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Aviation History
1948
1948 - 1264.PDF
FLIGHT August izttt, 1948 These details show at left, the method of compressor mounting ; centre, the attachment of blading ; and, right, the annular coupling in the main drive. Napier Naiad. Five combustion chambers are employed on the Naiad, the cans being simple cylinders fabricated in Nimonic 75 and enclosing fabricated flametubes of quite conventional form, located concentrically by three radial spigots and with spring lugs at the outlet mouth for accommodating linear expansion. A novel feature of this part of the unit is that the combustion chambers are interchangeable and their fixture is such as readily to permit the removal of individual chambers if required. Delivery of air from the compressor is carried through the main support plate by means of five transfer passages, the exit apertures of which match with orifices in the base of each of the individual cast cup-units which form the crown of each combustion chamber. It is in these crown cups that the blow-off valves are embodied. Fuel is fed through a manifold ring surrounding the compressor at the tenth-stage station, and having pigtail branches to each of the burners. These last are of conventional downstream-discharge pattern, the nozzles being surrounded with swirl vanes to impart helical motion to the primary airflow; secondary air is admitted through holes in the fiametube wall, both at the waist and at the skirt. Igniter plugs are furnished in two combustion chambers only, flame propagation to the remaining chambers being facili- tated by connecting bridge tubes. The combustion chambers discharge through passages formed in the turbine feed manifold, the passage form being such that, from the five inlet apertures, the gases are spread so as to enter the turbine nozzle ring with an equal cir- cumferential, distribution. Design of the turbine proper is in- teresting. The first-stage disc is machined from Jessop's G18B steel, which has characteristics not unlike Nimonic 80, and is flange-bolted to the tail of the turbine shaft, a 3375m dia- meter forging of Jessop's R20 steel; this material has a similar expansion coefficient to G18B, but is consider- ably easier to machine. The second- stage disc is splined to the tail shaft of the first-stage wheel and is machined from Jessop's H40 steel. Blades in each of the turbine stages are Nimonic 80 forgings secured is the discs with fir-tree roots of conventional form. Rear support for the turbine assembly is provided by a ball thrust bearing housed in the hub of the rear turbine casing. The latter is a four-spoked assembly which is through-bolted in conjunction with the second-stage nozzle-ring to the periphery of the turbine feed manifold. The first-stage nozzle-ring—a Firth-Vickers HR Crown Max casting—is recessed into the delivery annulus of the turbine feed manifold, being clamped into that position by the peripheral shrouding-ring surrounding the first-stage turbine blades. The second-stage nozzle-ring is of materi- ally similar construction to the first-stage ring, but is rim- welded to the second-stage turbine tip-shroud, and also serves to carry the director shroud nested between the tur- bine discs for the purpose of directing the cooling airflow over the adjacent disc faces. The front face of the first- COOLING AS? DIRECTOR SHROUD AIR FROM COMPRESSOR 7tK STAGE THRUST BALANCE CHAPRAGM
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