GB2129502A

GB2129502A - Counter rotation power turbine

Info

Publication number: GB2129502A
Application number: GB08328398A
Authority: GB
Inventors: Kenneth Odell Johnson
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1982-11-01
Filing date: 1983-10-24
Publication date: 1984-05-16
Also published as: FR2535394B1; AU2014683A; FR2535394A1; GB2129502B; JPH0351899B2; CA1233325A; AU594300B2; IT8323479A0; GB8328398D0; DE3338456A1; NL8303401A; SE8305993D0; SE8305993L; IT1171784B; JPS59103947A

Abstract

The power turbine includes a first rotor 38 having a plurality of first turbine blade rows extending radially outwardly therefrom, and a second rotor 48 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 for driving counterrotating fans or propellers disposed either at a forward end or at an aft end of the engine. <IMAGE>

Description

SPECIFICATION Counter rotation power turbine Field 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 counterrotating 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 power a fan, whereas the turboprop engine includes a gas generator to power 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 designed for lower cruise speeds, conventional turboprops are typically used inasmuch as they can provide superior performance and efficiency. For example, significant reductions in fuel burn, i.e., the amount of fuel consumed 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 efficiency at aircraft cruise speeds typical of turbofan powered aircraft.

However, a simple scaled up version of a conventional 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 and torque at relatively low speed to the propeller. The rotational speed of the large diameter propeller is a limiting factor four 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 propellors or fans without the use of a reduction gearbox are known in the prior art. They typically include relatively low speed, counterrotating tubine rotors having relatively few blade row stages driving a pair of counterrotating fans or propellors. These engines comprise various embodiments that utilize the fans or propellors for merely augmenting the thrust generated from the exhaust jet.

However, for propelling a modern, intermediatesized 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 counterrotating turbine rotor engines is required.

The present invention seeks to meet this requirement.

The gas turbine engine disclosed herein according to a preferred embodiment comprises a gas generator and a power 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 counterrotating directions.

According to several embodiments of the invention, the power turbine is effective for driving counterrotating fans or propellors disposed either at a forward end or at an aft end of the engine.

Embodiments of the invention are more particularly described in the following detailed description taken in conjunction with the accompanying drawings in which: Figure lisa sectional view of a gas turbine engine according to one embodiment of the present invention including a power turbine having counterrotating rotors effective for driving counterrotating aft mounted propellors.

Figure2 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 counterrotating 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 counterrotating forward mounted fans.

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 counterrotating propellors, wherein an annular gas generator is disposed parallel to and spaced from a longitudinal axis of the engine.

Detailed description Illustrated in Figure lisa 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 coaxially about the axis 12. The engine 10 also includes a conventional gas generator 16, which, for example, can comprise a booster compressor 18, a compressor 20, a combustor 22, a high pressure turbine (HPT) 24, and an intermediate pressure turbine (IPT) 26 all arranged coaxially about the longitudinal axis 12 of the engine 10 in serial, axial flow relationship. Afirst annular drive shaft 28 fixedly interconnects the compressor 20 and the HPT 24. A second annular drive shaft 30 fixedly interconnects the booster compressor 18 and the IPT 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 HPT 24 and the IPT 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 IPT26 at a mean discharge radius R1 from the longitudinal axis 12.

Attached to an aftmost 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 aft end of the casing 14. The support member 30 includes a plurality of circumferentially spaced strut members 32 extending radially inwardly from the aft end of the casing 14 and an annular hub member34fixedly 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 (LPT) 36, is rotatably mounted to the hub member 34.

The LPT 36 includes first annular drum rotor 38 rotatably 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 LPT 36 also includes a second annular drum rotor 48 disposed radially outwardly of the first rotor 38 and the first 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 rotatably mounted to the hub member 34 by suitable bearings 52 disposed at radially inner ends of a forwardmost blade row 50a of the second blade rows 50 and at radially inner ends of an aftmost blade row 50b which is rotatably 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 circumferentially 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 flowpath radius R2 which, by definition, represents a blade radius at which resultant work loads of the LPT 36 are assumed to be concentrated. For example, radius R2 can be defined as the mean pitch line radius of all the blade rows of the LPT 36.

