CA1233325A - Counter rotation power turbine - Google Patents
Counter rotation power turbineInfo
- Publication number
- CA1233325A CA1233325A CA000438676A CA438676A CA1233325A CA 1233325 A CA1233325 A CA 1233325A CA 000438676 A CA000438676 A CA 000438676A CA 438676 A CA438676 A CA 438676A CA 1233325 A CA1233325 A CA 1233325A
- Authority
- CA
- Canada
- Prior art keywords
- turbine
- rotor
- blade rows
- power
- engine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/072—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with counter-rotating, e.g. fan rotors
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C11/00—Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
- B64C11/30—Blade pitch-changing mechanisms
- B64C11/306—Blade pitch-changing mechanisms specially adapted for contrarotating propellers
- B64C11/308—Blade pitch-changing mechanisms specially adapted for contrarotating propellers automatic
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/06—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
- F02C3/067—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages having counter-rotating rotors
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENTS OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plant in aircraft; Aircraft characterised thereby
- B64D2027/005—Aircraft with an unducted turbofan comprising contra-rotating rotors, e.g. contra-rotating open rotors [CROR]
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Abstract
COUNTER ROTATION POWER TURBINE
ABSTACT OF THE DISCLOSURE
A new and improved gas turbine engine including a gas generator and a power turbine is disclosed. The power turbine includes a first rotor having a plurality of first turbine blade rows extending radially outwardly therefrom, and a second rotor having a plurality of second turbine blade rows extending radially inwardly therefrom.
The power turbine is supported aft of the gas generator and is effective for receiving combustion gases therefrom and expanding the gases through the first and second turbine blade rows for extracting substantially all output power therefrom for driving the first and second rotors in counterrotating directions.
According to several embodiments of the invention, the power turbine is effective for driving counterrotating fans or propellers disposed either at a forward end or at an aft end of the engine.
ABSTACT OF THE DISCLOSURE
A new and improved gas turbine engine including a gas generator and a power turbine is disclosed. The power turbine includes a first rotor having a plurality of first turbine blade rows extending radially outwardly therefrom, and a second rotor having a plurality of second turbine blade rows extending radially inwardly therefrom.
The power turbine is supported aft of the gas generator and is effective for receiving combustion gases therefrom and expanding the gases through the first and second turbine blade rows for extracting substantially all output power therefrom for driving the first and second rotors in counterrotating directions.
According to several embodiments of the invention, the power turbine is effective for driving counterrotating fans or propellers disposed either at a forward end or at an aft end of the engine.
Description
~3313~2~
- l - 13Dv 8253 COUNTER ROTATION POWER TURBINE
-FIEND OF THE INVENTION
This invention relates to gas turbine engines and, more particularly, to a new and improved gas turbine engine including a power turbine having counter-rotating rotors effective for providing output shaft power at relatively low speeds.
BACKGROUND OF THE INVENTION
While not limited thereto the present invention is particularly applicable -to gas turbine engines such as used for the propulsion of aircraft.
Several types of gas turbine engines are currently available for powering aircraft. The turbofan and the turboprop are two examples of such engines. The turbofan engine includes a core engine, i.e., gas generator, to drive a fan, whereas the turboprop engine includes a gas generator and power turbine which drives a propeller. Inasmuch as these engines drive propellers or fans for generating thrust they are -typically more fuel efficient at subsonic speeds than pure turbojet engines which generate thrust only through their exhaust jets.
Intermediate-sized transport aircraft, for example, 100 -to 180 passenger transports, typically utilize turbofan engines for propulsion. Turbofans provide the relatively high thrust required for powering these aircraft at relatively high altitudes and at cruise speeds of about Mach 0.6 to about Mach 0.8. For aircraft
- l - 13Dv 8253 COUNTER ROTATION POWER TURBINE
-FIEND OF THE INVENTION
This invention relates to gas turbine engines and, more particularly, to a new and improved gas turbine engine including a power turbine having counter-rotating rotors effective for providing output shaft power at relatively low speeds.
BACKGROUND OF THE INVENTION
While not limited thereto the present invention is particularly applicable -to gas turbine engines such as used for the propulsion of aircraft.
Several types of gas turbine engines are currently available for powering aircraft. The turbofan and the turboprop are two examples of such engines. The turbofan engine includes a core engine, i.e., gas generator, to drive a fan, whereas the turboprop engine includes a gas generator and power turbine which drives a propeller. Inasmuch as these engines drive propellers or fans for generating thrust they are -typically more fuel efficient at subsonic speeds than pure turbojet engines which generate thrust only through their exhaust jets.
Intermediate-sized transport aircraft, for example, 100 -to 180 passenger transports, typically utilize turbofan engines for propulsion. Turbofans provide the relatively high thrust required for powering these aircraft at relatively high altitudes and at cruise speeds of about Mach 0.6 to about Mach 0.8. For aircraft
- 2 - 13DV g253 designed for lower cruise speeds, conventional turboprop are -typically used inasmuch as they can provide superior performance and efficiency. For example, significant reductions inflow burn, i.e., the amount of fuel consumed S per passenger mile, are possible from the use of -the aerodynamically more efficient -turboprop over the turbofan.
Accordingly, it would be desirable to combine the advantages of the turbofan and the turboprop for obtaining a compound engine having improved overall engine efficiency at aircraft cruise speeds typical of turbofan powered aircraft.
However, a simple scaled up version of a convent tonal turboprop engine suitable for powering an intermediate-sized transport aircraft at the cruise speeds and altitudes typical of turbofan powered aircraft would require a single propeller of about 16 feet in diameter. It would also require the capability of generating about 15,000 shaft horsepower, which is several times the power output of conventional turboprop engines.
A conventional turboprop engine built to these requirements would further require the development of a relatively large and undesirably heavy reduction gearbox for transmitting the required power end torque at relatively low speed to -the propeller. The rotational speed of the large diameter propeller is a limiting factor for keeping the helical velocity of the propeller tip, i.e., aircraft velocity plus tangential velocity of the propeller tip, below supersonic speeds. This is desirable inasmuch as a propeller tip operating at supersonic speeds generates a significant amount of undesirable noise and results in a loss of aerodynamic efficiency.
Gas -turbine engines effective for driving propellers or fans without the use of a reduction gearbox are known in -the prior art. They typically include relatively low speed, counter rotating -turbine rotors having relatively few blade row stages driving a pair of counter-rotating fans or propellers. These engines comprise Lo l3Dv-8253 various embodiments that utilize the fans or propellers for merely augmenting the thrust generated prom the exhaust jet.
However, for propelling a modern, intermediate-sized aircraft that requires relatively large power output, a practical and relatively fuel efficient new generation engine having significant performance increases over conventional turbofan and turboprop engines and these counter rotating turbine rotor engines is required.
Accordingly, one object of the present invention is to provide a new and improved gas turbine engine.
Another object of the present invention is to provide a new and improved gas turbine engine including a power turbine having counter rotating rotors.
Another object of the present invention is to provide a new and improved gas turbine engine including a power turbine having a plurality of counter rotating turbine blade row stages wherein substantially all output power is obtained from expanding combustion gases through the stages and substantially little power remains in the exhaust gases leaving the engine.
Another object of the present invention is to provide a new and improved gas turbine engine wherein output power is obtainable without the use of a reduction gearbox.
Another object of the present invention is to provide a new and improved gas turbine engine including a gas generator r and a power turbine having counter rotating rotors, the power turbine being fixedly supported aft of the gas generator.
Another object of the present invention is to provide a new and improved gas turbine engine effective for providing counter rotating airfoil members such as propellers and fan blades.
SUMMERY OF THE INVENTION
The present invention comprises a new and improved gas turbine engine including a gas generator and a power --do--turbine. The power -turbine includes a first rotor and a plurality of first turbine blade rows extending radially outwardly therefrom, and a second rotor and a plurality of second turbine blade rows extending radially inwardly therefrom. The power turbine is supported aft of the gas generator and is effective for receiving combustion gases therefrom and expanding the gases through the first and second turbine blade rows for extracting substantially all output power therefrom for driving -the first and second rotors in counter rotating directions.
According to several embodiments of the invention, the power turbine is effective for driving counter rotating fans or propellers disposed either at a forward end or at an aft end of the engine.
BRIEF DESCRIPTION OF THE DRAWINGS
The invention, together with further objects and advantages thereof; is more particularly described in the following detailed description taken in conjunction with the accompanying drawings in which:
Figure 1 is a sectional view of a gas turbine engine according to one embodiment of the present invention including a power turbine having counter rotating rotors effective for driving counter rotating aft mounted propellers.
Figure 2 illustrates an aircraft including -two gas turbine engines such as in Figure 1 mounted to an aft end thereof.
