Showing papers on "Turbofan published in 1970"

Fan engine with counter rotating geared core booster

Abstract: The disclosure describes a turbofan engine having high and low pressure compressors wherein the low pressure compressor, or booster, comprises two counter rotating elements, a rotating duct which carries fan blades on its exterior surface and compressor blades on its interior surface and a conventional compressor rotor which rotates in a direction counter to the direction of rotation of the rotating duct. Alternative gearing schemes for coupling the rotation of the fan to the rotation of the low pressure compressor are shown.

Shortened afterburner construction for turbine engine

Abstract: An afterburner construction for a turbine engine, such as a turbofan engine, which is foreshortened by using a construction which utilizes swirl flow phenomena to rapidly mix the engine products of combustion and coolant flow, such as fan air, and/or to rapidly accomplish the afterburning combustion process in the afterburner, while maintaining engine performance and structural part integrity.

Gas turbine engine mixer

Abstract: An afterburning turbofan engine includes amixer for mixing the engine generated hot gas and fan streams prior to discharge through the engine nozzle. The mixer includes a fixed upstream portion and a movable downstream portion which may be indexed to selectively interchange flow communication between the chutes of the two portions to effect a temperature reduction in the downstream mixer portion and, hence, a reduction in infrared radiation.

Ducted fan engine exhaust mixer

Abstract: A turbofan engine has an exhaust arrangement for the bypassed air and turbine exhaust gases providing for mixing of the two before a common propulsion nozzle. The turbine exhaust gas is directed to a number of circumferentially spaced exhausts, the outlets of which are inclined to the axis of the engine. A baffle extending from the end of the bypass duct into the common exhaust duct divides the bypass air into a portion inside and a portion outside of the baffle, the former mixing initially with the turbine exhaust gases and the resulting mixture then being mixed with the remainder of the bypass air. The baffle includes lobes which extend downstream and inwardly so as to obscure the outlets of the turbine exhaust from the jet nozzle, thus minimizing thermal radiation through the nozzle.

Turbofan-powered stol aircraft

Abstract: A short takeoff and landing (STOL) aircraft employing turbofan engines and a wing and flap arrangement to provide propulsive lift, a high lift capability, and a steeper landing approach flight path. Improvements in these characteristics are achieved by a particular positioning of the engine fan and primary exhaust efflux flow relative to the wing and by locating the flaps for operation in consort with the engine-derived gaseous flow and the wing.

Noise reduction in jet engines having fans or low pressure compressors

Abstract: A TURBOFAN JET ENGINE EMPLOYED IN THE PROPULSION OF AIRCRAFT. THE AIR INLET TO A TYPICAL FAN-JET ENGINE IS DEFINED BY THE COWL AND COWL LIP RING WHICH DIRECTS THE INCOMING AIR TO AND THROUGH THE FAN, OR LOW PRESSURE COMPRESSOR, HAVING ROTOR AND STATOR BLADES. THE PRIMARY FLOW IS DIRECTED INTO THE COMBUSTION CHAMBER WHERE HOT GASES RESLUTING FROM COMBUSTION ARE USED TO DRIVE THE FAN OR LOW PRESSURE COMPRESSOR. THE SECONDARY OUTER PORTION OF THE AIR PRESSURIZED BY THE FAN IS DIRECTED REARWARD AROUND THE OUTER EDGE OF THE INNER ENGINE DEFINED BY THE INNER COWL AND THE OUTER WALL ADJACENT THE COMBUSTION CHAMBER AND DIRECTED TO PROVIDE A PROPULSIVE FORCE FROM THE ENGINE. AT LEAST TWO ANNULAR SPLITTER RINGS CONCENTRICALLY SPACED ARE PROVIDED WITHIN THE COWL FORWARD OF THE FAN OR LOW PRESSURE COMPRESSOR.