Combustion gases being discharged from the gas generator 16 at the mean flowpath radius R1 are channeled through the strut members 32 to the LPT 36. The LPT 36 is effective for expanding the combustion gases through the first and second turbine blade rows 46 and 50 along the mean flowpath radius R2 for extracting substantially all output power from the gases for driving the first and second rotors 38 and 48 in counterrotating directions at rotational speeds relatively lower than those of the first drive shaft 28.

The gas generator 16 and the LPT 36 as above arranged and described results in a new and improved gas turbine engine having counterrotating 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 HPT 24 is disposed aft of the combustor 22 for first receiving the relatively high pressure combustion gases being discharged therefrom. The HPT 24 is most efficient when it and the first 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 HPT 24 are at a reduced, intermediate pressure. The intermediate pressure gases then flow through the IPT 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 at speeds relatively lower than those of the HPT 24.

Finally, the low pressure combustion gases are channeled to the LPT 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 which is discharged from the LPT 36. Furthermore, inasmuch as the LPT 36 is the last element in the engine 10, it is subject to the lowesttemperation combustion gases and therefore, thermally induced stresses are reduced allowing for a less complex LPT 36.

For more efficiently extracting energy from the combustion gases in the LPT 36 it is preferable that the mean flowpath radius R2 thereof be greater than the mean discharge radius R1 of the gas generator 16. In the embodiment illustrated in Figure 1, the mean flowpath radius R2 is about double the magnitude of the mean discharge radius R1. 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 thereover.

Inasmuch as the LPT 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 LPT 36 is effective for driving counterrotating, oppositely pitched forward propellers 54 and aft propellers 56. More specifically, extending from an aftmost end of the first rotor 38 is an aft blade row 46a which extends radially outwardly to about the radial position of the second rotor 48. Attached to radially outer ends of the aft blade row 46a is an annular shroud member 58. The aft propellers 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 provided 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 LPT 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 (not shown) includes a single rotor typically rotating at about 10,000 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 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 1 of the LPT 36 to rotate in a direction opposite the first rotor 38, two output shafts, shaft, first rotor 38 and second rotor 48, are provided which rotate at about one quarter the speed of an analogous single rotor, conventional LPT, thus providing speed reduction.

Furthermore, additional speed reduction is obtainable by increasing the number of the first and second turbine blade rows 46 and 50, i.e., the number of stages. This is so in that at lower rotational speeds of the rotors 38 and 48 less energy can be extracted from the combustion gases per stage of the LPT 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/R1 for providing the combustion gases to the LPT 36 at a larger mean flowpath radius R2. Too many stages is undesirable because of the increased complexity, size and weight therefrom, and an LPT 36 having fewer stages and a relatively high R2/R1 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/R1 ratio of about 2.0 is preferable.

Furthermore, in the embodiment illustrated in Figure 1 for driving the counterrotating propellers 54 and 56, the LPT 36 having about 14 stages is preferred for obtaining output shaft speeds of the first and second rotors 38 and 48 of about 1200 RPM.

This speed is much less than the rotational speeds of the first and second drive shafts 28 and 30.

In the embodiment illustrated in Figure 1, the counterrotating 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 R4from the longitudinal axis 12. In the embodiment of the engine 10 including an LPT 36 driving propellers and having about 14 stages, it is also preferred that R1/R4, R2/R4, and R3/R4, equal about 0.18, 0.35, and 0.45, respectively. However, the number of stages of the LPT36 can range between about 10 and about 18 stages, and R1/R4, R2/R4 and R3/R4 can range between about 0.2 to 0.16, 0.4 to 0.3, and 0.5 to 0.4, all respectively.These relationships are preferred for obtaining an engine 10 suitable for most efficiently driving the counterrotating propellers 54 and 56 at rotational speeds of about 1200 RPM.