Figure 3 is a view illustrating an alternative arrangement for mounting a gas turbine engine such as illustrated in Figure 1 to a wing of an aircraft.
Figure 4 is a sectional view of a gas turbine engine according to another embodiment of the present invention including a power turbine effective for driving counter rotating aft mounted fans Figure 5 is a sectional view of a gas turbine engine according to another embodiment of the present invention including a power turbine effective for driving counter rotating forward mounted fans.
I
Figure 6 is a sectional view of a gas turbine engine according to another embodiment of -the present invention wherein a booster compressor and an intermediate pressure turbine share a common drive shaft with a forward mounted fan and a rotor of a power turbine.
Figure 7 is a sectional view of a gas turbine engine according to another embodiment of the present invention including a power turbine effective for driving forward mounted counter rotating propellers, wherein an annular gas generator is disposed parallel to and spaced from a longitudinal axis of the engine.
DETAILED DESCRIPTION
Illustrated in Figure 1 is a gas turbine engine 10 according to one embodiment of the present invention. The engine 10 includes a longitudinal centerline axis 12 and an annular casing 14 disposed coccal about the axis I
The engine 10 also includes a conventional gas generator 16, which, for example, can comprise a booster compressor 18, a compressor 2Q, a combustor 22, a high pressure turbine I (HUT) 24/ and an intermediate pressure turbine (IT) 26 all arranged coccal about the longitudinal axis 12 of the engine 10 in serial, axial flow relationship. A first annular drive shaft 28 fixedly interconnects the compressor 20 and the HUT 24. A second annular drive shaft 30 fixedly interconnects the booster compressor 18 and the IT 26.
In operation, the gas generator 16 is effective for providing pressurized air from the booster 18 and the compressor 20 to the combustor 22 where it is mixed with fuel and suitably ignited for generating combustion gases.
The combustion gases drive the HUT 24 and the IT 26 which in turn drive the compressor 20 and the booster 18, respectively The combustion gases are discharged from the gas generator 16 through the IT 26 at a mean discharge radius from the longitudinal axis 12.
Attached to an aft most end of the casing 14 and aft of the gas generator 16 is an annular support member 30. The support member 30 extends radially inwardly and in an aft direction from the await end of the casing 14.
The support member 30 includes a plurality of circus-ferentially spaced strut members 32 extending radially inwardly from the aft end of the casing 14 and an annular hub member 34 fixedly attached to radially inner ends of the strut members 32 and extending in an aft direction.
The strut members 32 are effective for supporting the hub member 34 and channeling combustion gases from the gas generator 16 to a power -turbine 36 constructed in accordance with one embodiment of the present invention.
The power turbine 36, or simply low pressure turbine LOT
36~ is rotatable mounted to the hub member 34.
The LOT 36 includes a first annular drum rotor 38 rotatable mounted by suitable bearings 40 to the hub member 34 at forward and aft ends 42 and 44 thereof. The first rotor 38 includes a plurality of first turbine blade rows 46 extending radially outwardly therefrom and spaced axially thereon.
The IOTA also includes a second annular drum rotor 48 disposed radially outwardly of the first rotor 38 and the firs-t blade rows 46. The second rotor 48 includes a plurality of second turbine blade rows 50 extending radially inwardly therefrom and spaced axially thereon. The second rotor 48 is rotatable mounted to the hub member 34 by suitably bearings 52 disposed at radially inner ends of a forwardmostblade row aye of the second blade rows 50 and at radially inner ends of an aft most blade row 50b which is rotatable disposed on the first rotor 38 mounted to the hub member 34.
Each of the first and second turbine blade rows 46 and 50 comprises a plurality of circumferential spaced turbine blades, with -the first blade rows 46 alternately spaced with respective ones of the second blade rows 50.
Combustion gases flowing through the blade rows 46 and 50 flow along a mean flow path radius R2 which, by definition, represents a blade radius at which resultant work loads of the LOT 36 are assumed to be concentrated. or example, I
radius can be defined as the mean pitch line radius of all the blade rows of the LOT 36.
Cohesion gases being discharged from the gas generator 16 at the mean flow path radius Al are channeled through the strut members 32 to the LOT 36. The LOT 36 is effective for expanding the combustion gases through the first and second turbine blade rows 46 and 50 along the mean flow path radius R2 for extracting substantially all output power from the gases for driving the first and second rotors 38 and 48 in counter rotating directions at rotational speeds relatively lower than those of the first drive shaft 28.
The gas generator 16 and the LOT 36 as above arranged and described results in a new and improved gas turbine engine having counter rotating rotors effective for providing output shaft power at relatively low rotational speeds. Significant features of -the present invention include the complimentary arrangement of the engine elements. More specifically, the HUT 24 is disposed at of -the combustor 22 for first receiving the relatively high pressure combustion gases being discharged therefrom. The HUT 24 is most efficient when it and the firs-t drive shaft 28 are designed -to rotate at about 10,000 -to 15,000 RPM in a 15,000 shaft horsepower engine for most efficiently utilizing the high pressure combustion gases from the combustor 22.
The combustion gases after passing -through the HUT 24 are at a reduced, intermediate pressure. The intermediate pressure gases then flow through the IT
26 which further reduces the pressure of the gases to a relatively low pressure while most efficiently extracting power for rotating the second drive shaft 30 and the booster compressor 18 a-t speeds relatively lower than those of the HUT 24.
Finally, the low pressure combustion gases are channeled -to the LOT 36 where they are further expanded and substantially all of the remaining energy -thereof is extracted for rotating the first and second rotors 38 and 48 for providing Output shaft power. Little energy remains, and is thereby used less efficiently, in -the exhaust jet discharged from the LOT 36. Furthermore, inasmuch as the LOT 36 is the last element in the engine 10, it is subject to the lowest temperature combustion gases and therefore, thermally induced stresses are reduced.
For more efficiently extracting energy from the combustion gases in the LOT 36 it is preferable that -the mean flow path radius R2 thereof be greater than the mean discharge radius Al of the gas generator 16. In the embodiment illustrated in Figure 1, the mean flow path radius R2 is about double the magnitude of the mean discharge radius Al. This arrangement is effective for placing the turbine blade rows 46 and 50 at an increased radius from the longitudinal axis 12 for increasing the relative tangential velocities thereof for more efficiently extracting power from the gases flowing there over.
Inasmuch as the LOT 36 is a power turbine effective for providing substantially all output power through the rotors 38 and 48 and is preferably disposed aft of the gas generator 16, a suitable and effective mounting system is required. The support member 30 extending from the aft end of the casing 14 as above described is therefore also a significant feature of the present invention.
In the exemplary embodiment shown in Figure 1, the LOT 36 is effective for driving contorting oppositely pitched forward propellers 54 and aft propellers 56. More specifically, extending from an aft most end of the first rotor 38 is an aft blade row aye which extends radially outwardly to about the radial position of the second rotor 48. attached to radially outer ends of -the aft blade row aye it an annular shroud member 58. The aft propellers I r;
- 9 13DV ~352 56 are suitably attached to -the shroud member 58.
Similarly, the forward propellers 54 are suitably attached -to a forward end of the second rotor 48.
Suitable pitch varying means 60 are providing for independently controlling the pitch of the forward and aft propellers 54 and 56.
A most significant feature of the present invention is a gas -turbine engine 10 including an LOT
36 effective for providing relatively high output power and torque at relatively low rotational speeds without the use of a reduction gearbox. A reduction gearbox, and related accessories, would add a significant amount of weight and complexity to an engine capable of generating the relatively large thrust required for powering a transport aircraft such as the 150 passenger transport.
Speed reduction is required where a gas turbine engine is used for driving airfoil members such as propellers or fans. A conventional low pressure turbine (no-t shown) includes a single rotor typically rotating a-t about Lowe to 15,000 RPM. These rotational speeds must be reduced to relatively low speeds of about 1,000 to about 2,000 RPM for driving airfoil members. Propellers and fans are designed for moving a relatively large amount of air at relatively low axial speeds for generating a thrust, and operate more efficiently at the relatively low rotational speeds. Additionally, the low rotational speeds are required for limiting the helical tip speed of the propellers to below supersonic speeds.
According to the present invention, by allowing the second rotor 48 in Figure l of the LOT 36 to rotate in a direction opposite the first rotor 38, two output shafts, first rotor 38 and second rotor 48, are provided which rotate at about one quarter the speed of a single rotor, conventional OPT, thus providing speed reduction.
Furthermore, additional speed reduction is I
Accordingly, it would be desirable to combine the advantages of the turbofan and the turboprop for obtaining a compound engine having improved overall engine efficiency at aircraft cruise speeds typical of turbofan powered aircraft.
