US2671315A

US2671315A - Internal-combustion geared turbine

Info

Publication number: US2671315A
Application number: US59496A
Authority: US
Inventors: Charles F Rocheville
Original assignee: ROCHEVILLE ENGINEERING Inc
Current assignee: ROCHEVILLE ENGINEERING Inc
Priority date: 1948-11-12
Filing date: 1948-11-12
Publication date: 1954-03-09
Anticipated expiration: 1971-03-09

Description

March 9, 1954 c. F. RocHl-:vlLLE 2,671,315

INTERNAL-COMBUSTION GEARED TURBINE Filed Nov. l2, 1948 3 Sheets-Sheet l IIIIII IN VEN TOR C/)a e: F.l FPocbevi/k ATrR/IE Ys ,E willlllllmm,

March 9, 1954 Q F; ROCHEWLLE 2,671,315

INTERNAL-COMBUSTION GERED TURBINE Filed Nov. l2, 1948 5 Sheets-Sheet 3 IN VEN TOR. Char/es F.' Fac/7e vil/e Patented Mar. 9, 1954 UNITED STATES FATENT OFFlCE INTERNAL-COMBUSTION GEARED TURBINE Application November l2, 1948, Serial No. 59,496

7 Claims.

This invention is directed to internal combustion turbine engines for aircraft, and particularly refers to an improved arrangement in which a plurality of fuel burners and combustion cham bers are disposed in a rotor with their longitudinal axes substantially in the plane of rotation at right angles to the axis of the turbine, and are arcuately positioned in helically overlapping relation to discharge combustion products tangentially to indu-ce rotation by reaction, the rotor being positively geared to counter-rotating varied turbine wheels which receive the combustion products to be rotated thereby, both rotating elements being adapted to drive a propeller shaft. Desirably, but not necessarily, clutch provisions and free-wheeling means for driving air compressing means may be included.

It is an object of this invention to provide a construction of the nature set forth which will permit high relative speed of the combustion gas driven counter-rotating elements, with a substantially lower speed of the propeller drive shaft.

Another object is to provide an improved arrangement of the burner and combustion chambers of a turbine of this type, to give a relatively long path for combustion to take place, in a structure of relatively short axial length.

Another object is to provide an improved arrangement of stationary fuel-air nozzles around the periphery of the turbine, said nozzles feeding continuously into a heat exchange passage and thence into a perforated burner tube to initiate combustion, which is subsequently substantially completed in a combustion chamber.

Another object is to provide an improved ar# rangement of gearing and lubricating means for a turbine unit of this type.

Another object is to provide an improved control means for the air compressing means of a turbine of this type, that can be selectively connected to furnish air for combustion at air speeds below that which ram-jet action occurs, and may be disconnected from the power unit at speeds where air compression is not required to `consume the fuel at the desired rate.

These and other objects and advantages of the invention will 'be further apparent from the following description and the attached drawings, which form a part of this specification, and illustrate a preferred embodiment as applied to an aircraft engine.

(Cl. Gli-39.35)

In the drawings, Figure 1 is a longitudinal part sectional view of the upper half of the forward or compressor section of the engine of this example, and illustrates a preferred arrangement of starter, clutch and free-wheeling means. Figure 2 is a rearward continuation view of the central part of the arrangement of Figure l, also in longitudinal and substantially complete cross section, illustrating the gearing, combustion chambers and turbine elements. Figure 3 is fa diagrammatic and somewhat simplified trans-- verse sectional view taken on line III-II of Figure 2, looking forward and taken substantially in the plane of the longitudinal axis 'of one of the two combustion chambers, showing the heli'- cally overlapping arrangement of the said ychl'ain'- bers.

Referring to Figure l, reference numeral i0 designates -a generally cylindrical metal outer casing for the engine unit, which 'may be faired into the wing structure ii of an airplane. An inner metal housing l2 is "suitably spaced and supported from the outer casing' and forms an annular passage I3, open at its forward end 'to admit cooling air for the housing and some engine parts to be described below. Extending inwardly from inner housing l2 are a plurality of struts l5 to support bearing I6 for the fori ward end of propeller shaft Il, the latter desira-v 'bly being hollow as at I8 to admit cooling air. A conventional propeller spinner or hub 'I9 supporting propeller blades 20 is secured 'to the outer end of shaft Il.

