US2724238A

US2724238A - Jet propulsive devices

Info

Publication number: US2724238A
Application number: US191979A
Authority: US
Inventors: Tresham D Gregg
Original assignee: Individual
Current assignee: Individual
Priority date: 1950-10-25
Filing date: 1950-10-25
Publication date: 1955-11-22
Anticipated expiration: 1972-11-22

Description

Nov. 22, 1955 T. D. GREGG JET PROPULSIVEDEVICES 4 Sheets-Sheet 1 Filed Oct. 25, 19.50

Nov. 22, 1955 T. D. GREGG JET PROPULSIVE DEVICES 4 Sheets-Sheet 2 Filed Oct. 25, 1950 INVE TOR..

Nov. 22, 1955 Filed Oct. 25, 1950 ENTHALP: 5.711 PER 1.5

T. D. GREGG JET PROPULSIVE DEVICES 4 Sheets-Sheet 3 6 ENT'ROPY FIG. 8

IN NT R.

1955 T. D. GREGG 2,724,31

JET PROPULSIVE DEVICES Filed Oct. 25, 1950 4 Sheets-Sheet 4 2,724,238 v JET PROPULSIVE DEVICES Tresham D. Gregg, New York, N. Y.

Application October 25, 1950, Serial No. 191,979

6 Claims. 01. 60--35.6)

Thisinvention relates to jet propulsive devices and particularly to means of producing such propulsive reaction by a jet from a gas turbine engine having a nozzle whose area of cross-section may be varied at will in conjunction with a segmentally-adjustable aerofoil ring thrust augmentor.

These and other objects of thisinvention are now summarized as follows: 3

1. In typical jet propulsion engines such as the Goblin II described infthe issue of February 24, 1946 of the magazine Flight, the tail pipe or propulsion nozzle is of fixed area. This fact limits the range of thrust economically obtainable from the engine and also limits its propulsive efficiency at speeds less than400 miles per hour. This invention increases such range of thrust and also increases the propulsive efficiency of the engine at moderate speeds by providing the engine with a segmentally-adjustable nozzle or tailpipe having a wide range of area variation. This is a tube consisting of a plurality of cyliudricallyor comically-curved plates over-lappingly assembled and movable radially with respect to the axis of the nozzle, being maintained in close contact at the overlaps by the pressure of the fluid within the tube.

2. The principal loss of energy and thrust in all fluid reaction propulsive devices now in use is theso-called leaving loss, that is, the kinetic energyof the propulsive jet as it leaves theprojecting nozzle. One of the objects of this invention, therefore, is to reduce to a minimum the leaving loss of the propulsive jet from a gas turbine jet engine. This object is bestaecomplished by adding to the propulsion system an aerofoil ring, preferably one that is segmentally-adjustable. Such a device utilizes the otherwise wasted kinetic energy of the driving jet as it leaves the nozzle by entraining and accelerating through the ring a large volume of outside airwhich, as it passes through the ring, mixes with the hot gas of the jet, cools it, reduces its speed, and at the sametime increases the flow of momentum of the system and henceits thrustybut without at the same time increasing the fuel consumption of the engine. a

3. In order to make effective the addition of such an aerofoil ring to the system, the stream of outside air thus drawn into the aerofoil ring duct should be guided smoothly into contact with the driving jet at the critical point in the ring channel for maximum thrust augmentation, through an annular duct having no sudden variations of area and adaptable for all possible nozzle diameters. The invention accomplishes this by attaching to that portion of the nozzle which projects beyond the nacelle or engine fuselage a segmentally adjustable conoidal sur- 2,724,238 Patented Nov, 22 1955 the inducted air must pass through an air gap between the fuselage and the leading edge of the aerofoil ring in a direction normal to the ring axis. Such an arrangement makes the ratio of the upstream reaction or thrust boost of the ring to the thrust of the jet substantially independent of the aeroplane speed. i

5. Since the aerofoil ring when in efiicient operation can deliver an independent thrust of 40% or more of the normal jet thrust at all speeds, it must be supported in the correct position with respect to the aeroplane under all conditions of speed and direction. This is accomplished by connecting the leading edge of the aerofoil ring to the nacelle or fuselage by means of a plurality of streamlined struts or links passing through close-fitting slots in the nacelle or fuselage and strong enough not only to transmit the ring reaction to the aeroplane but also to resist all stresses due to the weight of the augmentor and to the bending moments and shears and tortional forces due to sudden changes in direction of the aeroplane in flight. These struts or links must also supply means of controlling the width of the air gap and thus supply the means of making quick changes in thrust such as are frequently needed during landing and take-off, without requiring corresponding changes in the engine throttle with the consequent danger of killing the engine at a critical time. This device also serves the important function of giving the pilot a continuous measure of the net aerofoil ring thrust. The method of support also permits the necessary expansion and contraction of the ring.

