US2810544A

US2810544A - Gas turbine rotor

Info

Publication number: US2810544A
Application number: US237229A
Authority: US
Inventors: Schorner Christian
Original assignee: MAN Maschinenfabrik Augsburg Nuernberg AG
Current assignee: MAN AG
Priority date: 1951-01-20
Filing date: 1951-07-17
Publication date: 1957-10-22
Anticipated expiration: 1974-10-22
Also published as: DE943328C; CH295196A

Description

Oct. 22, 1957 c. scH6RNER 2,81

cns TURBINE ROTOR Filed July 1'1. 195:

CHRISTIAN ScHb'RNEIa BY I Q '5 Fig.6. WMQQIM Z0 l I ATTORNEY-5 J5 I mvem'oa United rates Patent GAS TURBINE ROTOR Christian Schiirner, Augsburg, Germany, assiguor to Maschineufabrik Augsburg-Numberg A. G., Augsburg, Germany, a corporation of Germany Application July 17, 1951, Serial No. 237,229

Claims priority, application Germany January 20, 1951 4 Claims. (Cl. Z53-39.15)

This invention relates to certain improvements in regard to the thermal protection of steel rotors in gas turbines with high temperatures of the propellant gas.

It is an object of the present invention to provide means for protecting the blade-carrying elements, i. e. the steel discs or drums of an axial-flow oneor multi-stage turbine rotor against the very high temperatures of the propellant which necessitate the provision of moving blades of special ceramic materials or other materials adapted for such high temperatures. The provision of ceramic blades is urged by the well-known'inferiority of metallic materials as regards creep strength, and also by the tendency of reducing to a minimum the loss of heat energy (e. g. by conduction into the metal rotor) which is suitable for the production of mechanical energy.

In view of the exacting conditions of operation existing in the hot rim of the turbine rotor rotating at a high speed it is very diflicult to provide effective and reliable heat barrier zones on the surface thereof.

According to the present invention this task is solved in such a way that more particularly the superficies or surface portions of the rotor in front of and behind the blade-grids or stages are interrupted, by cutting grooves of a suitable cross section and depth, into comb-like strips, extending in an axial and/ or in a peripheral direction, depending on the local conditions such as direction of gas flow. These strips are shaped so that they are still in a reliable mechanical coherence with the rotor bodies. The upper ends of the walls between the grooves are then turned over, as by rolling, whereby cavities with practically stationary gas layers are obtained. These gas layers together with the reduced cross sections for thermal conduction to the solid disc body result in an efiective reduction of the rim temperatures, from the value determined by the driving gas, to sufiiciently lower values in the portion carrying the blades, such lower values being required to ensure a satisfactory creep strength of the steel. Thus the thermal tensions in the blade-carrying part of the rotor, especially when starting and on rapid'changes of the load, also become smaller. It is also possible, by the construction according to the present invention, to increase the wall thickness of the drum and thus to increase the load-carrying capacity of the rotor. By grinding the outer diameter of the rolled-down intermediate walls it is easily possible to compensate for small plays of the guide blades with respect to the rotor thus protected. It is also possible, by known means, to feed an air current through the cavities obtained by rolling over the intermediate walls, whereby the insulation is increased in known manner. By the formation of air screens issuing from the cavities in a radial outward direction through gaps in the rolled over walls, the transfer of heat from the passing hot propellant gases to the rotor can be substantially reduced.

According to a further modification of the invention .a

Patented Oct. 22, 1957 transmission of heat from the heated outer surfaces of the rotor to the supporting solid core thereof are reduced, by transverse and crosswise grooves, to cross sections in the form of thin studs. The possibility of a flow of gas between the studs, which flow would impair the intended insulating effect, is practicallyeliminated by coating the studs with a ceramic or enamel layer by which the free space between the studs is greatly narrowed down. Owing to the spiral movement of the heating gas envelope around the rotor and by a corresponding direction of the grooves the formation of a convection current of the propellant gas within the narrow gaps left is avoided and thus a substantial reduction of the insulating effect of the layer is avoided. Owing to the fact that the studs are substantially unaffected by the mechanical strains within the carrier disc or drum a reliable durability of the heat insulating layer can be ensured if the expansion characteristics of the coating layer are approximately adapted to the data of the steel material.

According to a further feature of the invention it is also contemplated to fill up the entire interspace between the studs with a self-hardening heat-resistant cement of poor thermal conductivity and suitable characteristics of heat elongation, whereby the heat-insulating layer at the surface of the rotor will be nearly perfect.

Other and further objects, features and advantages of the invention will be pointed out hereinafter and appear in the appended claims forming part of the application.

In the accompanying drawings several now preferred embodiments of the invention are shown by way of illustration and not by way of limitation.

Fig. 1 is a fragmentary axial section through a turbine rotor,

Fig. 2 is an axial section through the heat-insulating layer 3 of a rotor, with the intermediate walls partly bent over,

Fig. 3 is an axial section through a modified form of heat insulating layer, with the intermediate walls partly bent over,

Fig. 4 is a plan view of a part of the rotor rim with longitudinal and transverse grooves,

Fig. 5 is an axial or transverse section through one row of studs 13 of Fig. 4 showing a heat insulating layer having longitudinal and transverse grooves, and

Fig. 6 is a fragmentary plan view of the surface of the rotor of Fig. 1.

in known manner including, for example, the spacing inserts 17. The heat insulating layers at the superficies of the disc or drum before and behind the blade grids or stages are denoted 3, and are illustrated as the form shown in'Fig. 2. For some high temperature operating conditions requiring additional heat insulation of the rotor, an axial boring 18 is provided in the rotor from which radial passages 19 branch off and open into an axial groove 20 at the rotor surface covered by cover plate 21. The circumferential grooves or spaces 6 open into the axial groove 20 so that a coolant may be circulated through the groove 6 by being introduced into axial boring 18 and, thence, through radial channels 19 to groove 20. According to Fig. 2 these insulating layers are formed by annular grooves 4 worked into the surface, the outer portions of the intermediate walls 5 between these grooves being bent over to form hollow spaces 6 containing either a stagnant gas layer or being traversed by cooling air.

