US2608821A - Contrarotating turbojet engine having independent bearing supports for each turbocompressor - Google Patents

Description

Sept. 2, 1952 E, HUNSAKER 2,608,821

CONTRAROTATING TURBOJET ENGINE HAVING INDEPENDENT BERNG SUPRORTS FOR EACH TURBOCOMPRESSOR Filed OC. 8, 1949 Patented Sept. 2, 1952 UNITED f STATE v 2,608,821,l Y

coNTnAnoTArING TUn'BolJE'I nncmirf` 1. Y HAVING INDEPENDENT BEARING sur roars ron EACH TUaBocoMrREsson vEugene L. Hunsaker, Manhattan Beach, Calif., assigner, by mesne assignments, to General," Electric Company, Schenectady, N. Y., a corporation of New York Application october s, 194e, sentirne. 120,247,

1 Claim. 1

My invention relates to axial flow compressorturbines of the gaseous combustion type, and more particularly to an improved compressor.- turbine suitable for use in aircraft driven at high subsonic or supersonic speeds.

When combustionV turbines are utilized in such aircraft, high airow capacity is necessary to increase the engine thrust coecient and to de,- crease the external drag coelicient. At present, the required airflows are obtained by the use of high blade speeds in axial new compressors of compressor-turbine combinations used to drive the high speed airplanes.

High blade speeds `Vresult in large values of centrifugal stress on the. rotating blades as Stress (centrifugal):

l( blade tip speed)2 X blade length constant coinpressordiarneter This stress relationship in the compressor limits the blade length and hence limits flow area and capacity fora given diameter of compressor. A directly coupled turbine Will also have large values of blade tip speed, and hence high values of centrifugal stress. The physical characteristics or" the hot'turbine bladesparticularly limits blade length, and hence airflow rate ior a. given turbine diameter.

It is an object of the present invention to provide a compressor-turbine combination which is capable of higher air capacities than are conventionally obtainable with compressors. using high blade speeds. and supersonic air velocities relative to the blades.Y y

It is another object of thev present invention to providean improved compressor-turbine combination having relatively low blade tip speeds.

It is still another objectof the invention to provide a compressor-turbine..combinationjhaving relatively long blade lengths, with aconsequent signiicant increase in airow capacity.

It is a further objectoi thepresent invention to provide a compressor-turbine,- combination of the combustion type suitable for driving aircraft at high subsonic and supersonic speeds.

A still further object of the invention is to provide an improved compressor-turbine in which high compression ratios can be produced in a single compressor stage, while maintaining high air capacity and relatively low blade tip speeds.

Another object of the present invention is to provide higher compressor eiciencies in compressor-turbines operating at supersonic relative velocities.

Briefly, the present invention involves a com- (Cl. (iO-35.6).

bustionturbine 'engineVv -having contra-rotating compressor blade..discs..directly ,driven by contrarotating turbine blade ,discsf The terms ycompressor rotor blades and turbine rotor blades are used .in the conventional sense of meaning bladeshaving the.. proper. contours Vfor rotating blades of. compressorsandturbines of compressor-turbine engines;-- whereastheterms compres-` sor.- counter-rotating,blades nd turbine counterrotating` blades are usedixr an unconventional sense indicating only thatthe counter-rotating blades have respective proper contours. .for v compressor stator blades and! turbine stator blades.

even though rotating. i Y

My present.v invention provides for ay compressor-turbine in which thecompressorjbladesfof, the

stator typev are; contrafrotait-xi with respect to the compresser-rcter` bladesjso.- that.Y the compressor-rotor speed` canbey decreased for: a' given overall stage pressure ratio. I prefer to drive the; contra-rotating compressor blades by a turbine-v disc having rturbinerotor bl'ading thereon, working in conjunction-with a. contra-rotating disc having turbineystatorftypeblades thereon, this disc beingconnectedfxto drivethe compressor-rotorblading. "The compressor rotor blading is preferably directly connected' to .a turbine. discY having turbine.' stator. type bladesthereion, so that both compressor: and'. turbine comprise, contrarotat-ing blading'. .In;this wiay thefrtip speeds of both compressor anduturbine .bladin'g are relatively low., and: anv increased. air: .capacity is obtained. forY a. givenv engine'. diameter.

lnzthisfdesigrr, thefstator .type blades are contra-rotated, with respect' to'thezrotor type blades and the ,bla-1deboundary. layer energizedV through this relative rotation.: suppresses separation of' thef fiowfrom. thefstator; type; blades and:

improves efficiency.

My'invention. will be;4 morcef: fully understood by reference: to the. drawing-3. showing in` diagrammatic for-1n,Y a longitudinal. sectional view of a combustion turbo-jet engine embodying a preierred'f form: ofv the: present. invention.

Referring; to'A the.drawing,.an= outer turbine casing. IV is-provided having Yat one end an airflowv inlet 2 and at the opposite end, a jet outlet V3. Attached to casing I just inside of the airflow inlet, a plurality of radial guide vanes 4 are provided, extending inwardly to support an inlet cone 6, which is hollow, and in turn supports a plurality of inlet bearing hangers 1 supporting an outer race 8 of a forward bearing 9.