However, a simple scaled up version of a convent tonal turboprop engine suitable for powering an intermediate-sized transport aircraft at the cruise speeds and altitudes typical of turbofan powered aircraft would require a single propeller of about 16 feet in diameter. It would also require the capability of generating about 15,000 shaft horsepower, which is several times the power output of conventional turboprop engines.
A conventional turboprop engine built to these requirements would further require the development of a relatively large and undesirably heavy reduction gearbox for transmitting the required power end torque at relatively low speed to -the propeller. The rotational speed of the large diameter propeller is a limiting factor for keeping the helical velocity of the propeller tip, i.e., aircraft velocity plus tangential velocity of the propeller tip, below supersonic speeds. This is desirable inasmuch as a propeller tip operating at supersonic speeds generates a significant amount of undesirable noise and results in a loss of aerodynamic efficiency.
Gas -turbine engines effective for driving propellers or fans without the use of a reduction gearbox are known in -the prior art. They typically include relatively low speed, counter rotating -turbine rotors having relatively few blade row stages driving a pair of counter-rotating fans or propellers. These engines comprise Lo l3Dv-8253 various embodiments that utilize the fans or propellers for merely augmenting the thrust generated prom the exhaust jet.
However, for propelling a modern, intermediate-sized aircraft that requires relatively large power output, a practical and relatively fuel efficient new generation engine having significant performance increases over conventional turbofan and turboprop engines and these counter rotating turbine rotor engines is required.
Accordingly, one object of the present invention is to provide a new and improved gas turbine engine.
Another object of the present invention is to provide a new and improved gas turbine engine including a power turbine having counter rotating rotors.
Another object of the present invention is to provide a new and improved gas turbine engine including a power turbine having a plurality of counter rotating turbine blade row stages wherein substantially all output power is obtained from expanding combustion gases through the stages and substantially little power remains in the exhaust gases leaving the engine.
Another object of the present invention is to provide a new and improved gas turbine engine wherein output power is obtainable without the use of a reduction gearbox.
Another object of the present invention is to provide a new and improved gas turbine engine including a gas generator r and a power turbine having counter rotating rotors, the power turbine being fixedly supported aft of the gas generator.
Another object of the present invention is to provide a new and improved gas turbine engine effective for providing counter rotating airfoil members such as propellers and fan blades.
SUMMERY OF THE INVENTION
The present invention comprises a new and improved gas turbine engine including a gas generator and a power --do--turbine. The power -turbine includes a first rotor and a plurality of first turbine blade rows extending radially outwardly therefrom, and a second rotor and a plurality of second turbine blade rows extending radially inwardly therefrom. The power turbine is supported aft of the gas generator and is effective for receiving combustion gases therefrom and expanding the gases through the first and second turbine blade rows for extracting substantially all output power therefrom for driving -the first and second rotors in counter rotating directions.
According to several embodiments of the invention, the power turbine is effective for driving counter rotating fans or propellers disposed either at a forward end or at an aft end of the engine.
BRIEF DESCRIPTION OF THE DRAWINGS
The invention, together with further objects and advantages thereof; is more particularly described in the following detailed description taken in conjunction with the accompanying drawings in which:
Figure 1 is a sectional view of a gas turbine engine according to one embodiment of the present invention including a power turbine having counter rotating rotors effective for driving counter rotating aft mounted propellers.
Figure 2 illustrates an aircraft including -two gas turbine engines such as in Figure 1 mounted to an aft end thereof.
Figure 3 is a view illustrating an alternative arrangement for mounting a gas turbine engine such as illustrated in Figure 1 to a wing of an aircraft.
Figure 4 is a sectional view of a gas turbine engine according to another embodiment of the present invention including a power turbine effective for driving counter rotating aft mounted fans Figure 5 is a sectional view of a gas turbine engine according to another embodiment of the present invention including a power turbine effective for driving counter rotating forward mounted fans.
I
Figure 6 is a sectional view of a gas turbine engine according to another embodiment of -the present invention wherein a booster compressor and an intermediate pressure turbine share a common drive shaft with a forward mounted fan and a rotor of a power turbine.
Figure 7 is a sectional view of a gas turbine engine according to another embodiment of the present invention including a power turbine effective for driving forward mounted counter rotating propellers, wherein an annular gas generator is disposed parallel to and spaced from a longitudinal axis of the engine.
DETAILED DESCRIPTION
Illustrated in Figure 1 is a gas turbine engine 10 according to one embodiment of the present invention. The engine 10 includes a longitudinal centerline axis 12 and an annular casing 14 disposed coccal about the axis I
The engine 10 also includes a conventional gas generator 16, which, for example, can comprise a booster compressor 18, a compressor 2Q, a combustor 22, a high pressure turbine I (HUT) 24/ and an intermediate pressure turbine (IT) 26 all arranged coccal about the longitudinal axis 12 of the engine 10 in serial, axial flow relationship. A first annular drive shaft 28 fixedly interconnects the compressor 20 and the HUT 24. A second annular drive shaft 30 fixedly interconnects the booster compressor 18 and the IT 26.
In operation, the gas generator 16 is effective for providing pressurized air from the booster 18 and the compressor 20 to the combustor 22 where it is mixed with fuel and suitably ignited for generating combustion gases.
The combustion gases drive the HUT 24 and the IT 26 which in turn drive the compressor 20 and the booster 18, respectively The combustion gases are discharged from the gas generator 16 through the IT 26 at a mean discharge radius from the longitudinal axis 12.
Attached to an aft most end of the casing 14 and aft of the gas generator 16 is an annular support member 30. The support member 30 extends radially inwardly and in an aft direction from the await end of the casing 14.
The support member 30 includes a plurality of circus-ferentially spaced strut members 32 extending radially inwardly from the aft end of the casing 14 and an annular hub member 34 fixedly attached to radially inner ends of the strut members 32 and extending in an aft direction.
The strut members 32 are effective for supporting the hub member 34 and channeling combustion gases from the gas generator 16 to a power -turbine 36 constructed in accordance with one embodiment of the present invention.
The power turbine 36, or simply low pressure turbine LOT
36~ is rotatable mounted to the hub member 34.
The LOT 36 includes a first annular drum rotor 38 rotatable mounted by suitable bearings 40 to the hub member 34 at forward and aft ends 42 and 44 thereof. The first rotor 38 includes a plurality of first turbine blade rows 46 extending radially outwardly therefrom and spaced axially thereon.
The IOTA also includes a second annular drum rotor 48 disposed radially outwardly of the first rotor 38 and the firs-t blade rows 46. The second rotor 48 includes a plurality of second turbine blade rows 50 extending radially inwardly therefrom and spaced axially thereon. The second rotor 48 is rotatable mounted to the hub member 34 by suitably bearings 52 disposed at radially inner ends of a forwardmostblade row aye of the second blade rows 50 and at radially inner ends of an aft most blade row 50b which is rotatable disposed on the first rotor 38 mounted to the hub member 34.
Each of the first and second turbine blade rows 46 and 50 comprises a plurality of circumferential spaced turbine blades, with -the first blade rows 46 alternately spaced with respective ones of the second blade rows 50.
Combustion gases flowing through the blade rows 46 and 50 flow along a mean flow path radius R2 which, by definition, represents a blade radius at which resultant work loads of the LOT 36 are assumed to be concentrated. or example, I
radius can be defined as the mean pitch line radius of all the blade rows of the LOT 36.
Cohesion gases being discharged from the gas generator 16 at the mean flow path radius Al are channeled through the strut members 32 to the LOT 36. The LOT 36 is effective for expanding the combustion gases through the first and second turbine blade rows 46 and 50 along the mean flow path radius R2 for extracting substantially all output power from the gases for driving the first and second rotors 38 and 48 in counter rotating directions at rotational speeds relatively lower than those of the first drive shaft 28.
The gas generator 16 and the LOT 36 as above arranged and described results in a new and improved gas turbine engine having counter rotating rotors effective for providing output shaft power at relatively low rotational speeds. Significant features of -the present invention include the complimentary arrangement of the engine elements. More specifically, the HUT 24 is disposed at of -the combustor 22 for first receiving the relatively high pressure combustion gases being discharged therefrom. The HUT 24 is most efficient when it and the firs-t drive shaft 28 are designed -to rotate at about 10,000 -to 15,000 RPM in a 15,000 shaft horsepower engine for most efficiently utilizing the high pressure combustion gases from the combustor 22.
The combustion gases after passing -through the HUT 24 are at a reduced, intermediate pressure. The intermediate pressure gases then flow through the IT
26 which further reduces the pressure of the gases to a relatively low pressure while most efficiently extracting power for rotating the second drive shaft 30 and the booster compressor 18 a-t speeds relatively lower than those of the HUT 24.
