초록 ▼

An aircraft accessory system includes an aircraft engine powered direct air turbine driven accessory and an air turbine drivingly directly connected by an air turbine shaft to the accessory. The air turbine includes a variable geometry turbine nozzle in selectable direct flow communication with at l

An aircraft accessory system includes an aircraft engine powered direct air turbine driven accessory and an air turbine drivingly directly connected by an air turbine shaft to the accessory. The air turbine includes a variable geometry turbine nozzle in selectable direct flow communication with at least two compressed engine air sources. The two compressed engine air sources may be an HPC interstage bleed and an HPC compressor discharge stage bleed. The variable geometry turbine nozzle may be in selectable direct flow communication with a third compressed engine air source such as a bypass duct or an engine inlet duct. The air turbine includes a turbine exit which may be in selectable direct flow communication with at least two relatively lower pressure engine air sinks. The air sinks may be located in the aft end of a bypass duct and in a divergent section of the exhaust nozzle.

대표청구항 ▼

What is claimed is: 1. An aircraft accessory system includes: an aircraft engine powered direct air turbine driven accessory, an air turbine drivingly directly connected by an air turbine shaft to the accessory, the air turbine having a variable geometry turbine nozzle, and the variable geometry tu

What is claimed is: 1. An aircraft accessory system includes: an aircraft engine powered direct air turbine driven accessory, an air turbine drivingly directly connected by an air turbine shaft to the accessory, the air turbine having a variable geometry turbine nozzle, and the variable geometry turbine nozzle being in selectable direct flow communication with at least two compressed engine air sources of a single aircraft propulsive engine. 2. A system as claimed in claim 1 further comprising the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 3. An aircraft accessory system includes: an aircraft engine powered direct air turbine driven accessory, an air turbine driving directly connected by an air turbine shaft to the accessory, the air turbine having a variable geometry turbine nozzle, the variable geometry turbine nozzle being in selectable direct flow communication with at least two compressed engine air sources, the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed, and the variable geometry turbine nozzle being in selectable direct flow communication with a third compressed engine air source wherein the third compressed engine air source is a bypass duct or an engine inlet duct. 4. A system as claimed in claim 1 further comprising the air turbine including a turbine exit in selectable direct flow communication with at least two relatively lower pressure engine air sinks. 5. An aircraft accessory system includes: an aircraft engine powered direct air turbine driven accessory, an air turbine drivingly directly connected by an air turbine shaft to the accessory, the air turbine having a variable geometry turbine nozzle, the variable geometry turbine nozzle being in selectable direct flow communication with at least two compressed engine air sources, the air turbine including a turbine exit in selectable direct flow communication with at least two relatively lower pressure engine air sinks, and a first one of the two relatively lower pressure engine air sinks being located in the aft end of a bypass duct and a second one of the two relatively lower pressure engine air sinks being located in a divergent section of the exhaust nozzle. 6. A system as claimed in claim 5 further comprising the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 7. A system as claimed in claim 6 further comprising the variable geometry turbine nozzle being in selectable direct flow communication with a third compressed engine air source wherein the third compressed engine air source is a bypass duct or an engine inlet duct. 8. A system as claimed in claim 1 wherein the air turbine driven accessory is a constant voltage electrical power generator. 9. An aircraft accessory system includes: an aircraft engine powered direct constant voltage electrical power generator, an air turbine drivingly directly connected by an air turbine shaft to the constant voltage electrical power generator, the air turbine having a variable geometry turbine nozzle, the variable geometry turbine nozzle being in selectable direct flow communication with at least two compressed engine air sources, and the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 10. A system as claimed in claim 9 further comprising the variable geometry turbine nozzle being in selectable direct flow communication with a third compressed engine air source wherein the third compressed engine air source is a bypass duct or an engine inlet duct. 11. A system as claimed in claim 8 further comprising the air turbine including a turbine exit in selectable direct flow communication with at least two relatively lower pressure engine air sinks. 12. A system as claimed in claim 11 further comprising a first one of the two relatively lower pressure engine air sinks being located in the aft end of a bypass duct and a second one of the two relatively lower pressure engine air sinks being located in a divergent section of the exhaust nozzle. 13. A system as claimed in claim 12 further comprising the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 14. A system as claimed in claim 13 further comprising the variable geometry turbine nozzle being in selectable direct flow communication with a third compressed engine air source wherein the third compressed engine air source is a bypass duct or an engine inlet duct. 15. A system as claimed in claim 8 further comprising a constant voltage generator control system for controlling the constant voltage electrical power generator, the control system including: a rotational speed sensor positioned to measure turbine discharge rotor speed of the air turbine and output a rotor speed signal, a converter operable to filter and then convert the rotor speed signal from a frequency signal to an analog signal indicative of the turbine discharge rotor speed of the air turbine in RPM, a comparator operably connected to the converter for receiving the analog signal, and comparing the analog signal to an acceleration and deceleration schedule and speed limits, and calculating a resulting error signal, a torque motor drive operably connected to an air servo valve for powering a pneumatic actuator operably connected to and for adjusting vanes of the variable geometry turbine nozzle, and the comparator being operable to use the resulting error signal for effecting compensation and gain of the torque motor drive to adjust the vanes. 