In the embodiment shown in Fig. 3 the intermediate walls 7 are provided with two brims or

tabs

8 and 9 which are bent over to opposite sides and together with thetabs 8 and'9 of 'tlieadjacent walls form the upper closure ofthe hollow spaces 10. Depending onthedirection of the gas flow throughlhe turbine, the grooves and the

hollow spaces

6, or 10, extend transveresly to the longitudinal axis of 'the "rotor or "parallel thereto and across-the direction of-gasfiow.

In the modification of the heat-insulating layer asper' Figs. 4 and 5 transverse and longitudinal 'grooves 11' and 12 are provided,subdividing the insulating layer'into single studs 13 which are provided with a heat=resistant coating 14 of ceramic material, enamel or the like, whereby the grooves 11 and 1'2"are11arrowed down "to small gaps, Fig. '-5. If desiredit'is'a1so-possibleto fill up the grooves l-l-and- '12-'entirely with aheat-Tesistant'insulating-layer.

The production 'of the"heat'-insuiating' layer according tothe invention is very "simple-and'oiferstheadvantage that loose screening elements are avoided whose. reliable fastening on the rotor is oftentimes very'difiicult'owing' to the high speeds.

While'the invention has *beemdescribedin fietai'l'iwith respect to certain 'now preferred examples *and'enibodiments of the invention it will *be'understoodtby those skilled in the art after understanding the invention that various changes" and' modificationsrnay *be'madewithout departing from-the spirit and-scope ofthe invention and it -'is intended, therefore; *to'cover all'such changes and modificationsin the appended'elairns.

What is claimed-is:

1. Ida high temperature 'gasturbine :having a metal rotor with a plurality of "stages" er turbine blades thereon and in which the-surface of'saidrotor between 'said"adjacent stages 'is directly exposed to-"high temperature gas flow through said'turbine, a heat insulatirrg'layer "forsaid exposed surface of said'rotorbetweensaid adjacentst'ages' comprising a plurality of-thinprojections "at said exposed rotor surface defining therebetween aplurality of .circumterential annular grooves forming dead gas 'spaces', said projections being relative1y1hin' for reducing heat conduction from said high temperature gas .flow to said rotor and the cross-section'of said spaces "between said jprojections being relativelysrnall' avoiding convection currents in said spaces, and means for at leastpartially closIng sai'd' spaces at the'radially'outer portion thereof for reducing aerodynamic flow resistauce'ofsai'd exposed rotor surface.

2. In 'ahightemperature gas'turbine having. ametal rotor with a pluralityof stages of turbinebla'des thereon and in whichthe surface-ofsaid rotor'between said adjacent stages is directly'exposed to "high temperature gas flow through said turbine, aheat insulatinglayer for. said exposedsufface ofsaidrotorbetween said adjacent stages comprising a plurality ofthinprojections on said exposedrotor surface definingitherebetween a plurality of circumferentialannular gro'ovesforming insulating gas spaces for reducing heat conductionfrom. said high temperature gas flow to said-rotor, the radially outer portionsof said" projections forming'atrleast'a partial closure "for said spaces to reduce aerodynamic flow resistance thereof,

and the radially inner portions of said spaces defining a heat insulating layer interposed "between "said 'high "temperature gas flow and said rotor.

3. In a high temperature gas turbine having a metal rotor with a plurality of stages of turbine blades thereon and in which the surface of said rotor between said adjacent stages is directly exposedto high temperature gas flow through said turbineyasheat insulating layer for said exposed surface of said rotor between said adjacent stages comprising a plurality of thin projections on said exposed rotor surface-definingitherebetweena plurality :of circumferential annular grooves forming insulating gas spaces .for reducing'heat conduction-from saidihigh temperature gas flow to said rotor, the radially outer portions of said 1 projections forming at least "a partial closure at said spaces to reduceaerodynamic ilowrresistance thereof, the radially inner portions of said spaces defining a heat insulating layer interposed between said high temperature gasflow -and.said rotor, and means .forsupplying a heat insulating coolant into said spaces.

4. 'In a hightemperature gasturbine havinga metal rotor with. a plurality of stages. of turbine blades thereon andflin which thesur'faceof said rotor-between said adjacent stages is directly exposed to high temperature gas flowihrough :said turbine, .a. heat insulatinglayer for said exposed surfaceofisaid rotorbetweensaid adjacent stages comprising a plurality ofithinprojectionson said exposed rotor surface defining therebetweena plurality of circumferential annular, grooves .iorming insulating .gas spaces for reducing heat conductionfrom.saidhightemperature gas "flow tosaid rotor, the radially .outer portions of said projections forming at least .a partial .closurefor said spaces to reduce aerodynamicflowrresistance thereof, the radially innerportions .of saidflspaces defining a heat insulatinglayerinterposed'vbetweensaid high temperature gasfiowfand sai d rotor,,and passagemeans through said partial closure communicating between said spaces and the. outer surface of said .closurejfor passage of low l temperature gasiromrsaidspaces to "form. a heat insulating layer over the outer surface .otseiid rotor.

References (iitedrinthefile of this patent UNITED STATES -.EA'1 ENTS 102,46 1v OTHER 15 Serial No; 385,333, Schutte (Am B. 0.), published May 25', .1943.