US2689681A - Reversely rotating screw type multiple impeller compressor - Google Patents

Description

A. SABATIUK TATING SCR Sept. 2l, 1954 2,689,681 Ew TYPE MULTIPLE PREssoR l REVERSELY RO 'Filed Sept.

IMPELLER COM 3 Sheets-Sheet 1 @gew 2 Sept 21. 1954 sABATluK A2,689,681

A. REVERSELY ROTATING SCREW TYPE MULTIPLE i Y IMPELLER COMPRESSOR Flled Sept. 17, 1949 -3 Sheets-Sheet 2 Eil?, 5.

M/L E7' ABSOLUTE COMPU/VENT RELATIVE E/Y/ Hasan/rf 15x/r Q S550/v0 Hora/E eggs-aa? Sept. 21, 1954 A. sABAnUK 2,689,681

REVERSELY ROTATING SCREW TYPE MULTIPLE IMPELLER COMPRESSOR Filed Sept. '17, 1949 3 Sheets-Sheet 3 Eig 5c.

@geni atented S'ept. 2.1., 11954' UNITED STATES REVERSELY ROTATING SCREW TYPE MULTIPLE IMPELLER COMPRESSOR Andrew Sabatiuk, New Britain, Conn., assignor to United Aircraft Corporation, `East Hartford, Conn., a corporation of Delaware Application September 17, 1949, Serial No. 116,286

1 Claim. (Cl. E30- 123) This invention relates to turbo-jet engines and more specically to axial flow compressors for `such power plants which provide high compresling at supersonic speeds relative to the moving rotor blades with the shocks occurring in the passages formed by the blades.

It is another object of this invention to provide an improved compressor for turbo-jet engines` whereby high compression ratios are produced with a minimum of compression stages and a minimum of space.

Another object of this invention resides in the provision of a compressor which exhausts high pressure fluid of such low relative velocity that the diffuser can be completely omitted or reduced to a minimum length.

A further object of this invention is to providea multiple stage axial flow compressor wherein each stage includes a rotor operating on the shock-in-rotor principle.

Another object of this invention resides in the employment of the self stabilizing inlet phenomenon of a supersonic rotor passage thereby eliminating the necessity of statorsbetween rotors in the compressor.

These and other objects will become readily apparent from the following description of the accompanying drawings in which,

Fig. l is a partial cross sectional view of a `turbo-jet engine including the compressor form- @and the rotors shown in Fig. 4;

Figs. 5a, 5b and 5c are partial views of the rotor blades illustratingvarious now conditions through the blades.

Fig. 5d is a detailed view of a rotor blade indicating in detail the iiow through a shock and expansion eld; and

Fig. 6 is a diagrammatic illustration of the compressor pressure ratio and efiiciency `that may be expected of the fluid passing through a shock at various relative inlet Mach numbers.

Although the general practice has been to avoid supersonic relative ilow of compressible uids in compressors due vto-the high loss expec tations, it is known that the supersonic cornlressors are capable of operating at relatively high compression ratios Without exceeding practicable loss limitations. pressors have utilized the principles wherein supersonic relative flo-w is obtained in the compressor and the working uid is passed `through a shock to increase the pressure of the fluid and reduce its velocity below that of sound.

Generally speaking, these supersonic compressors are classied as the shock-in-stator type or the shock-in-rotor type. However, in either ease it has been the practice to utilize stators or turning vanes to obtain proper operation and uid ilow in the rotors. Since the use of stator-s entails a `certain amount of `flow bending, denite losses of various types usually result thereby causing inefciencies which in addition affect the performance of the rotorblades.

To this end, this invention provides a compressor for a turbo-jet engine which compressor in its preferred form operatesy on the shock-in-rotor principle but eliminates the usual losses which accompany the use of stators.

