Showing papers on "Afterburner published in 1983"

System for controlling the thrust nozzle adjustment of dual cycle gas turbine jet propulsion engines

Abstract: The thrust nozzle opening angle of two cycle gas turbine jet engines is crolled in a system which comprises a low pressure compressor driven by a low pressure turbine. A first compressor is arranged radially inward for a first flow cycle or circuit and a second compressor is arranged radially outward for a second flow cycle or circuit. A high pressure compressor is driven by a high pressure turbine. A combustion chamber is situated upstream of the high pressure turbine. An afterburner is supplied with the turbine gases from the first, hot flow cycle and with compressed air from the second, relatively cool flow cycle or circuit. The system may have multiple shafts. The control is effected with reference to the instantaneous pressure relationship (π NVZ ) between the static pressure (p sD ) downstream of the radially outer compressor (1a) of the second flow cycle (K2) and the static pressure (p sE ) of the air flowing into the propulsion system. This ratio constitutes a variable rated control value (π NVZ-Soll ) and is calculated as a function of the reduced rpm (N r ) of the low pressure compressor (1). This type of control guaranties during a steady state operation particularly a minimal fuel consumption in the subsonic region and a maximum thrust in the supersonic region. This type of control further assures during a nonsteady type of operating, particularly the stability between the two aerodynamic limits, compressor surges (G1 B or G1 H ) on the one hand and afterburner quenching (G2 B or G2 H ) on the other hand.

System for controlling the thrust nozzle adjustment of dual cycle gas turbine jet propulsion engines

Abstract: The thrust nozzle opening angle of two cycle gas turbine jet engines is controlled in a system which comprises a low pressure compressor driven by a low pressure turbine. A first compressor is arranged radially inward for a first flow cycle or circuit and a second compressor is arranged radially outward for a second flow cycle or circuit. A high pressure compressor is driven by a high pressure turbine. A combustion chamber is situated upstream of the high pressure turbine. An afterburner is supplied with the turbine gases from the first, hot flow cycle and with compressed air from the second, relatively cool flow cycle or circuit. The system may have multiple shafts. The control is effected with reference to the instantaneous pressure relationship (πNVZ) between the static pressure (p sD ) downstream of the radially outer compressor (1a) of the second flow cycle (k2) and the static pressure (p sE ) of the air flowing into the propulsion system. This ratio constitutes a variable rated control value (π NVZ-Soll ) and is calculated as a function of the reduced rpm (N r ) of the low pressure compressor (1). This type of control guarantees during a steady state operation particularly a minimal fuel consumption in the subsonic region and a maximum thrust in the supersonic region. This type of control further assures during a nonsteady type of operating, particularly the stability between the two aerodynamic limits, compressor surges (G1 B or G1 H ) on the one hand and afterburner quenching (G2 B or G2 H ) on the other hand.

Fuel spray device for gas turbine augmentor or afterburner

Abstract: The fuel spray device is adapted for use in the thrust augmentor or afterburner of a gas turbine engine. The fuel spray device includes a fuel spray manifold having a plurality of spaced apart ejection orifices and a plurality of nozzle pintles each having a decreasing cross-sectional portion, such as a conical portion, against which the orifice wall seats when there is no fuel pressure and relative to which the orifice wall moves as fuel pressure increases to increase fuel flow rate and each pintle having a substantially constant cross-sectional portion, such as a cylindrical portion, downstream of the conical portion to accommodate movement of the orifice wall of the manifold with a minimal effect on fuel flow rate through the orifice.

Model test and full-scale checkout of dry-cooled jet runup sound suppressors

Abstract: Jet aircraft runup sound suppressors featuring complete enclosure of the aircraft, dry cooling of the exhaust sound suppressor (even during afterburner operation), and adaptability to a variety of aircraft types are now employed at a number of U.S. Navy and Air Force airfields. The design, model testing, and full-scale checkout of the existing U.S. Navy hush house is described herein. Also the extension of the hush house concept to unusual aircraft designs and to dry-cooled jet engine test cells is covered.

Design study of a dual-cycle turbofan-ramjet engine for a hypersonic aircraft

Abstract: Computer modelling was used with two different designs of an advanced turbofan-ramjet in order to derive performance predictions. The engine would enable an aircraft to take-off, accelerate to Mach 5.0, and climb to 90,000 ft. The two concepts included a turbofan with a ramjet annularly wrapped around it and a side-by-side configuration with the ramjet having a rectangular shape and mounted alongside the turbofan. The studies were performed to model weight, length, fuel efficiency, and the requirements of the thrust/drag ratio to exceed unity over the entire flight path. LH2 would be used for fuel and to regeneratively cool the combustion chamber. Turbofan operation with and without afterburner and with and without the ramjet inlet open were examined, as were variable areas for the burners. A side-by-side configuration displayed the best performance predictions, with a ramjet mass flow being 75 percent that of the turbofan and maximum temperatures being equal.