From the Rolls-Royce experimental archive: a quarter of a million communications from Rolls-Royce, 1906 to 1960's. Documents from the Sir Henry Royce Memorial Foundation (SHRMF).
Patent specification for an aircraft propulsion system, detailing the compressor, turbine, and combustion components.
| Identifier | ExFiles\Box 147\2\ scan0200 | |
| Date | 20th April 1936 | |
| 4 456,980 and other losses between compressor and turbine. The compressor and turbine may be mounted coaxially with their rotors carried on the same shaft and the 5 combustion chamber may be of generally helical form. Other features of the invention will appear hereinafter. A specific embodiment of the present 10 invention will now be described by way of example, by reference to the accom- panying diagrammatic drawings, of which :— Figure 1 is a sectional plan of the com- 15 pressor, turbine and other parts for propelling an aeroplane, Figure 2 is a similar view showing the turbine and associated parts in greater detail, and 20 Figure 3 is a diagrammatic develop- ment of the compressor scroll, the com- bustion chamber, the turbine nozzle and the turbine delivery chamber. The compressor comprises an impellor 25 10, blades 11, and a casing 12 formed with strengthening ribs 13 and is of the general form described in the specification of co-pending patent application No. 14,285/35 (Serial No. 456,976). The 30 impellor shaft is formed in two parts 14 and 15, each of which is formed integrally with a flange 16 by which it is attached to the solid boss of the impellor 10. The shaft portion 15 also constitutes the 35 shaft for an impulse turbine 17 having a delivery chamber 19 of volute form. The composite shaft 14, 15 rotates in suitably arranged bearings 20. The impellor 10 of the compressor 40 discharges into a primary diffuser 21 which takes the form of a bladeless radially arranged chamber. The impellor is driven at a tip speed in excess of the velocity of sound (referred to the state of 45 the gas leaving the tips of the impellor blades) so that, as explained in the speci- fication of application No. 14,285/35 (Serial No. 456,976) aforesaid, it is desiable that the speed of the air leaving 50 the impellor should be reduced before it reaches any fixed part. The primary diffuser 21 discharges into a scroll 22 which is formed with an outlet at its largest part to which is connected a helical 55 combustion chamber 24. This first part of the combustion chamber is of tapering form so as to constitute a secondary diffuser for the compressor and the dis- charge end of the chamber is connected to 60 the turbine nozzle. Figure 3 is a diagram which illustrates the gas circuit from the compressor to the turbine, the parts being shown developed into a plane. The scroll 22 of the com- 65 pressor, into which air flows from the impellor in the direction of the arrows, discharges into the combustion chamber 24 the first part 60 of which is of tapering form to constitute the secondary diffuser above mentioned. The combustion cham- ber discharges into a nozzle 62 which is not shown in Figures 1 and 2. The nozzle discharges into the delivery chamber 19 from which it passes, as shown by the arrows, to the buckets of the turbine rotor. The combustion chamber 24 contains means for heating the air flowing through it, such as a fuel jet 25 surrounded by a cowl 26. The heated air passes from the delivery chamber 19 of the turbine nozzle to the buckets 27 of the turbine rotor through an orifice 28 which is continuous around the rotor so that the whole of the blast of gas through the orifice is always leaving the rotor, the air passes into an annular collecting chamber 29 which merges into a fore-and-aft conduit 30. The annular chamber 29 is formed between a cone 40 and the conduit 30 and is shaped as a diverging channel for the gas leaving the turbine rotor. The cone is supported from the conduit by brackets 41. The conduit passes rearwardly of the aeroplane to form a jet which discharges through a convergent-divergent propel- ling nozzle 31 in the tail of the aeroplane. The general form of the profile of the aeroplane body is shown by the numeral 34 and the compressor may be attached to the body by having lugs 35 formed ra- dially from its casing and to engage with strength members 36 running fore-and- aft inside the profile 34. The profile 34 forms the principal structural member for supporting the propulsion system but the turbine end of the system may be attached to the aircraft by means of the brackets 47 shown in Figure 2. A very high mass-flow of air through the system is obtained by the provision of double intakes 37, 39, one on each side of the