Where compression has been achieved by the shock-in-rotor principle, for example, the usual practice has been to utilize turning vanes in the compressor inlet, upstream of the rotor or rotors, to produce. proper directional l'low and obtain higher relative velocities at the leading edge ci the rotor blades. With this type of construction the velocity of the inlet air ahead of the rotors is adversely limited. In other words, depending upon the shape of the turning vanes and the amount of turning producedy local supersonic velocity will readily occur over the vanes as increased subsonic velocities lare approached thereby causing separation, possible shock and/ which units cooperate to form an air inlet I6V Various types of comand an annular passage I8. The inner housing I4 carries two concentric shafts 20 and 22 which are rotatably mounted therein by means of a plurality of bearings 24. An axial flow cornpressor is provided and includes in its preferred form a pair of adjacent rotors 28 and 30 which are xed respectively to the shafts 20 and 22. A burner section is provided downstream of the compressor wherein fuel may be injected and the fuel-air mixture ignited to produce an expanding high Velocity gaseous medium for rotating the turbine blades 34 and 33 and for further providing a propulsive jet stream in the exit passage 38. The blades 34 and 36 are operatively connected to the shafts 22 and 20 respectively whereby the rear turbine rotor drives the forward compressor rotor 28 and the forward turbine rotor drives the rear compressor rotor 33. The blades 34 and 36 of the adjacent turbine rotors are oppositely inclined relative to the axis of uid flow so that the shafts 2D and 22 and their respective compressor rotors 2B and 33 are counterrotating. For this reason the compressor rotor blades 5l) and-52 will also be oppositely inclined relative to the axis of flow. The principle operation of this compressor will be more fully described hereinafter.

Figs. 2 and 3 indicate a modified drive mechanism for interconnecting the compressor rotors 60 and 62 to the turbine wheel 64. A drive shaft 66 is directly connected to the turbine wheel 64 and the forward compressor rotor Bl] while the aft compressor rotor 62 is driven in a direction opposite to that of the rotor B0 by means of a planetary gear system. This gear system comprises an internally toothed ring gear l0 carried by the rotor S2, a pluralityof pinion gears l2 which are rotatably mounted in the xed central housing 16 and a sun gear '18 integral with the drive shaft 66. With this construction, then, a single turbine Wheel may be utilized to rotate both compressor rotors in opposite direction. The various components of the planetary system may be varied in size as desired to obtain proper relative speeds between compressor rotors 6B and 62. f

Where the rotor blades in each rotor are substantially similar in shape as shown, the rotor 3D, for example, will be rotated at a .somewhat lower speed than the upstream rotor 28 in order to obtain approximately identical supersonic velocities between the blades of both rotors.

Although the iuid leaving the upstream rotor 28 has an absolute velocity which is less than the absolute inlet velocity it leaves the rotor 28 in a direction such that due to the rotational Velocity of rotor 3c it has a relative direction to rotor 3B which is substantially parallel to the chord line of the oncoming blades 52 thereof. Thus the velocity of the fluid entering the rotor til is high relative to the moving rotor blades 52 and it is therefore possible to rotate the rotor 30 slower than rotor 2S and still maintain substantially the same supersonic relative velocities through each successive rotor.

It may further be desirable to decrease the pitch of the blades on successive rotors inasmuch as the relative entrance flow to the downstream rotor is more in line with the chord of the approaching blades.

If for certain reasons a higher Mach number were desirable between the rotor blades of the rear rotor 39 than that being obtained through the blading of rotor 28, the aft rotor 30 may be Vrotated at approximatelyV the same speed. Also,

such identical rotor velocities may be required in the event that the blade structure of each successive rotor varied sufficiently so as to create a different flow velocity through the blades. Thus, for example, should a third oppositely rotating rotor stage be used it would be desirable to have the absolute exit ow from the second rotor to have adirection more in line with the blades of the third rotor rather than in the axial direction shown in Fig. 5.

The utilization of successive counter rotating rotors with supersonic flow between the blade passages also provides a self-compensating effect which insures that the relative iiow through the blade passages will continue to move substantially in line with the chord of the blades and thus insure smoothness and proper directional movement of the exit air. This self-compensating effect is best illustrated by referring to Figs. 5a, 5b and 5c which show various conditions of relative ilow into the blades of the second rotor.

Fig. 5a. shows the relative air entering the blades of the second rotor St substantially chordwise of the blades and represents the optimum operating condition of this rotor. Fig. 5b illustrates, in exaggerated form, the direction of relative inlet now when the fluid approaching the blades 52 is such that the blades are at a negative angle of attack. As the flow enters the blade passages an expansion field is generated at the leading edge of the blades adjacent the face thereof. In moving through the expansion eld the air flow is deflected toward the face of the blades as shown so that it subsequently flows substantially parallel to the face of the blades. At the same time an oblique shock will be produced near the leading edge on the backside of the blade and the fluid flowing through this shock will also be deflected so that the flow will continue downstream substantially parallel to the chord of the blades. The bending of flow through ythe expansion eld and the shock is more clearly shown in detail in Fig. 5d.