A generally conical air guide or ram-jet cone 2i extends from the forward end of bearing iB to a set of stationary guide vanes 22 at the inlet of a multi-stage air `compressor formed by a plurality of rows of inwardly directed stationary vanes 23 and outwardly directed rotating vanes 2li, the latter supported on compressor rotor 25 turning on bearings 26 and adapted to be selectively driven from shaft ll through clutch 28 and free-wheeling means 29 which may be a conventional arrangement of spring-controlled balls on inclined wedges, so arranged that when more power is required than for idling or gliding, the increased speed of rotation of Vthe 'gas actuated power elements to be described below will engage the free-wheeling means and 4drive the compres'i scr rotor to give the desired air now to the com-I bustion chamber. Upon reaching air speeds in the neighborhood of 400 M. P. H., the ram-jet action of the air intake between i2 and 2| will be adequate for air supply, whereupon clutch 28 may be disengaged, leaving the air compressor rotor to idle or windmill. In this example a starting motor 29| is provided to drive the propeller shaft I1 through the medium of gears 30 and 3I. 'I'he clutch control means generally designated 32 may be of the conventional hydraulic type.

Referring to the main section of the turbine engine, and particularly to Figure 2, there is illustrated the after end of compressor rotor 25, supported at this point on shaft I1 by bearing 26. A stationary spider 33 is secured within housing I2 and supports a plurality of circumferentially spaced air-fuel mixing nozzles 34, supplied with liquid fuel such as gasoline, kerosene or fuel oil through a conduit or passage 35, which may surround housing I2, and be connected to any desired type of pressure feed system, not shown. The forward face of spider 33 and the after face of compressor rotor are desirably provided with sealing means, such as the labyrinth of interengaging grooves and landsI 3B, to prevent undue leakage of fluids at this point.

Stationary spider 33 also serves to support bearings 31 for a plurality of circumferentially spaced hollow shafts 38, each shaft supporting at its forward end a first planet geai` 33, and supporting at its after end a second planet gear 4B. At the rear end of propeller shaft i1 is keyed a sun gear 4I, adapted to engage and be driven by the forward planet gear 39. On the rearinost portion of shaft I1 is supported a bearing t2 for the forward extension of a second sun gear 43, the latter also being keyed to the forward end of a hollow turbine shaft 44, which extends rearwardly to a flange 45, onto the rim of which turbine wheels 46 and 41 are secured by a plurality of In this manner, there is provided a turbine shaft 44, supported at its forward end by bearing 42 which is mounted on the rearmost end of propeller shaft I1, through the axially elongated annular sun gear 43, and supported at its after end by bearing 50, through the medium of the turbine wheels 46 and 41 and flanged members 45 and 49. A supplemental bearing for the after end of propeller shaft I1 may be installed at 54, this bearing being supported by a plate secured to the forward face of the stationary spider 33 by through bolts 56, and also acting as a support for the forward ends of shafts 38 through the medium of bearings 51.

A plurality of bearings 58 on turbine shaft 44 serves to support a generally cylindrical hollow combustion chamber rotor 59, which is positioned between the after face of stationary spider 33 and the forward face of turbine wheel 46. An annular flange 60 projects forwardly from the front of rotor 59, and is provided with internal gear teeth to engage the after planet gears 49. This arrangement, whereby sun gear 43 on turbine shaft 44 engages planet gears 49 at their inner pitch circle, and the teeth of annular flange gear 60 engages the same gears 40 at their outer pitch circle, insures that the turbine shaft 44 and the turbine wheels 46 and 41 carried thereby, will rotate in the opposite direction from the combustion chamber rotor 59, for reasons which will b-e further apparent below. Desirably, the forward face of combustion chamber rotor 59 is sealed from stationary spider 33 by means such as the grooved labyrinth 6I, and the after face of rotor 59 is sealed from the turbine wheel 4S by means such as the interengaging grooves and lands 32.