Finally, the low pressure combustion gases are channeled -to the LOT 36 where they are further expanded and substantially all of the remaining energy -thereof is extracted for rotating the first and second rotors 38 and 48 for providing Output shaft power. Little energy remains, and is thereby used less efficiently, in -the exhaust jet discharged from the LOT 36. Furthermore, inasmuch as the LOT 36 is the last element in the engine 10, it is subject to the lowest temperature combustion gases and therefore, thermally induced stresses are reduced.
For more efficiently extracting energy from the combustion gases in the LOT 36 it is preferable that -the mean flow path radius R2 thereof be greater than the mean discharge radius Al of the gas generator 16. In the embodiment illustrated in Figure 1, the mean flow path radius R2 is about double the magnitude of the mean discharge radius Al. This arrangement is effective for placing the turbine blade rows 46 and 50 at an increased radius from the longitudinal axis 12 for increasing the relative tangential velocities thereof for more efficiently extracting power from the gases flowing there over.
Inasmuch as the LOT 36 is a power turbine effective for providing substantially all output power through the rotors 38 and 48 and is preferably disposed aft of the gas generator 16, a suitable and effective mounting system is required. The support member 30 extending from the aft end of the casing 14 as above described is therefore also a significant feature of the present invention.
In the exemplary embodiment shown in Figure 1, the LOT 36 is effective for driving contorting oppositely pitched forward propellers 54 and aft propellers 56. More specifically, extending from an aft most end of the first rotor 38 is an aft blade row aye which extends radially outwardly to about the radial position of the second rotor 48. attached to radially outer ends of -the aft blade row aye it an annular shroud member 58. The aft propellers I r;
- 9 13DV ~352 56 are suitably attached to -the shroud member 58.
Similarly, the forward propellers 54 are suitably attached -to a forward end of the second rotor 48.
Suitable pitch varying means 60 are providing for independently controlling the pitch of the forward and aft propellers 54 and 56.
A most significant feature of the present invention is a gas -turbine engine 10 including an LOT
36 effective for providing relatively high output power and torque at relatively low rotational speeds without the use of a reduction gearbox. A reduction gearbox, and related accessories, would add a significant amount of weight and complexity to an engine capable of generating the relatively large thrust required for powering a transport aircraft such as the 150 passenger transport.
Speed reduction is required where a gas turbine engine is used for driving airfoil members such as propellers or fans. A conventional low pressure turbine (no-t shown) includes a single rotor typically rotating a-t about Lowe to 15,000 RPM. These rotational speeds must be reduced to relatively low speeds of about 1,000 to about 2,000 RPM for driving airfoil members. Propellers and fans are designed for moving a relatively large amount of air at relatively low axial speeds for generating a thrust, and operate more efficiently at the relatively low rotational speeds. Additionally, the low rotational speeds are required for limiting the helical tip speed of the propellers to below supersonic speeds.
According to the present invention, by allowing the second rotor 48 in Figure l of the LOT 36 to rotate in a direction opposite the first rotor 38, two output shafts, first rotor 38 and second rotor 48, are provided which rotate at about one quarter the speed of a single rotor, conventional OPT, thus providing speed reduction.
Furthermore, additional speed reduction is I
3~3 obtainable by increasing the number of the first and second turbine blade rows 46 and So, i.e., the number of stages. This is so in that a-t lower rotational speeds of the rotors 38 and 48 less energy can be extracted from the combustion gases per stage of the LOT 36. For obtaining the desired reduced speeds and extracting substantially all remaining power from the combustion gases, an increased number of stages would be required.
However, a fewer number of stages could be used for accomplishing these objectives by having increased values of the ratio R2/Rl for providing the combustion gases to the LOT 36 at a larger mean flow path radius R2. Too many stages are undesirable because of the increased complexity, size and weight therefrom, and LOT 36 having fewer stages and a relatively high R2/Rl ratio is undesirable because of the increased frontal area and weight attributable thereto. As above-described and in accordance with the present invention, it has been determined that an R2/Rl ratio of about 2.0 is preferable.
Furthermore, in the embodiment illustrated in Figure 1 for driving the counter rotating propellers 54 and 56, the lot 36 having about 14 stages is preferred for obtaining output shaft speeds of the first and second 25 rotors 38 and 48 of about 1200 RPM. This speed is much less than the rotational speeds of the firs-t and second drive shafts 28 and 30.
In the embodiment illustrated in Figure 1, the counter rotating propellers 54 and 56 are aft mounted to the engine 10 radially outwardly of both -the first rotor 38 and the second rotor 48. These propellers have a hub radius R3 and a -tip radius R4 from the longitudinal axis 12.
In the embodiment of engine 10 including an LOT
36 driving propellers and having about 14 stages, it is also preferred that Rl/R4, R2/~4, and R3/R4, equal about I I
0.18~ 0.35, and 0.45~ respectively. Louvre, the number of stages of -the LOT 36 can range between about 10 and about 18 stages, and Rl/R4, R2/R4, and R3/R4 can range between it 0.2 to 0.16~ 0.4 to 0.3~ 0.5 to I all respectively. These relationships are preferred for obtaining an engine 10 suitable for most efficiently driving the counter rotating propellers 54 and 56 at rotational speeds of about 1200 RPM.
The reduction in speed of the rotors 38 and 48 of the LOT 36 results in a second order reduction of centrifugally generated stresses. For example a one quarter reduction in speed results in a one sixteenth reduction in centrifugal stress. This is significant in that the LOT 36 requires less material for accommodating centrifugal stress which results in a lighter LOT 36~ The overall effect of using a counter rotating LOT 36 is a significant reduction in engine weight as compared to an engine including a conventional LOT and reduction gearbox.
The embodiment of the engine 10 illustrated in Figure 1 results in additional advantages. For example, by mounting the propellers 54 and 56 to the aft end of the engine 10, an annular inlet region 62 of the engine 10 is relatively free of flow disturbing obstructions.
Accordingly, the inlet region 62 and an annular nacelle 64 surrounding the engine 10 can be suitably designed for obtaining increased aerodynamic performance of air entering the engine 10 as well as flowing there over.
The use of two propellers over a single propeller allows for propellers of lesser diameter, for example about 12 feet, i.e. t R4 = 6 feet, versus about 16 feet, respectively, for generating an equivalent amount of thrust at rotational speeds of about 1200 RPM and 900 RPM, respectively, and at aircraft cruise speeds of about Mach 0.7 to about Mach 0.8. The reduced diameter results in reduced propeller tip speeds and noise therefrom.
I
Mounting the propellers I and 56 radially outwardly of the second rotor 48 increases the hub to tip ratio R3/R~ of the propellers which provides an improvement in aerodynamic performance thereof.
Furthermore, the propellers do not obstruct the flow of combustion gases discharged from the LOT 36, which would otherwise reduce engine performance and require cooling schemes for preventing thermal damage to the propellers 54 and 56.
Illustrated in Figure 2 is an aircraft 66 including two engines 10 driving counter rotating propellers, such as the one illustrated in Figure 1, mounted to an aft most end of the aircraft 66. Aft mounted counter rotating propeller engines 10 according to the present invention are effective for providing an aircraft 66 having improved performance and fuel burn.
Furthermore, the engines 10 have reduced weight when compared with a conventional turboprop engine sized for identical thrust output. Reduced propeller noise is 2Q realizable which allows for a reduction in the amount of noise attenuation modifications to -the aircraft, and thus additionally reduces -total aircraft weight.
Illustrated in Figure 3 is an alternative arrangement for mounting counter rotating propeller engines 10, such as the one illustrated in Figure 1, to a wing 68 of an aircraft (not shown). In this embodiment, the hub member 34 of the engine 10 is extended in an aft direction and suitably mounted to the wing 68. A
stationary, annular exhaust duct 70 is suitably secured to the hub member 34 for suitably channeling the exhaust gases of the engine 10, for example, under -the wing 68.
The embodiment of the engine 10 illustrated in Figure 3 clearly illustrates a significant advantage of the support member 30 of the engine 10. More specifically, the support member 30 is not only effective for mounting the OPT 36 in the engine 10 but is also effective for mounting the entire engine 10 to a wing 68 of an aircraft.
3,~3~
Illustrated in Figure 4 is a gas turbine engine 72 according to another embodiment of the present invention.