Inside of the jet outlet 3 a plurality of radial tail cone hangers l0 are provided, supporting a tail cone II. Inside of tail cone II a plurality of outlet bearing hangers I2 are provided, supporting the outer race I4 of a rear bearing I5. Journalled in bearings 9 and I5 is an inner cylindrical rotor i6. The forward endof theV inner rotor I6 is provided with a compressor rotor disc I'I on which are mounted a plurality of compressor rotor blades I8 just rearwardly of guide vanes 4. Attached to the rear of inner rotor I6 is a turbine counter-rotating disc lI9 provided with blades 2D of turbine stator type, these blades being the rearmost of the turbine assembly.

Between the compressor rotor `blades I8 and the turbine stator type blades 20, vinside of casing I and adjacent the inner surface thereof, is positioned a, ring of combustion chambers 2| of any conventional type, each supplied with fuel through a, fuel inlet jet 22. y

Forwardly of the combustion chamber ring, attached to casing I and extending radially inwardly is a ring of stationary compressor outlet guide vanes 2 4 supporting an inner shroud 25, forming an annular space for combustion chambers 2I. Y

Rearwardly of the combustion chambers 2l, attached to casing l and extending radially inwardly is a ring of stationary turbine inlet nozzles 26 also supporting the Ainner shroud 25.V Extending inwardly from kshroud 25 at the guide.

vane and nozzle stations, are a plurality of radially arranged struts 21 and 28 supporting an outer rotor 30 rotating coaxially with inner rotor I6 on front and rear outer rotor bearings 32 and 34 respectively. ,l l Y Compressor Y rotor blades I8'T'an'd' stationary guide vanes 24 are spaced to permit insertion therebetween of a plurality of compressor stator type blades 35 attached tothe forward end of outer rotor 3U, by compressor counter-rotating disc 36.

Similarly at the turbine, nozzles 26 and Yturbine stator type blades 20 are spaced to provide for a plurality of turbine rotor blades 37 mounted on a rear rotor disc 38 of youter rotor 30, Thus both the compressor and rotor have two sets of rotating blades. As one set of blades of both compressor and turbine are shaped as rotor type blades, and since another set of blades in each unit are shaped as stator type blades, the inner and outer rotors will be contra-rotating, with both rotor and stator type'blades of the compressor beingv driven bythe turbine blades, but in opposite directions. i It will be apparent to one skilled in the art that, for a given overall stage pressure ratio, the rotor speed is decreased by the use of such an arrangement. lThe decreased speed permits greatercompressor and turbine blade length, thereby increasing air capacity for a given engine diameter.

While in order to comply with the statute, the invention has been described in language more or less specific as to structural features, it is to be understood that the invention is not limited to the specic features shown. but that the means and 'construction herein disclosed comprise a preferred form oi' putting the invention into eiect, and the invention is, therefore, claimed in any of its forms or modifications within the legitimate and valid scope of the appended claim.

What is claimed is:

A gas turbine combination comprising an outer casing having an air entrance aperture and a jet aperture, a nose cone in said entrance aperture, a plurality of guide vanes extending from said casing and supporting said nose cone, a forward bearing supported by and within said nose cone, a tail cone in said jet aperture, a plurality of struts extending from said casing and supporting said tail cone, a rear bearing supported by and within said tail cone, an inner rotor journalled in said forward and rear bearings, a first compressor disc on the forward end of said inner rotor, a plurality of blades of compressor rotor type on said first compressor disc, a second turbine disc on the rear end of said inner rotor, a plurality of blades of the turbine stator type on said second turbine disc, a shroud positioned in said casing between said discs, supports extending from said casing and holding said shroud in said casing to provide an annular space between said shroud and casing, front and rear shroud bearings inside of and supported by said shroud, an outer rotor supported mechanically independent of said inner rotor and mounted in said latter shroud bearings to rotate concentrically with and spaced from said inner rotor, a second compressor disc on the forward end of said outer rotor, a plurality of blades of the compressorstator type on said forward disc, said latter blades being positioned adjacent and to the rear of the blades on said rst compressor disc, a rst turbine disc on the rear end of said outer rotor, a plurality of blades of turbine rotor type on said rst turbine disc, said latter blades being positioned adjacent and forward of the blades on said second turbine disc, and combustion chamber means in the annular space between said shroud and said casing.

EUGENE L. HUNSAKER.

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

UNITED STATES PATENTS Number Name Date 2,360,130 Heppner Oct.` 10, 1944 2,396,911 Anxionnaz Mar. 19, 1946 2,404,767 Heppner July 23, 1946 2,405,723 Way Aug. 13, 1946 2,409,176 Allen Oct. 15, 1946 2,430,399 Heppner Nov. 4, 1947 2,476,179 Cameron July 12, 1949 2,483,401 Cole Oct. 4, 1949 2,505,660 Baumann Apr. 25, 1950 2,563,744 Price Aug. 7, 1951 2,575,682 Price Nov. 20, 1951 Y FOREIGN PATENTS Number Country Date 879,123 France Nov. l0, 1942

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