6. All bodies moving through a fluid, whether streamlined or not, are subject to the drag of skin friction. Skin friction is proportional in intensity to the thickness of the boundary layer, a thin layer of air which virtually attaches itself to the surface of the moving body. The boundary layer begins to form at the forward end of the body and gradually increases in thickness toward the rear. Friction drag can be materially reduced by causing the boundary layer to be sucked into the interior of the body at intervals over its surface. The air gap between the nacelle and the aerofoil ring permits this sucking in of the boundary layer that has formed upstream so that the skin friction drag intensity on the outside surface of the ring is thereby materially reduced.

7. It is an important object of one form of this'invention to make possible a tail pipe or nozzle of variable discharge area having a shape and length suitable for after burning either with or without the addition to the system of an aerofoil ring thrust augmentor. After burning, as the name implies, is the mixing of new fuel with the turbine exhaust gas in the tail pipe to be there ignited and burnt at a very high temperature, the greatly expanded gas being then discharged through an enlarged nozzle at correspondingly increased velocity and augmented thrust. Before igniting this gas mixture two things are essential, namely, the tail pipe channel must be so shaped as to diffuse or recompress the gas after it has passed through the turbine and, second, a device or devices must be inserted into the channel after the fuel has been introduced to mix it thoroughly with the exhaust gas before it is ignited. Since such devices materially reduce the: thrust, especially for normal operation, they should be automatically removable after performing their function. These objects are accomplished in this invention first by forming the whole tail pipe from a plurality of overlapping segments movable radially in the same manner as the adjustable nozzle above described, second, by giving the pipe assembly the longitudinal variations in diameter required for diffusing the gas, and finally by utilizing the radial movement of the component segments by a simple mechanism to raise and lower a plurality of hinged plates capable of stirring up the passing gas during after burn} ing and .of .being folded .out of .the .channel .for normal operation.

These and other objects of'this invention will be understood when the following description is read in connection with :the attached drawings.

Fig. :1 .showsaside elevationof the engine nacelle and the attached aerofoil ring thrust augmentor.

Fig. 12 shows the venginewwith its nacelle partly broken away, the adjustable nozzle in a contracted position in axial section together with a portion of the forward end :of the aerofoil ring, also in a contracted position.

Fig. 2a shows a part of the rear end of the engine nacelle with a portion of the nozzle and of the forward part of the :aerofoil ring, both in an expanded position.

LF-ig. 2b shows in enlarged section a portion of the gas turbine with parts of the inner cone or bullet and its supporting struts together with the conical part, in seciion, of the segmentally adjustable tail pipe.

Figs. 3 and 3a .show partial sections C--C and BB of the segmentally adjustable fairing cone F and the after-face of the nacelle.

Fig. 4 shows a portion of section A--A of the tailpipe with its adjusting engines in the contracted position.

Fig. 5 shows a cross-section of the cylinder of one of the engines which supports the aerofoil ring and controls the-air gap.

Fig. 6 shows an .enlarged partial elevation of the conoidal initial portion of the adjustable nozzle with its hinged component members and one dilating or adjusting engine.

Fig. 7 shows a partial cross section and upstream axial view D--D of the initial portion of the tail pipe in both contracted and dilated positions, shown in elevation in Fig. 6.

Fig. 8 shows a Mollier or enthalpy-entropy diagram defining graphically the characteristic changes in enthalpy, pressure and entropy which take place in a typical turbo jet engine alone and when combined with an aerofoil ring thrust augmentor both in the normal and 'afteriburning" or reheat cycles.

Fig. 9 shows the combination of a tail pipe arranged and. proportioned for after burning in conjunction with an aerofoil ring thrust augmentor as in the second embodiment of this invention.

Fig. 10 shows in enlarged detail a partial transverse section of the tail pipe with the stabilizers and their operat-ing' mechanism both in the after burning and normal position.

