Supersonic jet engine and method of operating the same

A supersonic jet engine adapted to cruise at supersonic speeds at a relatively high efficiency, to operate with turbine inlet temperatures of at least 2500.degree. F. for high power operation, and to cruise subsonically with a relatively low specific fuel consumption, said engine comprising:

• a. a housing structure having an upstream inlet end to receive intake air, and a downstream exhaust end to discharge jet exhaust,
• b. an air inlet arranged to receive intake air at subsonic velocity and also to receive intake air at supersonic velocities and reduce said air to subsonic velocity,
• c. a compressor mounted in said housing rearwardly of said inlet and arranged to compress air flowing into said inlet, said compressor having an inlet end and an outlet end arranged to operate between a maximum compression ratio and a minimum compression ratio,
• d. means defining a combustion chamber mounted in said housing downstream of said compressor to receive compressed air therefrom,
• e. fuel injection and ignition means in said combustion chamber arranged to burn fuel in air flow from the compressor and provide a gaseous flow from said combustion chamber,
• f. means defining a first generally annular passageway downstream of said combustion chamber to receive gaseous flow from said combustion chamber,
• g. means defining second and third generally coannular passageways, said second passageway being positioned radially inwardly of said third passageway and arranged to receive a first portion of gaseous flow from said first passageway, said third passageway arranged to receive a second portion of gaseous flow from said first passageway,
• h. a turbine first stage positioned in said first passageway and arranged to receive the gaseous flow from said combustion chamber so as to be driven thereby with said gaseous flow proceeding to said second and third passageways, said turbine first stage being arranged to receive said gaseous flow at temperatures at least as high as 2500.degree. F. and also at lower temperatures, said turbine first stage having a predetermined cross-sectional nozzle area,
• i. a turbine second stage positioned in said second passageway and arranged to receive said first portion of gaseous flow so as to be driven thereby,
• j. nozzle means positioned downstream of said second and third passageways to receive the first portion of gaseous flow and discharge said first portion at a radially inward location at a relatively low velocity, and to receive the second portion of gaseous flow and discharge said second portion at a relatively high velocity in a generally annular pattern radially outward of the radially inward location of the first portion of gaseous flow,
• k. turbine bypass means to receive flow from a location downstream of said compressor as bypass flow, and to direct said bypass flow along a path bypassing said turbine first stage, and exhaust said bypass flow from said engine to produce a thrust,
• l. bypass valve means to control the amount of flow bypassed into said turbine bypass means,
• m. engine control means operatively connected to said fuel injection and ignition means and to said bypass valve means, in a manner to control the amount of fuel directed to said fuel injection and ignition means and to control said bypass valve means in a manner to control the amount of bypass flow through said turbine bypass means,
• n. said turbine being matched to said compressor in a manner that with said engine operating at subsonic cruise velocity, with said compressor operating at maximum compression ratio, with said bypass valve means positioned so that there is substantially no flow through said turbine bypass means, and with fuel flow being adequate to create thrust to match airplane drag at said subsonic cruise velocity, the tubine first stage has a flow area sized to allow gaseous flow therethrough at the speed of sound in said gaseous flow, o. said engine control means being arranged to set said bypass valve means at a more open position at higher engine thrust settings where higher temperatures are created in said combustion chamber, and to set said bypass valve means at a more closed position at lower engine thrust settings where lower temperatures are created in said combustion chamber, in a manner that there is substantially constant corrected gaseous flow into said turbine first stage, said corrected flow being measured according to the formula:
• ##EQU4## where:
• W=Total mass flow rate in lbs. per second
• .theta.+=Observed temperature (absolute) divider by standard temperature (518.67.degree.R)
• .delta.+=Observed pressure divided by standard pressure (2116.22 lbs./sq.ft.)
• whereby during subsonic cruise said engine can operate at a relatively low specific fuel consumption with high compression ratio and low combustion exit temperature, in a condition where said valve bypass means is at a substantially closed position, during high power acceleration mode said engine can operate at high combustion exit temperatures in a condition where said valve bypass means is in a more open position to maintain constant corrected flow through said turbine first stage to satisfy the turbine first stage requirements, and during supersonic cruise mode the engine can operate efficiently at relatively high combustion exit temperatures, and said engine discharges its jet exhaust in a pattern to alleviate noise generated by said engine.