'Ihe forward face of rotor 59 is provided adjacent its periphery with an annular opening 63, extending substantially throughout its circumference, whereby the fuel-air mixture from the nozzles 34 may enter the interior passages G4 and 65 of the rotor. These passages are formed by symmetrical annular metal housings 83 supported within the rotor by ported rings 51, the housings being generally rectangular in crosssection, with an enlarged entrance scoop (Figure 3) and, in addition, similarly ported as at 63 (Figure 2) to permit the fuel-air combustible mixture, after it is heated by the means to be described below, to pass into the spaces-59 within the housings and thence into the circumferentially arranged burner tubes 10 (Figure 3) which lie in helically overlapping relation within the housings 65 and rotor 59. In this example, only two of such housings and tubes are illustrated. and the transverse section of rotor 59 in Figure 3, which illustrates diagrammatically how they are disposed, has been somewhat simplified to facilitate this description. Obviously, if additional burners are desired, they would be symmetrically arranged in similar helically overlapping relation Within the rotor. Desirably, but not necessarily, the inlets and outlets of rotor 59 are at the same distance from the axis of its rotation, thus reducing the diameter of the unit, and facilitating its construction and maintenance.

Referring to Figure 3, it will be noted that the fuel-air mixture from space 69 enters the burner tubes '.'0 through the ports 1I, to be ignited by means such as the spark plug 12, the latter supplied by high voltage electric current by means which will be described below. After the combustion is initiated in burner tube 19, the heated gases pass into the combustion chamber or space designated 13, where expansion takes place to give a high velocity to the gases, which are then directed out of the discharge end of the combustion space by vanes 14 to impinge against the blades 15 of the rst wheel 46 or stage of the turbine element of the unit (Figure 2). After leaving that stage, the gases are straightened by stationary vanes 16, the latter supported by the after extension 11 of housing I2, and are directed into the vanes 18 of the second wheel 41 or stage of the turbine element of the engine unit. In this example, where only two such stages are illustrated, the exhaust gas, from which the greater part of the energy has been abstracted by the turbine vanes, emerges into the exhaust passage 19 formed between the housing extension 's1 and exhaust cone 52, to be released to the atmosphere, where it may impart added energy to the airplane by its reaction as a jet.

The helically overlapping arrangement of the housings 66, burner tubes 10 and combustion chambers 13, with the surrounding

passages

64 and 65, insure a maximum heat transfer' from these highly heated surfaces to the incoming fuelair mixture, thus servingto cool the parts below those temperatures at which damage would occur and also adding heat energy to the combustible mixture and increasing the ,overall efdciency of the turbine engine. From Figure 2 it is. apparent that the last-named mixture divides, as it enters the rotor, part of it going into passages dfi and the remainder of it into passages 65 to. circulate within the rotor as indicated by the arrows and nally to merge again as it passes through ports 68 into spaces 69 inside neusingst. From those spaces the now-heated fuel-air mixture enters theY burner tubes 'i8 of the next circumferentially adjacent housings 66, as best shown schematically in Figure 3. The materials of these partsl are desirably of heat-resistant alloy, such as an alloy of chromium, iron and molybdenum.

After the turbine engine is once in operation, it will not ordinarily be necessary to continue the supply of high tension electric current to spark plugs l2, but foi` starting, and for reignition the current is transmitted from any suitable source through a brush 8d (Figure 2) to an insulated annular metal collector ring 8l on the forward face of combustion chamber rotor 59. An insulated lead 82 extends from ring 8i to the terminal of each spark plug l2.

As stated above, the propeller shaft El is desirably hollow throughout its length. In this example a sleeve 84 is mounted within the turbine shaft flri, and extends within but is spaced from the bore IB of propeller shaft il. rlhis not only serves to conduct cooling air into the turbine shaft, to cool its bearings 553 and 58, but also supplies cooling air to the turbine wheels 46 and 4l, which may be ported as at 85 and 35, respectively, to admit air to the hollow vanes or blades 'i5 and i8, from which it escapes into exhaust passage i9 through notches 8l and 88 in the trailing edges of the blades. Sleeve 315 also serves to conduct lubricant from the gear and bearing space within the stationary spider 33 through the helical passage formed by blades 89 to bearings 58.