Figs. 11, 12 and 13 are enlarged details of tangentially operatingpiston engines for use if space is too small for the radial engines, shown in partial cross-section, both in open and closed position, and in elevation, of a segmentally-adjustable structure composed of overlapping plates.

Fig. 2 shows in sectional form a normal gas turbine engine having an adjustable tail pipe or nozzle in conjunction with an aerofoil ring thrust augmentor partially shown, also in section, which forms the first embodiment of the present invention.

In Fig. 2, 6 represents one of the air intake orifices which are connected by expanding ducts to a turbocompressor, as 7. The compressor is connected to sixteen burners, two of which, designated 8, are shown. These are joined at their downstream ends to a nozzle ring casting 24 shown in enlarged section in Fig. 2b. This casting is supported partially by struts 26 and partially by the cone 18. This cone also supports thedownstream bearing'of the hollowshaft 12 which carries the gas turbine rotor 11. A plurality of stream-lined struts 25, integral with the nozzle ring casting 24, support the cone or bullet 21 which forms the inner wall of the annular channel 20. The outer wall 22 of this channel, formed of a plurality of overlapping conoidally curved plates is connected by hinges 23, Fig. 2b, to the nozzle ring casting 24. The downstream ends of plates 22 are slidingly connected by means of the slotted double hinges .30., Figs. .2 and 6, to corresponding segmental plates forming the discharge nozzle 27. Each of these plates has a longitudinal stiffener 31 to which are attached the piston rods of two independently controlled sets of radially operating double-acting

piston engines

32 and 32a, shown also in Figs. 4 and 6 with two fluid supply manifolds 49 and manifold feed pipe 50, supported by a cylindrical plate or drum 33 which bears against a plurality of blocks 17 fastened to the engine nacelle 19. This nacelle is squared with the .engineaxis at its downstream end and has a rounded corner, forming a windshield 1%, Figs.

2, 2a and 3a for the aerofoil ring thrust augmentor 2,

Figs. 1, 2 and 2a, which is held in position by a plurality of struts or links 3 attached by pins 16, Figs. 2 and 2a, to the leading edge of the ring and by pins 15, Figs. 2 and 2a, to the piston rods 13 of a corresponding number of double-acting hydraulic engines 14, Figs. 2, 2a, 4 and 6, attached to the nacelle 19. Slots 1%, Figs. 3 and 3a, are provided in the corner of the nacelle and in the blocks supporting the drum '33 to permit the struts 3 to move radially as well as longitudinally and fitting the struts or links 3 closely so as to prevent any twisting motion of the aerofoil ring which they support. Two fluid supply pipes .51 provide the driving fluid for the said engines 14.

The outer or downstream portion of the nozzle extends beyond the squared end of the nacelle to the critical section of the aerofoil ring duct, called the aerodynamic center of the" ring, the position of this center being fully described in my Patent No. 2,475,022, column 3, lines 44 to 51, and column 7, lines 7 to 30. To each of the segments of the nozzle 27, Figs. 2, 2n, 3 and 3a, is fastened a radial diaphragm 36 which has a curved outer edge to which is fastened a conoidally curved fairing plate 35 extending from the, nacelle to the nozzle discharge orifice. These plates when overlappingly assembled form a smooth fairing cone F, as shown enlarged in Figs. 3 and 3a, integral with the diaphragms 36 and the nozzle plates 27, and bear slid-ingl'y, through an anti-friction device if required, upon the squared rear end of the nacelle 19b, and 1 Fig. 2a shows the nozzle fairing cone assembly F in its dilated or expanded position in relation to the forward end of the aerofoil ring correspondingly expanded as for maximum thrust, which forms, with the inner surface of the ring, the annular channel 28. In each of Figs. 1, 2 and 2a the variable air gap 4 between the engine nacelle at and the leading edge of the aerofoil ring is shown As above indicated, a second embodiment of this invention consists of a gas turbine engine with a segmentally adjustable tailpipe suitably shaped to serve in expanded position as an after burner in conjunction with an aerofoil ring thrust augmentor; and also in contracted position as a normaljet projector. Fig. 9 shows such an arrangement. The

