The T31 engine has a single-stage turbine that delivered about 5000 horsepower at stsndard sea-level conditions and an engine speed of 13,000 rpm. The T-31 engine was the first American turboprop engine to power an aircraft. It made its initial flight in the Consolidated Vultee XP-81 on December 21, 1945. The T-31 was mounted in the nose; a J-33 turbojet engine mounted in the rear fuselage provided added thrust. The T-31 was also used on the Navy XF2R-1 Dark Shark, similarly powered by a turboprop/turbojet engine combination. The engine was to have been flown experimentally on a Curtiss XC-113 (a converted C-46) but the experiment was abandoned after the XC-113 was involved in a ground accident. Only 28 T-31s were built, and none were used in production aircraft, but improved production turboprop engines were developed from the technology pioneered by the T-31.
The T31 engine is equipped with a 14-stage axial-flow compressor, which hae a sea-level air-flow rating of about 21 pounds per second at an engine speed of 13,000 rpm. The campressor rotor consists of 14 wheels shrunk on a hollow shaft. The compressor blades are dovetailed into the rotor wheels. The blade-tip diameter of the rotor is 1% inches and the over-all length is 25 inches. The hub-to-tip diameter ratio at the first rotor stage is 0.73 and increases to 0.88 at the fourteenth rotor stage. A balance pressure is applied to the forward end of the rotor by air bled from the fifth stage of the compressor. This air leaks out through two labyrinth seals into the compressor air passage aft of the inlet guide vanes. Another labyrinth seal is located at the aft end of the rotor. Air leaking through this seal is used to cool the forward face of the turbine wheel.
The Stator stages consist of half rings Into which the stator blades are dovetailed. These half rings, assermbled around the rotor with spacers and clamping bolta, compose the compressor stator assembly. Air enters the compressor through an annular inlet, which is divided into six equal segments by radial support struts. The flow area of the compressor inlet is approximately 95 square incbes. A single row of guide vanes turns the air in the direction of rotation of the rotor. Air is discharged from the compressor through two rows of straightening vanes into as annular exhaust.
The turbine has a solid steel disk that tapers in thickness from 3.70 inches at the hub to 0.57 inch at the thinnest section near the rim. The turbine blades, which are welded to the wheel rim, are 1.6 inches in length. The blade chord t8pers from 1.0 fnoh at the root to 0.75 inch at the tip. The blade forgings are so designed that the rectangular tips in the assembled wheel form the turbine shroud ring. The over-all diameter of the wheel including the shroud ring is 28 inches.
The turbine nozzle, which consists of 36 equally spaced hollow steel vanes, has an 8CtIl81 flow area of about 25 square inches 8nd an elcpansion ratio of 1.065. The vsnes are welded to inner and outer shroud rings. A portion of the air that enters the combustion chambers first flows through the hollow vages to provide cooling. Gases discharged from the turbine enter an 8nnular exhaust cone hsving an area of about 154 square inches at the location of the turbine-outlet. The inner cone is supported by four struts and by a series of small angle brsces extending along the entire length of the inner cone.
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