardly in this manner is discharged into the power stream at the outer ends of the turbine blades.

The rotation of the turbine rotor 60 with the blades 53 and 64 produces a pumping action on the air in the passages I65 and the blades, which increases the pressure of the air. Such pressure, when added to the pressure at which this stream of cooling air is withdrawn from the main air stream intermediate the two compressors, is sufficient to produce a total pressure at the outer ends of the turbine blades to cause the air to be discharged into the power stream. Because of the pumping action of the turbine rotor on this stream of air, the air is withdrawn from the power stream at a point below maximum pressure, and less energy is expended thereon than if it were withdrawn from the main air stream at a point of maximum pressure. The total pressure of the air at its point of discharge, however, is sufficient to effect its entrance into the power stream.

I claim:

1. An internal combustion gas turbine engine having at least two stages of air compressing means associated therewith comprising: a forwardly sloping generally conically shaped front housing; a first stage axial flow air compressor at the forward end of said housing; a second stage axial and radial flow compressor to the rear of said first stage compressor; an annular passage within said housing to conduct the partially compressed air from the first stage compressor to said second stage compressor; a compressed air collector ring at the discharge end of said second stage compressor, said ring forming a compressed air passage; an outer annular passage Within said front housing, said outer annular passage surrounding said first named will be discharged outwardly annular passage as well as said second stage compressor and said air collector ring; the common wall between said two annular passages having ports therein to permit a portion of the partially compressed air to flow into the outer annular passage; a rearwardly sloping generally conically shaped rear housing; said rear housing having an annular air passage inside thereof that is substantially co-extensive with its inner wall and which defines a space within the rear housing; said outer annular passage in the front housing communicating with said co-extensive passage in the rear housing; a combustion chamber disposed rearwardly from said collector ring and within said space; a hollow bladed gas driven turbine having a rotor disposed within the rear portion of said rear housing; means to feed fuel to said combustion chamber; said collector ring feeding compressed air to the space inside of said co-extensive passage; generally concentrically arranged annular air flow passages within said space and surrounding said combustion chamber; concentric passages having infeed openings at their forward ends whereby separate portions of the compressed air issuing from said ring enters each of said concentric passages; auxiliary passage means cooperating with said concentric passages to deliver air from adjacent the rear ends thereof to cool the outside surfaces of said turbine blades; and other auxiliary passage means cooperating with said co-extensive passage to deliver air therefrom to the center of said turbine rotor so that the air may be pumped outwardly through the hollow blades of the turbine by their rotation.

2. An internal combustion gas turbine engine having at least two stages of air compressing means associated therewith comprising: a forwardly sloping generally conically shaped front housing; a first stage axial flow air compressor at the forward end of said housing; a second stage axial and radial flow compressor to the rear of said first stage compressor; an annular passage within said housing to conduct the partially compressed air from the first stage compressor to said second stage compressor; a compressed air collector ring at the discharge end of said second stage compressor, said ring forming a compressed air passage; an outer annular passage within said front housing, said outer annular passage surrounding said first named annular passage as well as said second stage compressor and said air collector ring; the common wall between said two annular passages having ports therein to permit a portion of the partially compressed air to flow into the outer annular passage; a rearwardly sloping generally conically shaped rear housing; said rear housing having an annular air passage inside thereof that is substantially co-extensive with its inner wall and which defines a space within the rear housing; said outer annular passage in the front housing communicating with said co-extensive passage in the rear housing; a combustion chamber disposed rearwardly from said collector ring and within said space; a two stage hollow bladed gas driven turbine having a rotor disposed within the rear portion of said rear housing; means to feed fuel to said combustion chamber; said collector ring feeding compressed air to the space inside of said co-extensive passage; generally concentrically arranged annular air flow passages within said space and surrounding said combustion chamber; the inner of said concentric passages having infeed openings for compressed air at its forward end; auxiliary passage means cooperating with said inner concentric passage to deliver air from adjacent the rear end thereof to flow over the forward edges of the first stage turbine blades and then rearwardly to the exhaust of the engine; the other of said concentric passages having infeed openings for compressed air at its forward end; auxiliary passages cooperating with said other concentric passage to deliver air from adjacent the rear end thereof to be directed outwardly into the gas stream of the turbine from near the center of said rotor and along the rear edges of the blades of the first stage of the turbine and the front edges of the second stage of the turbine; and other auxiliary passage means cooperating with said 00- extensive passage to deliver air therefrom to the center of said turbine rotor so that the air may be pumped outwardly through the hollow blades of the turbine by their rotation.