parts

21', 22, 23, .24, 25,26- and 30 are also shown in-Figsy2, 2b and 6. 20a is' anannular channel such as 20 of Figs. 2 and 2b but greatly expanded. In Fig. 9, channel 37 is formed of a plurality of overlapping conically curvedsegments connected to the corresponding. segments of channel20a by the slotted double hinges 30, Figs. 6 and 7-. The segments 37 are joined integrally as by welding or other means atthe point of maximum expansion with corresponding convergent-divergent segments of channel 370, the divergent-portion of which extends with its attached fairing coneF beyond the nacelle into the duct of the aerofoil ring 2, hereshown expanded for maximum thrust augmentation. At or near the junction of- 37 and 37a are located a plurality of-hinged plates or stabilizers 38 with arms 39;projecting outsidethe pipe. These plates which project into the pipe channel during after burning as shown in Fig. 9 and in enlarged partial cross-section; in Fig. 10 stir up or mix the'passing gas as previously mentioned. When the tail-pipecontracts for normal operation as shown by the broken lines in-Figs. 9 and 10, these stabilizers are folded back to lie fiat against the Wall of the pipe by means ofthe outside projecting arms 39"wliiciieon'nectb'y means of universal of blocks 43 attached to the nacelle.

. the gas.

joints 40 to other; arms 39!: which are held bysimilar joints 40 to support castings 41 resting uponacircular plate or drum 42. This drum is ,carried on a-plurality The tail pipe 37.--37a issupported and moved radially by piston engines arranged in two independently operating

sets

32 and 32a carried on drums, as 33, which arein turn supported on

blocks

44 and 45 the necessary driving fluid being conveyed to said engines by manifolds49 and manifold supply pipes 50. The outboard section .of the nozzle carries the fairing cone F as. previously described bearing slidingly on the after-face of the nacelle at 19b, Figs. 2a, 3 and 3b. e I

.An alternative device for effecting the circumferential expansion and contraction of the nozzle ducts of this .invention, or any structure composed of overlapping plates, where space is too small for radial engines, is shown contracted in Fig. 11 and expanded in Fig. 12, the tangentially-operating. piston engines 46 being connected to the adjacent overlapping plates by means of pins 47 and blocks 48. s The blocks 48 serve as stops to limit contraction and the piston cylinders serve as stops to. limit expansion. Fig. 13 is an outside elevation of three of these engines and attached plates. h

, In the operation of this invention in conjunction with a turbo-jet engine such as the Goblin II. air normally at high speed relative to the plane enters two intake orifices, one of which is shown in Figs. land 2 M6, at atmospheric pressure and heat content Po and Ho respec tively on the diagram Fig. 8. It passes through ducts to a compressor as 7, Fig. 2, which it enters at a pressure P1 called the ram pressure" with a heat content or enthalpy of H1 in B. t. u. per lb. of gas as noted on the diagram, Fig. 8. In the compressor it is raised in pressure and enthalpy to P2 lb. per square inch and H2 units of enthalpy and passes thence into 16 combustion chambers of which 2 are shown as at 8, Fig. 2, into which liquid fuel is fed from a manifold 9 which mixes with a the compressed air and is ignited, burning at a high temperature, its total heat represented by H3 and at a pressure P3. The gas thus produced expands continuously through turbine nozzles formed of aerofoil shaped vanes or blades, Fig. 2b at 10, in the nozzle casting 24, against the moving blades of the turbine rotor Fig. 211,11. The hot gas leaves the turbine at pressure P4 and enthalpy H4, Fig. 8, whence it continues to expand through the annular duct 20, Figs. 2 and 2b, formed by the, bullet 21 and the converging outer cone 22, and thence through the duct 27 to its exit in the central duct of the aerofoil ring where it has the pressure P6 and heat content He.

In the case of a normal jet engine, such as the Goblin II without the aerofoil ring, the discharge pressure would be asmospheric, P0 and the enthalpy H5, Fig. 8, the heat H5-Ho B. t. 11. per lb. of gas being rejected in this cycle of operation. This rejected heat is proportional to the leaving loss or kinetic energy of the discharged gas. In the first embodiment of the present invention, however, the discharge pressure P6 and enthalpy H0 at the nozzle exit do not represent the final state of After leaving the nozzle the gas continues to do useful work by drawing into the aerofoil ring through the air gap 4 and the annular duct 28, Figs. 1 and 2, a