The engine 72 includes a gas generator 16 which is substantially identical to the gas generator 16 of the engine 10 of Figure 1. In this embodiment, however, a LOT 74 drives counter rotating, forward and aft fans 76 and 78l respectively, mounted to an aft end of the engine 72. The fans 76 and 78 include a plurality of radially outwardly extending and circumferential spaced fan blades. An annular fan duct 80 is disposed radially outwardly of the fans 76 and 78 and is suitably attached by a plurality of strut members 82 to the casing 13 and the nacelle 64 of the engine 72. Suitable thrust reversing means (not shown) can be mounted to the hub member 34 and aft of the aft fan 78.
Inasmuch as fan blades operate differently than propeller blades, the LOT 74, although basically identical -to the LOT 36 of Figure 1, is preferably designed for driving fan blades. More specifically, the total number of stages of the first and second turbine blade rows 46 and 50 preferably ranges between about 6 stages to about 12 stages, with about 8 stages (shown in Figure 4) being Preferred. Correspondingly, Rl/R4 and R2/R4 preferably have values between 0.35 to 0.25 and 0.65 to 0.45, respectively. However for 8 stages, values of Rl/R4 and R2/R4 of 0.3 and 0.58l respectively are preferred. As in the embodiment illustrated in Figure 1, it is preferred that R2 have a larger value than Al and preferably a value twice as large.
Illustrated in Figure 4 is a gas turbine engine 84 according to another embodiment of the present invention.
The engine 84 includes a gas generator 16 which is sub-staunchly identical to the one illustrated in Figure 1.
The engine 84 also includes an LOT 86 which is substantially identical to the LOT 74 illustrated in Figure 4. However, in this embodiment, the LOT 86 preferably includes an additional, aft most blade row 50c -For a total of 9 stages, which stages are arranged for driving counter rotating -forward and aft fans 88 and 90, respectively, rotatable mounted to a forward most end of the engine 84. Disposed radially outwardly of the fans 88 and 90 is an annular fan duct 92 suitably secured by struts to the engine 84.
In contrast to the LOT 74 illustrated in Figure 4, an aft most end go of the first rotor 38 extends radially inwardly of the hub member 34 and is fixedly attached to a third annular drive shaft 98 which extends to the forward end of the engine 8~1 and is suitably attached to the aft fan 90. The aft most blade row 50c extends radially inwardly from the second rotor 48.
Radially inner ends 100 of the aft most blade row 50c are fixedly attached to a fourth drive shaft 102 which extends to the forward end of the engine 84 and is fixedly attached to the forward fan 88. The engine 84 thus includes four coccal mounted drive shafts 28~ 30, 98, and 102, with the LOT 86 being effective for driving the forward and aft fans 88 and 90, respectively in opposite directions.
The resulting engine 84 is capable of ultrahigh bypass ratios of greater than about 6 to 1.
Illustrated in Figure 6 is a gas turbine engine 104 according to another embodiment of the present invention.
In this embodiment, which is substantially identical to the embodiment illustrated in Figure 5, the aft fan 90 is fixedly connected to -the booster compressor 18, which are both driven by a common drive shaft, the third drive shaft 98 which is fixedly connected to the first rotor 38 of the LOT 86 and to the disc rotor of the LOT 26.
Illustrated in Figure 7 is a gas turbine engine 106 according to another embodiment of the present invention. This embodiment includes an LOT 108 which is substantially identical to the LOT 36 of Figure 1 -that includes 14 stages. However, the LOT I is arranged similarly to the LOT 86 of Figure 5 including the additional blade row 50c for a total of 15 stages and including the ~3;~3~3 13DV-8~53 third and fourth drive shafts 98 and 102. The drive shafts 98 and 102 are effective for driving counter rotating forward and aft variable pitch propellers 110 and 112, respectively, rotatable mounted to the forward most end of the engine 106.
In this embodiment, one, or a plurality of gas generators 114 are arranged for driving the LOT 108. The gas generator 114 is substantially identical to the gas generator 16 of Figure 1 and includes a longitudinal centerline axis 116. However, in contrast to the one illustrated in Figure 1, the gas generator 114 is mounted so that the longitudinal axis 116 thereof is parallel to and spaced from the longitudinal axis 12 of the engine 106.
A suitable annular duct 118 fluidly connects the gas generator 114 to the LOT 108 for providing combustion gases thereto In this embodiment, one or more gas generators 114 can be mounted circumferential about and parallel to the longitudinal axis 12 of the engine 106 for providing combustion gases to the PUT 108 for driving the counter-rotating propellers 110 and 112.
While there have been described herein what are considered to be preferred embodiments of the present invention, other embodiments will occur to those skilled in the art from the teachings herein.
For example, the gas generator 16 of Figure 1 without a booster compressor 18 and IT 26 can also be used for generating combustion gases. Furthermore, inasmuch as the counter rotating LOT 36 is effective for providing relatively large output power and torque at low speeds, gas turbine engines incorporating such Lots can be used for powering ships, generators, and large pumps, for example, which can be designed for having counter-rotating input shafts suitably attached to the first and second rotors 38 and 48 of the LOT 36.
Furthermore although the invention has been described as applied to a 15,000 shaft horsepower engine, it can also be sized for other engine classes. For example, 13~V-8253 in a smaller, 1500 shaft horsepower engine, powering shorter propellers 54 and 56, the HUT 24 would be designed to operate at about 30,000 RPM. The first rotor 38 and the second rotor 48 of the LOT 36 of Figure 1 would be correspondingly designed to operate at about a 10 to 1 speed reduction, i.e., at about 3,000 RPM. The propellers 54 and 56, although operating at about 3,000 RPM, have reduced tip radii R4 and therefore the helical tip speeds can be maintained below supersonic speeds.
However, a fewer number of stages could be used for accomplishing these objectives by having increased values of the ratio R2/Rl for providing the combustion gases to the LOT 36 at a larger mean flow path radius R2. Too many stages are undesirable because of the increased complexity, size and weight therefrom, and LOT 36 having fewer stages and a relatively high R2/Rl ratio is undesirable because of the increased frontal area and weight attributable thereto. As above-described and in accordance with the present invention, it has been determined that an R2/Rl ratio of about 2.0 is preferable.
Furthermore, in the embodiment illustrated in Figure 1 for driving the counter rotating propellers 54 and 56, the lot 36 having about 14 stages is preferred for obtaining output shaft speeds of the first and second 25 rotors 38 and 48 of about 1200 RPM. This speed is much less than the rotational speeds of the firs-t and second drive shafts 28 and 30.
In the embodiment illustrated in Figure 1, the counter rotating propellers 54 and 56 are aft mounted to the engine 10 radially outwardly of both -the first rotor 38 and the second rotor 48. These propellers have a hub radius R3 and a -tip radius R4 from the longitudinal axis 12.
In the embodiment of engine 10 including an LOT
36 driving propellers and having about 14 stages, it is also preferred that Rl/R4, R2/~4, and R3/R4, equal about I I
0.18~ 0.35, and 0.45~ respectively. Louvre, the number of stages of -the LOT 36 can range between about 10 and about 18 stages, and Rl/R4, R2/R4, and R3/R4 can range between it 0.2 to 0.16~ 0.4 to 0.3~ 0.5 to I all respectively. These relationships are preferred for obtaining an engine 10 suitable for most efficiently driving the counter rotating propellers 54 and 56 at rotational speeds of about 1200 RPM.
The reduction in speed of the rotors 38 and 48 of the LOT 36 results in a second order reduction of centrifugally generated stresses. For example a one quarter reduction in speed results in a one sixteenth reduction in centrifugal stress. This is significant in that the LOT 36 requires less material for accommodating centrifugal stress which results in a lighter LOT 36~ The overall effect of using a counter rotating LOT 36 is a significant reduction in engine weight as compared to an engine including a conventional LOT and reduction gearbox.
The embodiment of the engine 10 illustrated in Figure 1 results in additional advantages. For example, by mounting the propellers 54 and 56 to the aft end of the engine 10, an annular inlet region 62 of the engine 10 is relatively free of flow disturbing obstructions.
Accordingly, the inlet region 62 and an annular nacelle 64 surrounding the engine 10 can be suitably designed for obtaining increased aerodynamic performance of air entering the engine 10 as well as flowing there over.
The use of two propellers over a single propeller allows for propellers of lesser diameter, for example about 12 feet, i.e. t R4 = 6 feet, versus about 16 feet, respectively, for generating an equivalent amount of thrust at rotational speeds of about 1200 RPM and 900 RPM, respectively, and at aircraft cruise speeds of about Mach 0.7 to about Mach 0.8. The reduced diameter results in reduced propeller tip speeds and noise therefrom.
I
Mounting the propellers I and 56 radially outwardly of the second rotor 48 increases the hub to tip ratio R3/R~ of the propellers which provides an improvement in aerodynamic performance thereof.
