Patent GB-1069217-A (GB)
Title: Improvements relating to engines
Inventor: MACPHERSON IAN
IPC: F02C5/10
Filed: 1965-03-29
Abstract: 1,069,217. Gas turbine combustion chambers; gas turbine blades; gas turbine plant; jet propulsion plant. ROLLS-ROYCE Ltd. March 22, 1966 [March 29, 1965], No. 13295/65. Headings F1G, F1J, F1L and F1T. [Also in Division G3] A gas turbine engine has an annular combustion chamber 19 around which at least one combustion initiating detonation wave continuously rotates at supersonic velocity. Liquid fuel is continuously supplied by a fuel injector 30 upstream of the combustion chamber 19 so that an atomized mixture of fuel and air enters the chamber through ports 29 defined by baffles 27. The detonation wave is initiated by an igniter 31 and the combustion gases behind the rotating wave flow continuously, either by expansion from their initial high pressure or by scavenging by the incoming combustible mixture, through exhaust ports defined by baffles 28 to drive the turbine 20. The flow thus pulsates around the circumference of the turbine. The gases leaving the combustion chamber are at a pressure higher than the incoming mixture. The maintenance of the detonation wave is effected by suitable choice of the size of the inlet and outlet ports of the combustion chamber and the rate of supply of the combustible mixture. Part of the air from the compressor 14 by-passes the combustion chamber 19 and is further compressed by blades 23, 24 integral with the turbine rotor blades 21 before mixing with the combustion gases downstream of the turbine. The by-passed air cools the blades 21. In a modification. Fig. 3 (not shown), having a combustion chamber (40) similar to chamber 19, the turbine rotor blades (41) are integral with inwardly disposed blades (43) of a ducted fan (44). The axial flow compressor rotor blades (37), driven by the turbine, are also integral with inwardly disposed fan blades (45). The combustion chambers 19, (40) may be cooled by air or by a flow of fuel which is subsequently introduced into the combustion chamber. In a modification, chambers 19, (40) are replaced by a combustion chamber (46), Fig. 4 (not shown), in which to reduce flashback through the inlet conduits (50) a finned member (54) is mounted in an enlarged portion of each conduit. The member (54) has a central body (57), Fig. 5 (not shown) with narrow passages (60) therein closed at one end to trap blown back combustion gases. The trapped gases after cooling flow back into the combustion chamber as a buffer layer ahead of the incoming combustible mixture to prevent premature ignition thereof. The inlet conduits (50) communicate with the compressor through narrow passages (61). Walls (63) separating conduits (50) have cooling passages (64) therein supplied with fuel or air as the coolant fluid. A cooling jacket (62) also surrounds the combustion chamber (46). In a modification, the axial flow compressors 14, (37) are replaced by centrifugal compressors. Liquid fuel may be replaced by gaseous fuel or a liquid or gaseous monopropellant. The fuel and air may be supplied separately to the combustion chamber. In a further modification the combustion chamber is mounted in the by-pass duct of a gas turbine by-pass engine.