GEORGE] F. WISLICENUS.

References Cited in the file of this patent UNITED STATES PATENTS OTHER REFERENCES Design Analysis of Messerschmitt MELZ'SZ Jet Fighter, in Aviation, November 1945, pages 129 and 130.

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US2760338A (en) * 1952-02-02 1956-08-28 A V Roe Canada Ltd Annular combustion chamber for gas turbine engine
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US2812158A (en) * 1951-12-06 1957-11-05 United Aircraft Corp Stator ring construction
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US4668162A (en) * 1985-09-16 1987-05-26 Solar Turbines Incorporated Changeable cooling control system for a turbine shroud and rotor
US5094069A (en) * 1989-06-10 1992-03-10 Mtu Motoren Und Turbinen Union Muenchen Gmbh Gas turbine engine having a mixed flow compressor
US6488469B1 (en) * 2000-10-06 2002-12-03 Pratt & Whitney Canada Corp. Mixed flow and centrifugal compressor for gas turbine engine
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EP1167722A3 (en) * 2000-06-20 2003-05-28 General Electric Company Methods and apparatus for delivering cooling air within gas turbines
US20050002781A1 (en) * 2002-12-03 2005-01-06 Rolls-Royce Plc Compressor for a gas turbine engine
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US2812158A (en) * 1951-12-06 1957-11-05 United Aircraft Corp Stator ring construction
US2760338A (en) * 1952-02-02 1956-08-28 A V Roe Canada Ltd Annular combustion chamber for gas turbine engine
US2793832A (en) * 1952-04-30 1957-05-28 Gen Motors Corp Means for cooling stator vane assemblies
US2912222A (en) * 1952-08-02 1959-11-10 Gen Electric Turbomachine blading and method of manufacture thereof
US2942844A (en) * 1952-12-22 1960-06-28 Gen Motors Corp Turbine nozzle
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US4503668A (en) * 1983-04-12 1985-03-12 The United States Of America As Represented By The Secretary Of The Air Force Strutless diffuser for gas turbine engine
DE3514352A1 (en) * 1985-04-20 1986-10-23 Mtu Muenchen Gmbh Gas turbine engine with devices for diversing compressor air for cooling hot parts
US4761947A (en) * 1985-04-20 1988-08-09 Mtu Motoren- Und Turbinen- Union Munchen Gmbh Gas turbine propulsion unit with devices for branching off compressor air for cooling of hot parts
US4668162A (en) * 1985-09-16 1987-05-26 Solar Turbines Incorporated Changeable cooling control system for a turbine shroud and rotor
US5094069A (en) * 1989-06-10 1992-03-10 Mtu Motoren Und Turbinen Union Muenchen Gmbh Gas turbine engine having a mixed flow compressor
EP1167722A3 (en) * 2000-06-20 2003-05-28 General Electric Company Methods and apparatus for delivering cooling air within gas turbines
US6488469B1 (en) * 2000-10-06 2002-12-03 Pratt & Whitney Canada Corp. Mixed flow and centrifugal compressor for gas turbine engine
US6536201B2 (en) * 2000-12-11 2003-03-25 Pratt & Whitney Canada Corp. Combustor turbine successive dual cooling
US20050002781A1 (en) * 2002-12-03 2005-01-06 Rolls-Royce Plc Compressor for a gas turbine engine
US7437877B2 (en) * 2004-02-13 2008-10-21 Honda Motor Co., Ltd. Compressor having low-pressure and high-pressure compressor operating at optimum ratio between pressure ratios thereof and gas turbine engine adopting the same
US20050178105A1 (en) * 2004-02-13 2005-08-18 Honda Motor Co., Ltd. Compressor and gas turbine engine
EP2679783A3 (en) * 2012-06-29 2018-03-07 Rolls-Royce plc Spool for turbo machinery
GB2503495A (en) * 2012-06-29 2014-01-01 Rolls Royce Plc Spool for turbo machinery
GB2503495B (en) * 2012-06-29 2014-12-03 Rolls Royce Plc Spool for turbo machinery
US9366260B2 (en) 2012-06-29 2016-06-14 Rolls-Royce Plc Spool for turbo machinery

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