Furthermore, the propellers do not obstruct the flow of combustion gases discharged from the LOT 36, which would otherwise reduce engine performance and require cooling schemes for preventing thermal damage to the propellers 54 and 56.
Illustrated in Figure 2 is an aircraft 66 including two engines 10 driving counter rotating propellers, such as the one illustrated in Figure 1, mounted to an aft most end of the aircraft 66. Aft mounted counter rotating propeller engines 10 according to the present invention are effective for providing an aircraft 66 having improved performance and fuel burn.
Furthermore, the engines 10 have reduced weight when compared with a conventional turboprop engine sized for identical thrust output. Reduced propeller noise is 2Q realizable which allows for a reduction in the amount of noise attenuation modifications to -the aircraft, and thus additionally reduces -total aircraft weight.
Illustrated in Figure 3 is an alternative arrangement for mounting counter rotating propeller engines 10, such as the one illustrated in Figure 1, to a wing 68 of an aircraft (not shown). In this embodiment, the hub member 34 of the engine 10 is extended in an aft direction and suitably mounted to the wing 68. A
stationary, annular exhaust duct 70 is suitably secured to the hub member 34 for suitably channeling the exhaust gases of the engine 10, for example, under -the wing 68.
The embodiment of the engine 10 illustrated in Figure 3 clearly illustrates a significant advantage of the support member 30 of the engine 10. More specifically, the support member 30 is not only effective for mounting the OPT 36 in the engine 10 but is also effective for mounting the entire engine 10 to a wing 68 of an aircraft.
3,~3~
Illustrated in Figure 4 is a gas turbine engine 72 according to another embodiment of the present invention.
The engine 72 includes a gas generator 16 which is substantially identical to the gas generator 16 of the engine 10 of Figure 1. In this embodiment, however, a LOT 74 drives counter rotating, forward and aft fans 76 and 78l respectively, mounted to an aft end of the engine 72. The fans 76 and 78 include a plurality of radially outwardly extending and circumferential spaced fan blades. An annular fan duct 80 is disposed radially outwardly of the fans 76 and 78 and is suitably attached by a plurality of strut members 82 to the casing 13 and the nacelle 64 of the engine 72. Suitable thrust reversing means (not shown) can be mounted to the hub member 34 and aft of the aft fan 78.
Inasmuch as fan blades operate differently than propeller blades, the LOT 74, although basically identical -to the LOT 36 of Figure 1, is preferably designed for driving fan blades. More specifically, the total number of stages of the first and second turbine blade rows 46 and 50 preferably ranges between about 6 stages to about 12 stages, with about 8 stages (shown in Figure 4) being Preferred. Correspondingly, Rl/R4 and R2/R4 preferably have values between 0.35 to 0.25 and 0.65 to 0.45, respectively. However for 8 stages, values of Rl/R4 and R2/R4 of 0.3 and 0.58l respectively are preferred. As in the embodiment illustrated in Figure 1, it is preferred that R2 have a larger value than Al and preferably a value twice as large.
Illustrated in Figure 4 is a gas turbine engine 84 according to another embodiment of the present invention.
The engine 84 includes a gas generator 16 which is sub-staunchly identical to the one illustrated in Figure 1.
The engine 84 also includes an LOT 86 which is substantially identical to the LOT 74 illustrated in Figure 4. However, in this embodiment, the LOT 86 preferably includes an additional, aft most blade row 50c -For a total of 9 stages, which stages are arranged for driving counter rotating -forward and aft fans 88 and 90, respectively, rotatable mounted to a forward most end of the engine 84. Disposed radially outwardly of the fans 88 and 90 is an annular fan duct 92 suitably secured by struts to the engine 84.
In contrast to the LOT 74 illustrated in Figure 4, an aft most end go of the first rotor 38 extends radially inwardly of the hub member 34 and is fixedly attached to a third annular drive shaft 98 which extends to the forward end of the engine 8~1 and is suitably attached to the aft fan 90. The aft most blade row 50c extends radially inwardly from the second rotor 48.
Radially inner ends 100 of the aft most blade row 50c are fixedly attached to a fourth drive shaft 102 which extends to the forward end of the engine 84 and is fixedly attached to the forward fan 88. The engine 84 thus includes four coccal mounted drive shafts 28~ 30, 98, and 102, with the LOT 86 being effective for driving the forward and aft fans 88 and 90, respectively in opposite directions.
The resulting engine 84 is capable of ultrahigh bypass ratios of greater than about 6 to 1.
Illustrated in Figure 6 is a gas turbine engine 104 according to another embodiment of the present invention.
In this embodiment, which is substantially identical to the embodiment illustrated in Figure 5, the aft fan 90 is fixedly connected to -the booster compressor 18, which are both driven by a common drive shaft, the third drive shaft 98 which is fixedly connected to the first rotor 38 of the LOT 86 and to the disc rotor of the LOT 26.
Illustrated in Figure 7 is a gas turbine engine 106 according to another embodiment of the present invention. This embodiment includes an LOT 108 which is substantially identical to the LOT 36 of Figure 1 -that includes 14 stages. However, the LOT I is arranged similarly to the LOT 86 of Figure 5 including the additional blade row 50c for a total of 15 stages and including the ~3;~3~3 13DV-8~53 third and fourth drive shafts 98 and 102. The drive shafts 98 and 102 are effective for driving counter rotating forward and aft variable pitch propellers 110 and 112, respectively, rotatable mounted to the forward most end of the engine 106.
In this embodiment, one, or a plurality of gas generators 114 are arranged for driving the LOT 108. The gas generator 114 is substantially identical to the gas generator 16 of Figure 1 and includes a longitudinal centerline axis 116. However, in contrast to the one illustrated in Figure 1, the gas generator 114 is mounted so that the longitudinal axis 116 thereof is parallel to and spaced from the longitudinal axis 12 of the engine 106.
A suitable annular duct 118 fluidly connects the gas generator 114 to the LOT 108 for providing combustion gases thereto In this embodiment, one or more gas generators 114 can be mounted circumferential about and parallel to the longitudinal axis 12 of the engine 106 for providing combustion gases to the PUT 108 for driving the counter-rotating propellers 110 and 112.
While there have been described herein what are considered to be preferred embodiments of the present invention, other embodiments will occur to those skilled in the art from the teachings herein.
For example, the gas generator 16 of Figure 1 without a booster compressor 18 and IT 26 can also be used for generating combustion gases. Furthermore, inasmuch as the counter rotating LOT 36 is effective for providing relatively large output power and torque at low speeds, gas turbine engines incorporating such Lots can be used for powering ships, generators, and large pumps, for example, which can be designed for having counter-rotating input shafts suitably attached to the first and second rotors 38 and 48 of the LOT 36.
Furthermore although the invention has been described as applied to a 15,000 shaft horsepower engine, it can also be sized for other engine classes. For example, 13~V-8253 in a smaller, 1500 shaft horsepower engine, powering shorter propellers 54 and 56, the HUT 24 would be designed to operate at about 30,000 RPM. The first rotor 38 and the second rotor 48 of the LOT 36 of Figure 1 would be correspondingly designed to operate at about a 10 to 1 speed reduction, i.e., at about 3,000 RPM. The propellers 54 and 56, although operating at about 3,000 RPM, have reduced tip radii R4 and therefore the helical tip speeds can be maintained below supersonic speeds.
Claims (10)
1. A gas turbine engine comprising:
a gas generator effective for generating combustion gases; and a power turbine including a first rotor having a plurality of first turbine blade rows extending radially outwardly therefrom and a second rotor having a plurality of second turbine blade rows extending radially inwardly therefrom, wherein said first and second rotors are arranged so as to define inner and outer flowpath surfaces, respectively, for said combustion gases flowing through said power turbine; and said power turbine is effective for receiving said combustion gases and extracting substantially all output power therefrom for driving said first and second rotors in counterrotating directions; and an annular casing disposed circumferentially about said gas generator.
a gas generator effective for generating combustion gases; and a power turbine including a first rotor having a plurality of first turbine blade rows extending radially outwardly therefrom and a second rotor having a plurality of second turbine blade rows extending radially inwardly therefrom, wherein said first and second rotors are arranged so as to define inner and outer flowpath surfaces, respectively, for said combustion gases flowing through said power turbine; and said power turbine is effective for receiving said combustion gases and extracting substantially all output power therefrom for driving said first and second rotors in counterrotating directions; and an annular casing disposed circumferentially about said gas generator.