plane of rotation of the impellor 10. The air taken into the system may be admitted to the intakes 37 and 39 by any known means. For example, the air may be admitted to the interior of the aircraft structure through a suitable orifice (not shown in the drawing) so that the whole of the interior of the body constitutes a reservoir from which the intakes 37 and 39 can draw. If the orifice is arranged to face forwardly with respect to the direction of flight, then an initial com- pression will be applied to the air prior to its admission to the intakes of the com- pressor. By reason of the double intakes above described and the high speed at which the compressor impellor is driven, it is 456,980 5 possible to obtain a mass-flow load factor (as defined in the specification of application No. 14285/35 (Serial No. 456,976) aforesaid) of 0.25 or more and, as 5 described in the said specification, the impellor 10 is provided with rotating guide vanes and with a large number of ribbed radial blades 11. As shown in Figure 2, the turbine 17 10 may be formed with a shaft portion 42 which is held against rotation in the com- pressor shaft 15 by means of splines 43. The two shafts may be held against wise movement by means of a taper pin 15 or cotter 44 which can be passed into or out from its engaging position through the casing 45 which holes are normally closed by plugs 46. The casing 45 of the turbine is directly 20 attached, as shown, to the casing 12 of the compressor and is also attached by light aluminium-alloy brackets 47 of girder form which pass around the inlet volute 19 to engage lugs 49 affixed to a 25 ring 50 of U-shaped cross-section. The U-shaped ring is of strong construction and is attached at one side to the volute 19 which it supports and to the other side to a portion 51 of the conduit 30. By this 30 construction the volute 19 is relieved of all stresses except such as are caused by the gases within it and the arrangement is such that the parts can be assembled and disassembled. The brackets 47 are also 35 attached to lugs 59 formed on the ribs 13 of the compressor casing. The turbine casing may be formed also with water jackets 52. In an alternative arrangement (not 40 shown in the drawings) the compressor may be driven by the turbine through gearing, and preferably in the opposite direction to the turbine so as to reduce the net gyrostastic couples. 45 An aircraft may carry one or more pro- pulsion systems of the kind herein described. Where two systems are carried, the rotating parts of one may be arranged so as to rotate in the opposite direction 50 from the rotating parts of the other, whereby the resultant gyrostastic couples are reduced to a minimum. Having now particularly described and ascertained the nature of my said invention and in what manner the same is to 55 be performed, I declare that what I claim is :— 1. A propulsion system of the kind described comprising a compressor having two air intakes, one on each side of the 60 plane of rotation of the impellor. 2. A propulsion system of the kind described comprising a turbine for driving the compressor, wherein the mouth of the nozzle for the turbine extends sub- 65 stantially around the periphery of the turbine rotor. 3. A propulsion system of the kind described, wherein the outlet from the compressor is connected to the inlet to 70 the passage which constitutes a secondary diffuser for the compressor and a com- bustion chamber for fuel. 4. A propulsion system according to 75 any of the preceding claims, wherein the compressor and turbine are arranged coaxially and rotate at the same speed. 5. A propulsion system according to claim 3, wherein the said passage is of 80 generally helical form. 6. A propulsion system according to any of the preceding claims, wherein the compressor comprises a primary and a secondary diffuser, and wherein the 85 primary diffuser is constituted by a blade- less radial discharge chamber. 7. A propulsion system according to any of the preceding claims having a mass-flow load factor, as herein defined, 90 of 0.25 or more. 8. A propulsion system according to any of the preceding claims, wherein the compressor impellor and the turbine rotor 95 are interconnected by gearing. 9. An aircraft incorporating a propul- sion system according to any of the pre- ceding claims. 10. An aircraft propulsion system sub- 100 stantially as herein described with refer- ence to the accompanying drawings. Dated this 20th day of April, 1936. BOULT, WADE & TENNANT, Chartered Patent Agents, 111 & 112, Hatton Garden, London, E.C.1. Leamington Spa: Printed for His Majesty's Stationery Office, by the Courier Press.—1936. | ||