In those applications of this invention which are not carried out in a structure, such as an airplane, which moves relatively to the air used for combustion, the clutch means 23 would ordinarily not be required as the compressor rotor 25 would operate at all times that the power shaft Il is rotated. Similarly, circumstances might also be encountered wherein the freewheeling means 29 could be omitted, as will be obvious to those skilled in the art.

In conclusion, it will be appreciated from the foregoing description that this invention comprehends broadly the combination, in an internal combustion turbine engine, of burner tubes and combustion chambers operating by the reaction of gases and discharging into vaned turbine elements, said chambers and elements being positively and compactly geared to rotate in opposite directions and to jointly drive a single shaft at a lower speed than that at which either rotates to deliver useful power, for example, to an airplane propeller. It includes also the circumferential helically overlapping arrangement for the housings, burner tubes and combustion chambers with preheating and heat transfer passages for the fuel-air mixture. Although a single embodiment is described and illustrated as applied to an aircraft engine, it is obvious that numerous modifications and changes could be made in its construction and application to other uses without departing from the essential features of the invention, and all such as fall within the scope of the appended rotating u? claims are understood: to be embraced thereby.

I" claim:

1. In combination in an internal combustion turbine engine, a rotor Supporting a plurality of combustion chambers, means for supplying a combustible mixture to said chambers, at least one turbine wheel forreceiving products of combustion from al1- oi'saidv chambers, a power shaft coaxial; with said rotor, and direction reversing gear mea-ns' separately connecting said shaft to said rotor and said turbine wheel to rotate' them in Vopposite directions to deliver power to said 2. A combination according to claim -1,= in which said combustion chambers are arcuately disposed in helically overlapping relation around said rotor.

3. A combination according to claim 1, with the addition of a rotatable air compressor means surrounding said shaft, and means for selectively connecting said compressor means to said shaft.

4. In combination in an internal combustion turbine engine, a plurality of circumferentially arranged fuel-air nozzles, axial flow means for Ichambers in said rotor, a vaned turbine wheel for continuously receiving heated gases from said combustion chambers, a power shaft, and gear means connecting said rotor and said turbine wheel to rotate in opposite directions, and to rotate said power shaft.

5. A combination according to claim 4, with the addition of a, housing in said rotor for said tubes and chambers forming passages to conduct said combustible mixture in heat exchange relation thereto prior to its introduction into said burner tubes.

6. In combination in an internal combustion turbine engine having an annular air passage and a plurality of fuel-air mixing nozzles at the end of said passage, a rotor adjacent the end of said passage having an annular inlet for combustible mixtures in one end face, a plurality of heat exchange housings in said rotor in helically overlapping relation, a perforated burner tube communicating with and contained in each of said housings, ignition means in each of said tubes, a combustion passage leading from each of said tubes, means for passing said combustible mixture from said annular inlet first through said housings, around said combustion passages to be heated thereby, and then into said burner tubes to be ignited therein, and a combustion gas outlet for each said passage in the opposite end face of said rotor.

'7. A combustible mixture igniting and consuming means for an internal combustion engine having a turbine element to receive the products of combustion from said means, comprising a rotor having opposed axially spaced faces,

an annular passage formed in one of said faces, a plurality of circularly disposed helically overlapping housings in said rotor for continuously receiving a combustible mixture from said passage, a perforated burner tube in each of said housings, an enclosed combustion chamber for each of said burner tubes, an outlet for each of -said combustion chambers, each of said housings adapted continuously to conduct a portion of said combustible mixture in heat exchange relation with its associated combustion chamber and Ydirect said portion of said combustible mixture to a succeeding housing and burner tube to be ignited therein.

CHARLES F. ROCHEVILLE.

References Cited in the le of this patent UNITED STATES PATENTS Number Name Date Clark Mar. 6, 1906 Kothe Oct. 6, 1914 Holtz July 13, 1915 Lake Sept. 17, 1918 Tyler Jan. 14, 1919 Number 10 Number