2. A gas turbine engine including a longitudinal axis comprising:
an annular casing disposed coaxially about said axis;
a gas generator disposed in said casing and including a compressor, combustor and a high pressure turbine in serial flow relationship, said high pressure turbine being effective for first receiving combustion gases from said combustor for driving said compressor through a drive shaft fixedly connected thereto, said gas generator being effective for exhausting said combustion gases therefrom substantially at a mean discharge radius from said longitudional axis and in an aft direction;
a power turbine disposed coaxially about said longitudinal axis and including:
a first drum rotor arranged so as to define an inner flowpath surface for said combustion gases flowing through said power turbine;
a plurality of first turbine blade rows extending radially outwardly from said first rotor and spaced axially thereon;
a second drum rotor arranged so as to define an outer flowpath surface for said combustion gases flowing through said power turbine and disposed radially outwardly of said first rotor an dsaid first turbine blade rows; and a plurality of second turbine blade rows extending radially inwardly from said second rotor and alternately spaced with respective ones of said plurality of said first turbine blade rows;
said power turbine being effective for receiving said combustion gases from said gas generator and expanding said gases through said first and second turbine blade rows along a mean flowpath radius for extracting substantially all output power therefrom for driving said first and second rotors in counterrotating directions at speeds relatively lower than those of said drive shaft;
and first and second counterrotating propellers driven by said first and second rotors, respectively.
an annular casing disposed coaxially about said axis;
a gas generator disposed in said casing and including a compressor, combustor and a high pressure turbine in serial flow relationship, said high pressure turbine being effective for first receiving combustion gases from said combustor for driving said compressor through a drive shaft fixedly connected thereto, said gas generator being effective for exhausting said combustion gases therefrom substantially at a mean discharge radius from said longitudional axis and in an aft direction;
a power turbine disposed coaxially about said longitudinal axis and including:
a first drum rotor arranged so as to define an inner flowpath surface for said combustion gases flowing through said power turbine;
a plurality of first turbine blade rows extending radially outwardly from said first rotor and spaced axially thereon;
a second drum rotor arranged so as to define an outer flowpath surface for said combustion gases flowing through said power turbine and disposed radially outwardly of said first rotor an dsaid first turbine blade rows; and a plurality of second turbine blade rows extending radially inwardly from said second rotor and alternately spaced with respective ones of said plurality of said first turbine blade rows;
said power turbine being effective for receiving said combustion gases from said gas generator and expanding said gases through said first and second turbine blade rows along a mean flowpath radius for extracting substantially all output power therefrom for driving said first and second rotors in counterrotating directions at speeds relatively lower than those of said drive shaft;
and first and second counterrotating propellers driven by said first and second rotors, respectively.
3. A gas turbine engine according to claim 2 wherein the total number of said first and second turbine blade rows is less than about 18 rows and more than about 10 rows.
4. A gas turbine engine according to claim 2 wherein:
said propellers have a tip radius and a hub radius; and said mean discharge radius of said gas generator, said means flowpath radius of said power turbine and said hub radius of said propellers have magnitudes with respect to said tip radius of said propellers of between 0.2 to 0.16, 0.4 to 0.3, 0.5 to 0.4, respectively, corresponding to a total number of said first and second turbine blade rows greater than about 10 rows and less than about 18 rows, respectively.
said propellers have a tip radius and a hub radius; and said mean discharge radius of said gas generator, said means flowpath radius of said power turbine and said hub radius of said propellers have magnitudes with respect to said tip radius of said propellers of between 0.2 to 0.16, 0.4 to 0.3, 0.5 to 0.4, respectively, corresponding to a total number of said first and second turbine blade rows greater than about 10 rows and less than about 18 rows, respectively.
5. A gas turbine engine comprising:
a gas generator effective for generating combustion gases; and means for efficiently transferring the energy of said gases into a net engine thrust, said means including:
a power turbine including a first rotor having a plurality of first turbine blade rows extending radially outwardly therefrom and a second rotor having a plurality of second turbine blade rows extending radially inwardly therefrom, wherein:
said first and second rotors are arranged so as to define inner and outer flowpath surfaces, respectively, for said combustion gases flowing through said power turbine; and said power turbine is effective for receiving said combustion gases and extracting substantially all output power therefrom for driving said first and second rotors in counterrotating directions;
first and second counterrotating propellers, each with a plurality of blades attached to first and second rotatable nacelle rings, respectively, wherein:
said first and second propellers are directly coupled to and driven by said first and second rotors, respectively, and disposed radially outwardly of said power turbine; and each of said blades has a hub radius to tip radius ratio of between 0.5 and 0.4.
a gas generator effective for generating combustion gases; and means for efficiently transferring the energy of said gases into a net engine thrust, said means including:
a power turbine including a first rotor having a plurality of first turbine blade rows extending radially outwardly therefrom and a second rotor having a plurality of second turbine blade rows extending radially inwardly therefrom, wherein:
said first and second rotors are arranged so as to define inner and outer flowpath surfaces, respectively, for said combustion gases flowing through said power turbine; and said power turbine is effective for receiving said combustion gases and extracting substantially all output power therefrom for driving said first and second rotors in counterrotating directions;
first and second counterrotating propellers, each with a plurality of blades attached to first and second rotatable nacelle rings, respectively, wherein:
said first and second propellers are directly coupled to and driven by said first and second rotors, respectively, and disposed radially outwardly of said power turbine; and each of said blades has a hub radius to tip radius ratio of between 0.5 and 0.4.
6. A gas turbine engine including a longitudinal axis comprising:
an annular casing disposed coaxially about said axis;
a gas generator disposed in said casing and including a compressor, combustor and a high pressure turbine in serial flow relationship, said high pressure turbine being effective for first receiving combustion gases from said combustor for driving said compressor through a drive shaft fixedly connected thereto, said gas generator being effective for exhausting said combus-tion gases therefrom substantially at a mean discharge radius from said longitudinal axis and in an aft direction;
a power turbine disposed coaxially about said longitudinal axis and including:
a first drum rotor arranged so as to define an inner flowpath surface for said combustion gases flowing through said power turbine;
a plurality of direct turbine blade rows extending radially outwardly from said first rotor and spaced axially thereon;
a second drum rotor arranged so as to define an outer flowpath surface for said combustion gases flowing through said power turbine and disposed radially outwardly of said first rotor and said first turbine blade rows; and a plurality of second turbine blade rows extending radially inwardly from said second rotor and alternately spaced with respective ones of said plurality of said first turbine blade rows;
said power turbine being effective for receiving said combustion gases from said gas generator and expanding said gases through said first and second turbine blade rows along a mean flowpath radius for extracting substantially all output power therefrom for driving said first and second rotors in counterrotating directions at speeds relatively lower than those of said drive shaft;
and first and second counterrotating fans driven by said first and second rotors, respectively.
an annular casing disposed coaxially about said axis;
a gas generator disposed in said casing and including a compressor, combustor and a high pressure turbine in serial flow relationship, said high pressure turbine being effective for first receiving combustion gases from said combustor for driving said compressor through a drive shaft fixedly connected thereto, said gas generator being effective for exhausting said combus-tion gases therefrom substantially at a mean discharge radius from said longitudinal axis and in an aft direction;
a power turbine disposed coaxially about said longitudinal axis and including:
a first drum rotor arranged so as to define an inner flowpath surface for said combustion gases flowing through said power turbine;
a plurality of direct turbine blade rows extending radially outwardly from said first rotor and spaced axially thereon;
a second drum rotor arranged so as to define an outer flowpath surface for said combustion gases flowing through said power turbine and disposed radially outwardly of said first rotor and said first turbine blade rows; and a plurality of second turbine blade rows extending radially inwardly from said second rotor and alternately spaced with respective ones of said plurality of said first turbine blade rows;
said power turbine being effective for receiving said combustion gases from said gas generator and expanding said gases through said first and second turbine blade rows along a mean flowpath radius for extracting substantially all output power therefrom for driving said first and second rotors in counterrotating directions at speeds relatively lower than those of said drive shaft;
and first and second counterrotating fans driven by said first and second rotors, respectively.
7. A gas turbine according to claim 6 wherein the total number of said first and second turbine blade rows is less than about 12 rows and more than about 6 rows.
8. A gas turbine according to claim 6 wherein said fans have a tip radius and a hub radius and said mean discharge radius of said gas generator and said mean flowpath radius of said power turbine have magnitudes with respect to the tip radius of said fans of between 0.35 to 0.25 and 0.65 to 0.45, respectively, corresponding to a total number of said first and second turbine blade rows greater than about 6 rows and less than about 12 rows.
