Patent GB-1428726-A (GB)
Title: Combustion-stabilizing devices
IPC: F02B15/00
Filed: 1973-07-30
Abstract: 1428726 Combustion equipment eg for gas turbine engines; IC engines L E COLE and J M ZABSKY 30 July 1973 37302/73 Headings F1B F1G F1J and F1L [Also in Division F4] The invention relates to combustion equipment and to means for preventing instability of pressure in a combustion chamber in which a fluidic device is utilized to sense any such instability. In a first embodiment, the invention is applied to a liquid propellant rocket unit which comprises a combustion chamber 12 the end-wall of which comprises a propellant injector. The end-wall comprises walls 17, 20, 21, 22, the walls 17, 20 defining an oxidizer chamber 19 and the walls 20, 21 defining a fuel chamber 23. The wall 22 is formed with pairs of divergent passages 27, 28 and 33, 34, oxidizer passing from chamber 19 through duct 26 and passage 25 to the divergent passages 27, 28 and fuel passing from the chamber 23 through passage 32 to the divergent passages 33, 34. Further passages 30 extend from the point of divergence of the passages 27, 28 and open into the combustion chamber 12 at points 31 located adjacent the points at which the divergent passages 27, 28 open into the chamber. Similarly, further passages 35 extend from the point of divergence of the passages 33, 34 and open into the combustion chamber at points 36 located adjacent the points at which the divergent passages 33, 34 open into the chamber. In the event of a pressure wave developing in the chamber, such pressure is sensed by a passage such as 31 whereby the flow of oxidizer is caused to switch from passage 27 say, to passage 28. Similarly, pressure waves sensed at a point 36 will cause the fuel supply to switch from passage 33 to passage 34. The fuel and oxidizer discharge orifices and associated pressure-sensing orifices are disposed in concentric annular patterns on the wall of the propellant injector. In Fig. 4, a fuel supply device is shown and comprises a supply passage 44 which extends to two divergent passages 45, 46 opening to a combustion space 47. Further passages 48, 51 extend from the point 49 at which the passages 45, 46 diverge, to points 50, 52 along the passages 45, 46. In this arrangement if fuel is being delivered along passage 45 for example, the pressure sensed at point 50 will be communicated through passage 48 to the point 49 whereupon the fuel supply will be switched to the passage 46. Thus fuel is caused to discharge alternately through the passages 45 and 46. This arrangement is shown applied to an I.C. engine in Fig. 5 (not shown). In Fig. 7 liquid fuel is supplied from a chamber 82 and air or steam is supplied from a chamber 81, both to the point of divergence of passages 84, 85. Flow through one passage 84 is sensed at point 88 and communicated through passage 87 to the point of divergence 83 whereupon flow is switched to the other passage 85. In a combustion chamber of a gas turbine engine, Fig. 11 fuel is supplied through a device 141 which causes the fuel to switch alternately from one divergent passage to the other, and air under pressure is supplied through similar devices 142. Secondary air is supplied through similar  devices 143 disposed in the wall of the flame tube. Other embodiments are described.