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 a fluid reaction propulsion system for aircraft.
| Identifier | ExFiles\Box 147\2\ scan0191 | |
| Date | 17th February 1937 | |
| 4 471,368 and may generally follow the form described in relation to my British Patent Specification No. 456,980. It may be convenient to provide the whole device in the form of a unit comprising hollow nacelle with an entry opening to face the direction of travel (whereby incidentally a certain head of pressure termed “ Pilot pressure ” may be derived) and in such nacelle all those components apt to lose heat may be completely housed to conserve energy. The invention is further to be understood by reference to the following description. Where the description or claims seem to refer to the whole or a determined part, of a flow, it is to be understood that a portion thereof may be separated and utilised for such purposes as driving auxiliaries, cabin heating, etc. The accompanying drawings diagrammatically illustrate applications of the invention : Figure 1 is a partial section in plan of a twin aircraft propulsion system, employing a type of combustion unit more fully explained in my British Patent Specification No. 456,980; Figure 2 is a partial elevation section of a further alternative example. Referring to Figure 1 there is illustrated what may be termed a twin apparatus, that is, it is symmetrical about a central line. For aircraft use it may be important to select the directions of rotation of rotatable parts, having regard to their gyrostatic reactions. Considering this apparatus as comprising virtually a port and a starboard intake and accompanying reaction jets, only one side will be described, the other, apart from any question of rotational direction, being substantially identical. The initial compressors are of the axial flow single-stage type, whilst the internal combustion engine includes a constant pressure gas turbine of the type described below. The arrangement involves a first stage axial flow compressor 50 supplied by an annular intake orifice 51, and discharging into a nacelle 52 within which, behind the compressor 50 is driving gear 53 powered by a shaft 54 connected to a driving turbine 55 within a casing 56A. The main flow created in the first place by the compressor 50 and subject to any Pitot pressure at 51 of which advantage can be taken, passes rearwardly through the nacelle 52 to emerge from the nozzle 52A. The flow from the compressor 50 is divided; part of it is diverted, (and any suitable guiding means such as baffles may help to divert it) into a lateral trunk 56 which is common to both starboard and port nacelles. Within the trunk 56 is housed a two-stage bilateral intake compressor, a fuller description of which can be found in my British Patent Specification No. 456,976. In short, this comprises twin first stage compressors 57 outputting to the second stage compressor 58. From the secondary diffuser of the compressor 58 the second stage output is led as indicated by the arrow 59 to the delivery chamber 60 of a turbine, the mechanical output of which drives the compressor shaft 61 common to both 57 and 58. In the passage indicated by the arrow 59 is a combustion chamber with any suitable fuel-burning means, arranged for example somewhat in the manner of that described in relation to the above numbered Specification. This unit comprising compressor means, combustion means, and turbines, is an internal combustion engine. It is here noted that as between port and starboard sides of the whole apparatus the working directions must of course be appropriate for both port and starboard turbines to drive the shaft 61 mutually. The combustion effluent expands through the vanes of the gas turbine rotor 62 and by a passage 63, 64, 65 is delivered to the nozzle scroll of the turbines 55 through which these gases 8 expand further rejoining the main flow in the nacelle 52 through the exhaust 66. In the alternative illustrated diagrammatically in Figure 2, the propelling unit comprises a nacelle or duct 30 with a forwardly facing entry 30A and a rearwardly facing exit or propulsion nozzle 30B. This nacelle encloses within its entry 30A an axial flow compressor with rotor 31 driven by a shaft 32 from an axial flow turbine rotor 33. In practice there will be step-down gearing between the turbine 33 and rotor 31. Also within the nacelle 30 is a compression ignition engine represented at 34; this engine drives through its gear box 34A a centrifugal bilateral intake compressor 35 with its intakes (indicated by arrows) collecting from the interior of the nacelle 30. The output of the compressor 35 is led by a duct indicated at 36 which leads such output in the direction of the arrows 36A to the nozzle scroll 37 of the turbine 33, the effluent therefrom escaping through the duct 38 and a rearwardly facing propulsion nozzle 39 with such energy as remains after passage through the turbine 33. This unit therefore consists in a first compression with divided output, a second and centrifugal compressor intaking the diverted flow of air, driven by and supplying part of its air output to the compression ignition engine, from which rejoins the rest of the air supply for the said second compressor to form the working fluid in the turbine 33. To this end the engine 34 has its intake 34B in the duct 36A and its exhaust 34C returning to that duct. The engine 34 may also have its own supercharger, and of course its own auxiliary apparatus. The whole output of the first compressor is finally employed for reaction propulsion. It may be found possible to enhance the propulsive efficiency by deriving some or all of the air flow from the boundary layer at or over any desired part of the structure; and it may also be possible to make use of the discharged gases to modify or improve aerodynamic effects. Having now particularly described and ascertained the nature of my said invention and in what manner the same is to be performed, I declare that what I claim is :— 1. In a fluid reaction propulsion system for aircraft, the combination of an air compressor, an arrangement which in effect divides the output from the compressor into a first stream which is passed out through a propulsion nozzle and a second stream, an internal combustion engine supplied by the second stream, and a gas turbine supplied wholly or partly by the effluent gas from said engine and driving the said air compressor. 2. A system as set forth in Claim 1, in which the combustion products of the engine also contribute to the thrust by fluid reaction. 3. A system as set forth in Claim 2, further characterised in that the combustion products rejoin the remainder of the air output before final expansion. 4. A system according to Claims 1, 2, or 3, having a plurality of compressors each with its own related driving turbine, mutually combined with a single internal combustion engine. 5. A system according to any previous claim, in which the engine itself includes a gas turbine other than that which drives a compressor. 6. A propulsion system for aircraft as claimed in any previous claim 1-4 constructed as a unit, comprising an air duct with an entry opening to face the direction of travel, an axial flow compressor therein, and an internal combustion engine together with its cooling appurtenances wholly within said duct, being said compressor, the duct having an outlet orifice facing oppositely to the entry. 7. A unit according to Claim 6, substantially as described and illustrated in Figure 1. 8. A unit according to Claim 5, in which the gas turbine is a constant pressure gas turbine employing centrifugal compressor means for its compression, continuous combustion, and expansion through turbine means which provide power for the compression. 9. A unit according to Claim 8 in which the compressor means is a multi-stage centrifugal compressor. 10. A unit according to Claim 8, in which there are two mechanically independent turbine systems, one of which drives the compressor means of the gas turbine, and the other of which supplies the remainder of the power required for the energising of the whole air throughput. 11. A system according to Claim 5 in which the gas turbine comprises a centrifugal compressor driven by, and supplying part of its air output to a compression ignition engine the exhaust from which rejoins the rest of the air supply from the said compressor to form the working fluid in a turbine. 12. An aircraft provided with and adapted to be propelled by one or a plurality of the devices substantially as described. Dated this 17th day of February, 1937. For the Applicant, F.{Mr Friese} J.{Mr Johnson W.M.} CLEVELAND & COMPANY, Chartered Patent Agents, 29, Southampton Buildings, Chancery Lane, London, W.C.2. Leamington Spa : Printed for His Majesty’s Stationery Office, by the Courier Press.—1937. | ||