16. A system as claimed in claim 15 further comprising a rectifier for converting AC current from the constant voltage electrical power generator to DC current and matching the DC current to an aircraft electrical load. 17. A system as claimed in claim 1 wherein the air turbine driven accessory is a constant frequency electrical power generator. 18. An aircraft accessory system includes: an aircraft engine powered direct air turbine driven accessory, an air turbine drivingly directly connected by an air turbine shaft to the accessory, the air turbine having a variable geometry turbine nozzle, and the variable geometry turbine nozzle being in selectable direct flow communication with at least two compressed engine air sources, the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 19. A system as claimed in claim 18 further comprising the variable geometry turbine nozzle being in selectable direct flow communication with a third compressed engine air source wherein the third compressed engine air source is a bypass duct or an engine inlet duct. 20. A system as claimed in claim 17 further comprising the air turbine including a turbine exit in selectable direct flow communication with at least two relatively lower pressure engine air sinks. 21. A system as claimed in claim 20 further comprising a first one of the two relatively lower pressure engine air sinks being located in aft end of a bypass duct and a second one of the two relatively lower pressure engine air sinks being located in a divergent section of the exhaust nozzle. 22. A system as claimed in claim 21 further comprising the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 23. A system as claimed in claim 22 further comprising the variable geometry turbine nozzle being in selectable direct flow communication with a third compressed engine air source wherein the third compressed engine air source is a bypass duct or an engine inlet duct. 24. A system as claimed in claim 17 further comprising a constant frequency generator control system for controlling the constant frequency electrical power generator, the control system including: a rotational speed sensor positioned to measure turbine discharge rotor speed of the air turbine and output a rotor speed signal, a converter operable to filter and then convert the rotor speed signal from a frequency signal to an analog signal indicative of the turbine discharge rotor speed of the air turbine in RPM, a comparator operably connected to the converter for receiving the analog signal, and comparing the analog signal to an acceleration and deceleration schedule and a speed set point, and calculating a resulting error signal, a torque motor drive operably connected to an air servo valve for powering a pneumatic actuator operably connected to and for adjusting vanes of the variable geometry turbine nozzle, and the comparator being operable to use the resulting error signal for effecting compensation and gain of the torque motor drive to adjust the vanes. 25. A system as claimed in claim 24 further comprising a voltage regulator for matching AC current from the constant frequency electrical power generator to an aircraft electrical load. 26. A system as claimed in claim 1 wherein the air turbine driven accessory is a variable speed centrifugal fuel pump. 27. A system as claimed in claim 26 further comprising the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 28. A system as claimed in claim 27 further comprising the variable geometry turbine nozzle being in selectable direct flow communication with a third compressed engine air source wherein the third compressed engine air source is a bypass duct or an engine inlet duct. 29. A system as claimed in claim 26 further comprising the air turbine including a turbine exit in selectable direct flow communication with at least two relatively lower pressure engine air sinks. 30. A system as claimed in claim 29 further comprising a first one of the two relatively lower pressure engine air sinks being located in aft end of a bypass duct and a second one of the two relatively lower pressure engine air sinks being located in a divergent section of the exhaust nozzle. 31. A system as claimed in claim 30 further comprising the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 32. A system as claimed in claim 31 further comprising the variable geometry turbine nozzle being in selectable direct flow communication with a third compressed engine air source wherein the third compressed engine air source is a bypass duct or an engine inlet duct. 33. A system as claimed in claim 26 further comprising a variable speed centrifugal fuel pump control system for controlling a pump speed of the fuel pump to maintain a constant pressure decrease across a fuel metering valve which is fluid flow receiving communication with the fuel pump, the control system including: a rotational speed sensor positioned to measure turbine discharge rotor speed of the air turbine and output a rotor speed signal, a converter operable to filter and then convert the rotor speed signal from a frequency signal to an analog signal indicative of the turbine discharge rotor speed of the air turbine in RPM, a comparator operably connected to the converter for receiving the analog signal, and comparing the analog signal to an acceleration and deceleration schedule and a pressure decrease set point across the fuel metering valve, and calculating a resulting error signal, a torque motor drive operably connected to an air servo valve for powering a pneumatic actuator operably connected to and for adjusting vanes of the variable geometry turbine nozzle, and the comparator being operable to use the resulting error signal for effecting compensation and gain of the torque motor drive to adjust the vanes. 