much larger weight of outside air by reason of the pressure drop PO-PG; and accelerating it and compressing it to atmospheric pressure at the trailing edge orifice 29. When this air passes the exit orifice of the nozzle 27 it has the pressure P6 and enthalpy H7, Fig. 8. Between this point and the trailing edge orifice of the aerofoil ring 29, Fig. 1, it is mixed with the hot gas from the driving nozzle 27 and the mixture is discharged through the trailing edge orifice 29 at atmospheric pressure Po and enthalpy Ha as shown in Fig. 8. The lost heat Ha-Ho per lb. of driving gas is seen to be less than in the case of the normal jet engine as above described, and a consequent reduction .in the leaving loss with a correspondingincrease in thrust is thus achieved by the first embodiment of this invention.

In the operation of the after burner which constitutes a second embodiment of the invention, the gas leaving the turbine passes into a gradually enlarging annular duct at 20a, Fig. 9, formed by the bullet 21 and the same segmentally hinged conoidal surface as 22, Figs. 2 and 6, in the first embodiment above described, each of the component segments of which is also attached by means of slotted double hinges as 30, Figs. 2, 6 and 7, to corresponding segments of the after burner tail pipe, the first section of which, 37, continues to increase in diameter.

Between the hinged joint and the point of its maximum diameter liquid fuel is sprayed into duct 37, by means not shown. At or near the point of maximum diameter where 37 joins the convergent-divergent section 37a, Fig. 9, the, stabilizer plates 38, Fig. 9, stir up the passing gas to cause the fuel to mix thoroughly with it. The mixture is ignited and burns at a temperature generally very much higher than thatof the gas which entered the turbine, with maximum enthalpy being H'4 and pressure P4. and expands finally to atmospheric pressure P0 with a heat content H's for the jet engine without the aerofoil ring, rejecting the heat H's-Ho, and to P's and H's at the nozzle exit when such ring is present, whence it mixes with the outside air drawn into the ring as before described, the whole being discharged from the ring orifice 29, Fig. 1, at atmospheric pressure P0 and heat content H's. Again it is apparent that the heat given up to the atmosphere, H's-H0, i. =e., the leaving loss, is less than that of the after burning jet=engine Without the aerofoil ring, i. e., H's-Ho. Hence the statements in the second paragraph of this specification are justified for both of the embodiments of the present invention.

It is obvious that the thrust of a nozzle propelled jet is impressed upon the nozzle walls by the jet, and by the nozzle upon the engine and its nacelle, either directly or indirectly. It is understood, therefore, that the method and means of transfer of jet thrust to the engine nacelle and aeroplane must be adequate and not necessarily those shown and described in the specification.

The speed of the. driving jet is normally equal to or greater than that of sound. The jet therefore has great lateral stiffness due to its momentum and, together with the annular stream of inducted air surrounding it which also has great momentum, forms an elastic support to the aerofoil ring against sudden lateral movement and shocks occurring during flight.

While the specific means for applying pressure to the various piston engines described herein have been shown on the drawings, it is to be understood that any suitable means, either manually or automatically controllable, may be employed, such as those shown and described in my Patent No. 2,475,022 for tangentially controlling a segmentally adjustable aerofoil ring.

As at present constructed, tail pipes for after burning" and particularly the means now in use for varying the area of discharge are subject to severe temperature stresses due to the great variation in temperature that occurs between the entrance to and the exit from the tail pipe; and also between normal and after burning operations. Moreover, the devices now in use for varying the area of discharge are subject to the danger of over confining explosive gas mixtures. The present invention, because of the segmental character of the tail pipe, readily adjusts itself to unequal expansion and contraction due to large temperature differences, and the discharge orifice gives free egress to exploding gas.

As indicated in paragraph numbered 1 in column 1 it is assumed that the surfaces at the points of contact between the overlapping component segments of the adjustable nozzle or tail pipe will be smooth and accurately manufactured so that the gas pressure will make the junctures gas-tight.

Pipes for conveying cooling air from the compressor to ,7 the hollow turbine shaft may be passed through the struts 25 which support the inner cone as in the Goblin 'II engine.

.I claim:

1. A gas turbine jet aeroplane engine in combination withan aero'foil ring thrust augmentor, said engine having a nozzle and also havinga stream lined body immediately in front of the aerofoil ring said body being substantially larger in diameter thanjthe outer diameter of the 'aerofoil ri'ng, said body having a plane surface presented toward the leading'edge of the aerofoil ring and perpendicular to the axis of the said aerofoil ring, said nozzle being substantially smaller than the minimum diameter of the ring duct andof such length that its downstream end is positioned "at the aerodynamic center of the ring.