Referring to Fig. 5c a condition of iiow opposite to the Fig. 5b situation is shown wherein the relative inlet iiow to the rotor is such Ythat the approachingrotor blades are relatively at aV high positive angle of attack.v Under this condition an expansion eld will be produced near the leading edge and on .the backside of each blade while a shock will be generated adjacent the face of each blade to deflect the flow. Conditions shown in Figs. 5b and 5c will tend to decelerate and accelerate the inlet axial now respectively so as to approach condition shown in Fig. 5a.. n Under all the foregoing flow conditions the blade passages will still generate a normal shock downstream in the vicinity of theA diverging portion of the'pa'ssages. Y

It is to be understood that the same conditions may exist at the blade entrance of the i'lrst rotor. It is then apparent that with the self-stabilizing effect that the use of stators or turning vanes can be completely eliminated.

Fig. 6 illustrates the comparsion of theoretical pressure ratio and efficiency at various Mach numbers that may occur when a fluid passes through a normal shock which stabilizes in the blades passages of a supersonic compressor. A similar comparative chart may be used as a criterion for choosing design proportions. It will be noted that for example a pressure ratioof 3 is obtainable in each stage with an efiiciency approximately percent and this pressureA ratio may be further increased at higher Machunumbers with a corresponding decrease in elciency. However, even with some loss of efficiency, at increased Mach numbers within the rotors, extremely high compression ratios are available thereby permitting extremely high power output in a turbo-jet engine and the like. It is further apparent that with one stage producing a compression ratio of 3 the introduction of a second stage of similar theoretical efliciency and without having the normal losses usually encountered with stators, an overall compression ratio of 9 could readily be produced. A compressor of the type described would normally be designed for operation in a region between the lines A and B which are superimposed on the chart of Fig. 6. In other words, the shape of the blades, the size of the blades, the spacing of the rotors and the rotational velocity of the rotors would be of such values that most efficient operation would be obtained for example in the Mach number range between the lines A and B to thereby produce the corresponding compression ratios and eiliciencies. The compression ratios and efficiencies illustrated in Fig. 6 are by way of example only and are not limting ranges of this invention.

The axial spacing of the rotors of the compressor of this invention may be increased somewhat to permit pressure bleed-off and stabilization of flow conditions during starting and offdesign operation. Also the hub tip ratios of the compressor may be altered to obtain optimum passage flow conditions and radial equilibrium.

It is therefore apparent that as a result of this i invention it is possible to obtain a very high pressure rise across each compressor rotor thereby providing a high pressure operating unit having a high mass flow capacity with minimum space and weight requirements. Such high pressure and high mass flow capacities permit a turbo-jet engine or the like to operate at unusually high power output with the possible expense of some losses which are, however, within a practicable limit.

Further, as a result of this invention it is apparent that a high capacity compressor has been provided which is readily adaptable to a greater number of stages than that illustrated.

It is also apparent that as a result of this invention a compressor and power plant is provided Which is readily adaptable for aircraft inasmuch as the counterrotating compressor and turbine rotors produce no gyroscopic forces during aircraft maneuvers.

Although certain embodiments of this invention have been illustrated and described herein, it is apparent that various modifications and changes may be made in the arrangement and construction of the various parts without departing from the scope of this novel concept.

What it is desired to obtain by Letters Patent is:

A compressor comprising a casing structure having an annular substantially unobstructed airflow passage, a plurality of compressor rotors Liournalled axially in said casing structure, blading carried by each of said rotors and extending across said passage, means for rotating each successive downstream rotor at proportionally lower rotational speed and in a direction opposite to its preceding rotor, said blades having a thickness which reduces toward their trailing edges for at least a distance from mid-chord to said trailing edges thereby forming diverging passages therebetween so as to impart a normal shock to the gaseous medium entering said blades to provide an immediate reduction in the velocity of said medium relative to said rotor blades and thereby eiiecting an immediate increase in pressure of said medium.

References Cited in the le of this patent UNITED STATES PATENTS

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