The reduction in speed of the rotors 38 and 48 of the LPT 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 LPT 36 requires less material foraccommo- dating centrifugal stress which results in a lighter LPT 36. The overall effect of using a counterrotating LPT 36 is a significant reduction in engine weight as compared to an engine including a conventional LPT 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 thereover.

The use of two propellers over a single propeller allows for propellers of lesser diameter, for example about 12 feet, i.e. R4 = 6 feet, versus about 16 feet, respectively, for generating an equivalent amount of thurst 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.

Mounting the propellers 54 and 56 radially outwardly of the second rotor 48 increases the hub to tip ratio R3/R4 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 LPT36, 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 counterrotating propellers, such as the one illustrated in Figure 1, mounted to an aftmost end of the aircraft 66. Aft mounted counterrotating 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 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 counterrotating 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 LPT 36 in the engine 10 but is also effective for mounting the entire engine 10 to a wing 68 of an aircraft.

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 ofthe engine 10 of Figure 1. In this embodiment, however, a LPT 74 drives counterrotating, forward and aft fans 76 and 78, respectively, mounted to an aft end of the engine 72. The fans 76 and 78 include a plurality of radially outwardly extending and circumferentially 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 14 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 LPT 74, although basicaily identical to the LPT 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, R1/R4 and R2/R4 preferably have values between about 0.35 to about 0.25 and 0.65 to about 0.45, respectively. However, for 8 stages, values of R1/R4 and R2/R4 of about 0.3 and 0.58, respectively, are preferred. As in the embodiment illustrated in Figure 1, it is preferred that R2 have a larger value than R1 and preferably a value twice as large.

Illustrated in Figure 5 is a gas turbine engine 84 according to another embodiment of the present invention. The engine 84 includes a gas generator 16 which is substantially identical to the one illustrated in Figure 1.The engine 84 also includes an LPT 86 which is substantially identical to the LPT 74 illustrated in Figure 4. However, in this embodiment, the LPT 86 preferably includes an additional, aftmost blade row 50c for a total of 9 stages, which stages are arranged for driving counterrotating forward and aft fans 88 and 90, respectively, rotatably mounted to a forwardmost end of the engine 84. Disposed radially outwardly of the fans 88 and 90 is an annular fan duct 92 suitably secured by struts 94 to the engine 84.

In contrast to the LPT 74 illustrated in Figure 4, an aftmost end 96 of the first rotor 38 extends radially inwardly of the hub member 34 and is fixedly attached to a third annular driveshaft 98 which extends to the forward end of the engine 84 and is suitably attached to the aft fan 90. The aftmost blade row 50c extends radially inwardly from the second rotor 48. Radially inner ends 100 of the aftmost blade row 50c are fixedly attached to a fourth driveshaft 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 coaxially mounted driveshafts 28, 30, 98 and 102, with the LPT 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 driveshaft, the third driveshaft 98 which is fixedly connected to the first rotor 38 of the LPT 86 and to the disc rotor of the IPT 26.

Illustrated in Figure 7 is a gas turbine engine 106 according to another embodiment of the present invention. This embodiment includes an LPT 108 which is substantially identical to the LPT 36 of Figure 1 that includes 14 stages. However, the LPT 108 is arranged similarly to the LPT 86 of Figure 5 including theadditional blade row 50c for a total of 15 stages and including the third and fourth driveshafts 98 and 102. The driveshafts 98 and 102 are effective for driving counterrotating forward and aft variable pitch propellers 110 and 112, respectively, rotatably mounted to the forwardmost end of the engine 106.