The effect of inlet temperature distortion on the performance of a turbo-fan engine compressor system

Abstract: Turbofan engine compressor system performance dependence on circumferential extent, magnitude and rate of change of inlet temperature in altitude test facility

Performance and stall limits of an afterburner-equipped turbofan engine with and without inlet flow distortion

Abstract: Performance and stall limits of afterburner-type turbofan engine with and without inlet flow distortion

Improvements relating to Cowls for Turbo-Fan Aircraft Engines.

Abstract: 1,195,027. Gas turbine jet propulsion plant; axial flow fans. BRITISH AIRCRAFT CORP. Ltd. Feb. 5, 1968 [Nov. 8, 1966], No. 50031/66. Headings F1C and F1J. A gas turbine ducted fan aircraft engine has a fan cowl 23 formed by inlet and exhaust sections 25, 26, respectively, both fabricated from radially spaced skins of light-gauge sheet material and joined to each other by an intermediate section 27 consisting of a single skin surrounding the periphery of the fan 18 and of sufficient thickness to contain accidentally detached fan parts. The outer and inner skins 40, 42, Fig. 3, of the inlet section 25 are braced by longitudinally extending ribs 44 having holes 48 to receive securing bolts for lugs 46 of the intermediate section 27. Similar ribs 45 brace the skins 41, 43 of the outlet section 26 and have holes 49 to receive securing bolts for further lugs 47 of section 27. Straps 50, 51 cover the access openings to the bolts. The single skin of section 27 may be formed of a single thickness of material or of a sandwich of two or more materials in surface contact. As shown in Fig. 1 the cowl 23 is applied to a three-spool type aft fan engine and is supported from the engine by radial struts 24, 28. Compared with a conventional fan cowl, shown dotted at 30, the cowl 23 is shorter and has a smaller external diameter, thus reducing drag. The cowl may also be applied to a front fan engine, Fig. 2 (not shown).

The s/tol aircraft

Abstract: The invention concerns a turbo fan engine and in which the output of the fan can be normally directed downwardly to lift the aircraft from the ground, or optionally, can be caused to shift into a horizontal direction and thus augment the normal thrust of the engine. It has particular application to the so-called ''''vertical short takeoff and landing'''' type of aircraft. The fan is located forward of the engine and the movement of the air which is in excess of that furnished to the compressor in changed as to direction from the vertical to the horizontal by causing a jet of high pressure air, obtained from a source exterior of the fan, bodily to shift or divert the air column emanating from the fan traveling in the downward vertical direction to now travel in the horizontal direction after the plane has attained its proper height in space. The diverting jet of high pressure air is discharged from an annular opening in a chamber of ring-like character which surrounds the engine and to which the source of high pressure air, preferably taken from the compressor, is applied.

Determination of the Effect of Atmospheric Humidity on the Characteristics of a Turbofan Engine

Abstract: : A mathematical method is presented for the study of the humidity effect on parameters of any gas-turbine engine.

High Temperature Turbine Design Considerations: A discussion of the major technological developments which have led to the trend of using higher temperatures in modern gas turbine aero engines from a paper first presented at an AGARD conference in September

Abstract: THE effect of higher turbine inlet temperature on the performance of an aircraft gas turbine engine can be quite dramatic. It can be used to increase the exhaust velocity of a dry turbojet to providea higher specific thrust; to increase the bypass ratio of a turbofan engine to improve its propulsive efficiency; to optimize the thermodynamic cycle at a higher pressure ratio to improve its specific fuel consumption; to reduce the amount of afterburner fuel flow in an augmented turbojet to improve its specific fuel consumption, or to increase the work output of a turboshaft engine. If the thrust or power of the engine is held constant, a size, cost and/or weight reduction can result. If the size of the engine is held constant growth capability can be provided.