34. An aircraft ramjet engine comprising: in downstream serial fluid communication an annular engine inlet duct, fan duct circumscribing a fan section, a core engine, a low pressure turbine, and an exhaust duct, a bypass duct extending downstream from at least a portion of the fan section around the core engine and the low pressure turbine to an exhaust duct downstream of and in fluid communication with both the core engine and the bypass duct, ram burners operatively disposed in the engine and capable of operating the engine in a ramjet mode, an aircraft engine powered direct air turbine driven accessory, an air turbine drivingly directly connected by an air turbine shaft to the accessory, the air turbine having a variable geometry turbine nozzle, and the variable geometry turbine nozzle being in selectable direct flow communication with at least two compressed engine air sources of the engine. 35. An engine as claimed in claim 34 further comprising the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 36. An aircraft ramjet engine comprising: in downstream serial fluid communication an annular engine inlet duct, fan duct circumscribing a fan section, a core engine, a low pressure turbine, and an exhaust duct, a bypass duct extending downstream from at least a portion of the fan section around the core engine and the low pressure turbine to an exhaust duct downstream of and in fluid communication with both the core engine and the bypass duct, ram burners operatively disposed in the engine and capable of operating the engine in a ramjet mode, an aircraft engine powered direct air turbine driven accessory, an air turbine drivingly directly connected by an air turbine shaft to the accessory, the air turbine having a variable geometry turbine nozzle, the variable geometry turbine nozzle being in selectable direct flow communication with at least two compressed engine air sources, the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed, and the variable geometry turbine nozzle being in selectable direct flow communication with a third compressed engine air source wherein the third compressed engine air source is a bypass duct or the inlet duct. 37. An engine as claimed in claim 34 further comprising the air turbine including a turbine exit in selectable direct flow communication with at least two relatively lower pressure engine air sinks. 38. An aircraft ramjet engine comprising: in downstream serial fluid communication an annular engine inlet duct, fan duct circumscribing a fan section, a core engine, a low pressure turbine, and an exhaust duct, a bypass duct extending downstream from at least a portion of the fan section around the core engine and the low pressure turbine to an exhaust duct downstream of and in fluid communication with both the core engine and the bypass duct, ram burners operatively disposed in the engine and capable of operating the engine in a ramjet mode, an aircraft engine powered direct air turbine driven accessory, an air turbine drivingly directly connected by an air turbine shaft to the accessory, the air turbine having a variable geometry turbine nozzle, the variable geometry turbine nozzle being in selectable direct flow communication with at least two compressed engine air sources, the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed, the air turbine including a turbine exit in selectable direct flow communication with at least two relatively lower pressure engine air sinks, and a first one of the two relatively lower pressure engine air sinks being located in aft end of a bypass duct and a second one of the two relatively lower pressure engine air sinks being located in a divergent section of the exhaust nozzle. 39. An engine as claimed in claim 38 further comprising the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 40. An engine as claimed in claim 39 further comprising the variable geometry turbine nozzle being in selectable direct flow communication with a third compressed engine air source wherein the third compressed engine air source is a bypass duct or the inlet duct. 41. A bypass turbofan engine comprising: in downstream serial fluid communication a fan duct circumscribing a fan section, a core engine, a low pressure turbine, a bypass duct extending downstream from at least a portion of the fan section and circumscribing at least a part of the core engine, an aircraft engine powered direct air turbine driven accessory, an air turbine drivingly directly connected by an air turbine shaft to the accessory, the air turbine having a variable geometry turbine nozzle, and the variable geometry turbine nozzle being in selectable direct flow communication with at least two compressed engine air sources of the engine. 42. An engine as claimed in claim 41 further comprising the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed. 43. An engine as claimed in claim 42 further comprising the air turbine including a turbine exit in selectable direct flow communication with at least one relatively lower pressure engine air sinks. 44. A bypass turbofan engine comprising: in downstream serial fluid communication a fan duct circumscribing a fan section, a core engine, a low pressure turbine, a bypass duct extending downstream from at least a portion of the fan section and circumscribing at least a part of the core engine, an aircraft engine powered direct air turbine driven accessory, an air turbine drivingly directly connected by an air turbine shaft to the accessory, the air turbine having a variable geometry turbine nozzle, the variable geometry turbine nozzle being in selectable direct flow communication with at least two compressed engine air sources, the two compressed engine air sources being an HPC interstage bleed and an HPC compressor discharge stage bleed, the air turbine including a turbine exit in selectable direct flow communication with at least one relatively lower pressure engine air sinks, and the relatively lower pressure engine air sinks being located in aft end of a bypass duct and a second one of the two relatively lower pressure engine air sinks being located in a divergent section of the exhaust nozzle.