2. An airplane power plant including, in combination, a gas turbine ,jet engine, a nacelle for said engine having a squared after face, a jet propelling nozzle for said engine having a conoidal fairing extending from its orifice to the square after face of the said nacelle, an aerofoil ring thrust augmentor of segmentally variable diameter coaxial with said nozzle and means to fasten said augmentor to said engine, said fastening means being arranged to vary the width of the air gap between the leading edge of said augmentor and the after face of said nacelle to hold the variable aerodynamic center of said augmentor at the plane of the nozzle discharge orifice and to permit varying the diameter and profile of said augmentor.

3. In combination, a gas turbine jet aeroplane engine positioned in 'a hollow, stream-lined nacelle having a nozzle projecting axially downstream into. the duct of an aerofoil ring thrust augmentor, the rear diameter of said nacelle being substantially larger than the -maximum diameter of the said augmentor, the outer portion of the rear face of said nacelle being square with its axis, said nozzle forming a large annular duct with the inner surface of the augmentor, the entrance to said annular duct being defined by the squared rear face of the said nacelle and the leading edge of the aerofoil ring thrust augmentor adapted to be contracted and expanded; a plurality of, struts adjustably connecting said augmentor to the said nacelle and each arranged to vary its angular position toward the axis of the said augmentor at its forward end during contraction and expansion of said augmentor.

4. The combination with a gas turbine jet airplane engine having a nozzle the diameter of which may be varied of a fairing cone surrounding said nozzle and expandible and contractible therewith an aerofoil ring thrust augmentor into which the nozzle extends, the downstream end of said nozzle being substantially smaller than the minimum diameter of the augmentor duct and located substantially at the aerodynamic center of the said ring section, a nacelle for said engine the squared after face of 8 which ;is in sliding contact with the fairing cone of the noz: 21c and is substantially larger in diameter than the leading edge of the aerofoil ring, means to attach saidaugmentor to said engine comprising a plurality of struts, pivotally fastened at one end to the leading edge of said augmentor and passing through close fitting slots in the engine nacelle to pivots in the ends of the piston rods of a corresponding number 'of double acting hydraulic piston engines within the nacelle, the direction of motion of said piston being parallel to the jet 'engine'axis to control the width of the air gap between the lea-ding'edge of the aerofo il ring and the engine nacelle.

5. A gas turbine jet engine including, in combination, a nozzle of variable cross-section composed of a plura'lit'y of overlapping segments projecting beyond the engine nacelle, the projecting portion of the component overlapping segments being connected by radial diaphragmsto a corresponding plurality of conoidally curved overlapping segments forming an adjustable fairing cone; and means for moving said diaphragms and conoidally curved segments radially with respect to the nozzle axis in conjunction with the nozzle segments to which they are attached. 1

6. A gas turbine jet engine including, in combination, a nozzle of variable cross-section consisting of a succession of convergent and divergent conical sections composed of a plurality of overlapping segments; and means to dilate and contractsaid nozzle by moving the componentsegm'ents radially with respect to the nozzle'axis; a plurality of plates within said nozzle movably attached to 'the component segments .and mechanical means to project said plates into the nozzle duct as the nozzle expands and to depress or remove them as the nozzle contracts.

References 'Cited in the file of this patent UNITED STATES PATENTS 157. 526 Leggett Dec. 8, 1874 $43,182 Hunt July 23, 1895 2,390,161 Mercier Dec. 4, 1945 2,447,100 Stalker Aug. 17, 1948 2,462,953 Eaton et al. Mar. 4, 1949 2,475,022 Gregg July 5, 1949 2,487,588 -;Price Nov. 8, 1949 2,509,890 Stalker May 30, 1950 2,510,506 'Lindhagen et al. June 6, 1950 2,569,497 Schiesel Oct. 2, 1951 2,597,253 Melchior May 20, 1952 2,603,062 Weiler et al. July 15, 1952 2,648,192 Lee Aug. 11, 1953 FOREIGN PATENTS 617,173 Great Britain Feb. 2, 1949 922,032 France J an. 20, 1947 Wed MEN