Convertible Fan/Shaft Engine Variable Fan Geometry Investigation

Abstract: : Preliminary design studies have shown that advanced VTOL aircraft performance can be enhanced when the aircraft are coupled with a turbofan engine having the ability to provide shaft power to the vertical lift rotor (convertible fan/shaft engine). A program was conducted to determine the feasibility of using variable fan inlet and duct exit guide vanes to effectively reduce fan horsepower and eliminate thrust during periods of vertical lift and hover. An exploratory test program was conducted in an existing 0.5 hub-tip ratio compressor research rig modified to simulate a high-bypass-ratio fan configuration. This fan configuration consisted of variable inlet and exit guide vanes, a fan rotor, and a simulated gas generator flowpath providing a 4.2:1 bypass ratio. The fan rotor was an existing moderate-speed rotor blade with a tip diameter of 43 inches, tip speed of 1150 fps, and design pressure ratio and flow of 1.35 and 285 lb/sec, respectively. Results from the test program established the feasibility of using variable-geometry inlet and duct exit guide vanes to effectively reduce fan horsepower and thrust. Fan horsepower was reduced to 16 percent of the maximum (cruise) power absorbed by the fan, and fan thrust was reduced by 100 percent. The test results indicate that the potential exists for reducing fan horsepower to values of less than 10 percent of the cruise power. Inlet guide vane, rotor blade, and duct exit guide vane stresses were within safe operating limits over the test range of guide vane positions.

Improvements in gas turbine engines

Abstract: 1,211,064. Gas turbine ducted fan engines. GENERAL ELECTRIC CO. March 5, 1969, No.11822/69. Heading F1J. A turbofan engine comprises a core engine 10 having an axial flow compressor 18 downstream of which is a combustor 24, the hot gas stream from which drives a turbine 26 connected by a shaft 28 to the compressor rotor 20. The overall engine compression ratio, i.e. the pressure of air discharged from the compressor 18 to the pressure of air entering the turbofan engine is at least 20:1. A second turbine 14 is driven by this same hot gas stream and is connected by a shaft 16 to a fan 12 rotating within an annular duct. The hot gas stream is discharged from a nozzle 30 defined in part by a plug 32, and the fan air from a convergent nozzle 29 to provide a propulsive force. The by-pass ratio of the mass of air discharged by the fan 12 into the nozzle 29 to the mass of air entering the core engine 10 is between 5:1 and 10:1, which parameter in combination with the compression ratio of 20:1 or greater is effective in reducing specific fuel consumption. The temperature of the gas stream leaving the combustor 24 and entering the turbine 26 is at least 2000‹F, operation at which temperature minimizes the size and weight of an engine for a given thrust level with little or no increase in specific fuel consumption. The lighter weight and reduced aerodynamic drag of a smaller engine in turn result in lower rates of fuel consumption in propulsion of an aircraft. In order to cool the turbine, cooling air may be derived from the compressor without any substantial increase in specific fuel consumption.

Calculation of the Flow Field Downstream of the Fan Nozzle of a Turbofan Aero Engine

Abstract: The pursuit of better economics for subsonic transport aircraft has led in recent years to the development of high by-pass ratio turbofan aero engines. The improved efficiency of these engines results mainly from an improvement in the propulsive, or Froude efficiency brought about by the use of an increased airflow with a smaller jet velocity for a given thrust level. The specific thrust (thrust per unit airflow) for a high by-pass ratio engine is thus appreciably lower than for a conventional turbojet so that the ratio of engine gross thrust to installed net thrust is higher. Typically, for an engine of unit by-pass ratio, the ratio of gross to net thrust is about 1·6 under cruise conditions, while for an engine of by-pass ratio five it is between 2·6 and 3. Losses in the exhaust system effectively act on the gross thrust level so that any inefficiency is magnified by the gross to net thrust ratio. Thus the high values of this ratio in the new high by-pass engines render losses in the exhaust system of especial significance.

Effect of simulated downstream flow blockage doors on the performance of an axial-flow fan rotor

Abstract: Simulating downstream flow blockage doors in bypass air duct of turbofan engine with axial flow fan rotor