9. An unducted fan engine for subsonic aircraft comprising:
a gas generator including a compressor, a combustor and a high pressure turbine for driving said compressor, a power turbine for receiving the hot gases exhausted from said high pressure turbine, said power turbine including a first rotor having a plurality of turbine blade rows extending radially outwardly therefrom and a second rotor having a plurality of turbine blade rows extending radially inwardly therefrom and interspersed with the blade rows of the first rotor, said first and second rotors having generally annular inner and outer surfaces defining the flowpath for the hot gases flowing through said power turbine, first and second propeller systems each having an annular nacelle ring surrounding the power turbine and a plurality of propeller blades mounted on said ring, one of said rings being mounted on and driven directly by said first power turbine rotor and the second of said rings being mounted on and driven directly by the second power turbine rotor whereby said propeller systems are rotated in opposite directions by said rotors, and said power turbine being effective to remove substantially all of the usable energy from the gases passing through it so that said propeller blades are driven thereby to produce substantially all the thrust generated by said engine.
a gas generator including a compressor, a combustor and a high pressure turbine for driving said compressor, a power turbine for receiving the hot gases exhausted from said high pressure turbine, said power turbine including a first rotor having a plurality of turbine blade rows extending radially outwardly therefrom and a second rotor having a plurality of turbine blade rows extending radially inwardly therefrom and interspersed with the blade rows of the first rotor, said first and second rotors having generally annular inner and outer surfaces defining the flowpath for the hot gases flowing through said power turbine, first and second propeller systems each having an annular nacelle ring surrounding the power turbine and a plurality of propeller blades mounted on said ring, one of said rings being mounted on and driven directly by said first power turbine rotor and the second of said rings being mounted on and driven directly by the second power turbine rotor whereby said propeller systems are rotated in opposite directions by said rotors, and said power turbine being effective to remove substantially all of the usable energy from the gases passing through it so that said propeller blades are driven thereby to produce substantially all the thrust generated by said engine.
10. An unducted fan engine according to claim 9 wherein:
said engine further includes a nacelle surrounding said gas generator and forming the only surface for guiding the air to said propeller blades, and said annular nacelle rings form the only surfaces controlling the air flow in the region of said blades.
said engine further includes a nacelle surrounding said gas generator and forming the only surface for guiding the air to said propeller blades, and said annular nacelle rings form the only surfaces controlling the air flow in the region of said blades.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US43792382A | 1982-11-01 | 1982-11-01 | |
| US437,923 | 1982-11-01 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1233325A true CA1233325A (en) | 1988-03-01 |
Family
ID=23738483
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000438676A Expired CA1233325A (en) | 1982-11-01 | 1983-10-07 | Counter rotation power turbine |
Country Status (9)
| Country | Link |
|---|---|
| JP (1) | JPS59103947A (en) |
| AU (1) | AU594300B2 (en) |
| CA (1) | CA1233325A (en) |
| DE (1) | DE3338456A1 (en) |
| FR (1) | FR2535394B1 (en) |
| GB (1) | GB2129502B (en) |
| IT (1) | IT1171784B (en) |
| NL (1) | NL8303401A (en) |
| SE (1) | SE8305993L (en) |
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| SE456075B (en) * | 1984-11-29 | 1988-09-05 | Volvo Penta Ab | ROTOR SYSTEM, PREFERRED BAT PROPELLER SYSTEM |
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| US4860537A (en) * | 1986-08-29 | 1989-08-29 | Brandt, Inc. | High bypass ratio counterrotating gearless front fan engine |
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| DE3917970A1 (en) * | 1988-06-17 | 1989-12-21 | Gen Electric | MANIFOLD-FREE BLOW ENGINE FIXED TO THE SUPPORT |
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| US4927329A (en) * | 1988-10-21 | 1990-05-22 | General Electric Company | Aircraft engine unducted fan blade pitch control system |
| US4936748A (en) * | 1988-11-28 | 1990-06-26 | General Electric Company | Auxiliary power source in an unducted fan gas turbine engine |
| US4969325A (en) * | 1989-01-03 | 1990-11-13 | General Electric Company | Turbofan engine having a counterrotating partially geared fan drive turbine |
| US5010729A (en) * | 1989-01-03 | 1991-04-30 | General Electric Company | Geared counterrotating turbine/fan propulsion system |
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| US5082424A (en) * | 1989-06-05 | 1992-01-21 | General Electric Company | Connection system for aircraft propeller blades |
| GB2264907A (en) * | 1992-02-10 | 1993-09-15 | Peter Antony Hulmes | Multi-engined aircraft. |
| US6666017B2 (en) * | 2002-05-24 | 2003-12-23 | General Electric Company | Counterrotatable booster compressor assembly for a gas turbine engine |
| US7758303B1 (en) * | 2006-07-31 | 2010-07-20 | General Electric Company | FLADE fan with different inner and outer airfoil stagger angles at a shroud therebetween |
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| FR2918120B1 (en) † | 2007-06-28 | 2009-10-02 | Snecma Sa | DOUBLE BLOWER TURBOMACHINE |
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| FR2937678B1 (en) * | 2008-10-23 | 2013-11-22 | Snecma | DEVICE FOR CONTROLLING THE ORIENTATION OF BLOWER BLADES OF A TURBOPROPULSEUR |
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| FR2940247B1 (en) * | 2008-12-19 | 2011-01-21 | Snecma | SYSTEM OF CONTRAROTATIVE PROPELLERS DRAWN BY AN EPICYCLOIDAL TRAIN PROVIDING A BALANCED TORQUE DISTRIBUTION BETWEEN THE TWO PROPELLERS |
| FR2946011B1 (en) | 2009-05-29 | 2013-01-11 | Snecma | MOBILE ROLLER DEVICE FOR CONTROLLING THE ORIENTATION OF BLOWER BLADES OF A TURBOPROPULSOR |
| FR2946012B1 (en) | 2009-05-29 | 2011-06-24 | Snecma | DEVICE FOR CONTROLLING THE ORIENTATION OF BLOWER BLADES OF A TURBOPROPULSEUR |
| GB0911100D0 (en) * | 2009-06-29 | 2009-08-12 | Rolls Royce Plc | Propulsive fan system |
| FR2956854B1 (en) * | 2010-03-01 | 2012-08-17 | Snecma | DEVICE FOR CONTROLLING THE ORIENTATION OF BLOWER BLADES OF A COUNTERPROOF TURBOPROPULSOR. |
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| JP6657250B2 (en) * | 2015-04-03 | 2020-03-04 | ターボデン ソシエタ ペル アツィオーニＴｕｒｂｏｄｅｎ Ｓｐａ | Multi-stage turbine, preferably for an organic Rankine cycle ORC plant |
| US9771878B2 (en) * | 2015-10-19 | 2017-09-26 | General Electric Company | Thrust scheduling method for variable pitch fan engines and turbo-shaft, turbo-propeller engines |
| FR3095670B1 (en) * | 2019-04-30 | 2021-12-03 | Safran Aircraft Engines | Improved architecture of a contra-rotating turbine engine |
| CN113982781B (en) * | 2021-08-18 | 2023-08-11 | 高阳 | Multi-rotor impeller of air compressor and turbine multi-rotor full-contra-rotating aeroengine |
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| US2505660A (en) * | 1950-04-25 | Augmentor fob jet propulsion hav | ||
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-
1983
- 1983-10-04 NL NL8303401A patent/NL8303401A/en not_active Application Discontinuation
- 1983-10-07 CA CA000438676A patent/CA1233325A/en not_active Expired
- 1983-10-13 AU AU20146/83A patent/AU594300B2/en not_active Ceased
- 1983-10-22 DE DE19833338456 patent/DE3338456A1/en not_active Withdrawn
- 1983-10-24 GB GB8328398A patent/GB2129502B/en not_active Expired
- 1983-10-27 IT IT23479/83A patent/IT1171784B/en active
- 1983-10-28 FR FR838317250A patent/FR2535394B1/en not_active Expired - Fee Related
- 1983-11-01 SE SE8305993A patent/SE8305993L/en not_active Application Discontinuation
- 1983-11-01 JP JP58203763A patent/JPS59103947A/en active Granted
Also Published As
| Publication number | Publication date |
|---|---|
| AU594300B2 (en) | 1990-03-08 |
| JPS59103947A (en) | 1984-06-15 |
| SE8305993L (en) | 1984-05-02 |
| FR2535394B1 (en) | 1992-04-30 |
| FR2535394A1 (en) | 1984-05-04 |
| GB2129502A (en) | 1984-05-16 |
| GB2129502B (en) | 1989-10-18 |
| IT1171784B (en) | 1987-06-10 |
| DE3338456A1 (en) | 1984-05-03 |
| AU2014683A (en) | 1984-05-10 |
| GB8328398D0 (en) | 1983-11-23 |
| IT8323479A0 (en) | 1983-10-27 |
| SE8305993D0 (en) | 1983-11-01 |
| NL8303401A (en) | 1984-06-01 |
| JPH0351899B2 (en) | 1991-08-08 |
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