In this embodiment, one, or a plurality of gas generators 114 are arranged for driving the LPT 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 thatthe 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 1 to the LPT 108 for providing combustion gases thereto. In this embodiment, one or more gas generators 114 can be mounted circumferentially about and parallel to the longitudinal axis 12 of the engine 106 for providing combustion gases to the LPT 108 for driving the counterrotating 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 IPT 26 can also be used for generating combustion gases. Furthermore, inasmuch as the counterrotating LPT 36 is effective for providing relatively large output power and torque at low speeds, gas turbine engines incorporating such LPTs can be used for powering ships, generators, and large pumps, for example, which can be designed for having counterrotating input shafts suitably attached to the first and second rotors 38 and 48 of the LPT 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, in a smaller, 1500 shaft horsepower engine, powering shorter propellers 54 and 56, the HPT 24 wou Id be designed to operate at about 30,000 RPM. The first rotor 38 and the second rotor 48 of the LPT 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

1. Agasturbine 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; said power turbine being effective for receiving said combustion gases and extracting substantially all output power therefrom for driving said first and second rotors in counterrotating directions.

2. A gas turbine engine according to claim 1 further including: an annular casing disposed circumferentially about said gas generator; and an annular support member extending radially inwardly and in an aft direction from an aft end of said casing; said first and second rotors of said power turbine being rotatably supported at radially inner ends to said support member.

3. A gas turbine engine including a longitudinal axis comprising: an annular casing disposed coaxially about said axis; an annular support member extending radially inwardly and in an aft direction from an aft end of said casing; 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 longitudinal axis and in an aft direction; and a power turbine disposed coaxially about said longitudinal axis and including: a first drum rotor rotatably attached to said support member; a plurality of first turbine blade rows extending radially outwardly from said first rotor and spaced axially thereon; a second drum rotor rotatably attached to said support member 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.

4. A gas turbine engine according to claim 3 wherein said mean flowpath radius of said power turbine is greater than said mean discharge radius of said gas generator.

5. A gas turbine engine according to claim 4 wherein said mean flowpath radius is about double the magnitude of said mean discharge radius.

6. A gas turbine engine according to claim 3 wherein said support member comprises a plurality of circumferentially spaced strut members extending radially inwardly from said aft end of said casing and an annular hub member fixedly attached to radially inner ends of said strut members and extending in an aft direction, said strut members being effective for supporting said hub member and for channeling combustion gases from said gas generator to said power turbine, said hub member being effective for supporting said first and second rotors of said power turbine.

7. A gas turbine engine according to claim 3 wherein said second drum rotor is rotatably mounted to said support member at radially inner ends of a forwardmost blade row of said second blade rows and at radially inner ends of an aftmost blade row of said second blade rows, said aftmost blade row being rotatably disposed on said first drum rotor mounted to said support member.

8. A gas turbine engine according to claim 3 wherein said gas generator is disposed coaxially with said longitudinal axis and said power turbine.

9. A gas turbine engine according to claim 3 wherein said gas generator includes a longitudinal centerline axis being disposed parallel to and spaced from said longitudinal axis of said gas turbine engine.

10. A gas turbine engine according to claim 3 wherein said first rotor and said second rotor are effective for driving first and second counterrotating airfoil members.

11. A gas turbine engine according to claim 10 wherein said airfoil members comprise propellers.

12. A gas turbine engine according to claim 10 wherein said airfoil members comprise fans.

13. A gas turbine engine according to claim 10, wherein said airfoil members are disposed at a forward end of said gas turbine engine and forward of said gas generatorthereof.

14. A gas turbine engine according to claim 10 wherein said airfoil members are disposed at an aft end of said gas turbine engine and adjacent to said power turbine.

15. A gas turbine engine according to claim 3 wherein said first and second rotors of said power turbine are effective for driving first and second counterrotating propellers and a total number of said first and second turbine blade rows is less than about 18 rows and more than about 10 rows.

16. A gas turbine engine according to claim 3 wherein: said first and second rotors of said power turbine are effective for driving first and second counterrotating propellers, said propellers having a tip radius and a hub radius; and said mean discharge radius of said gas generator, said mean flowpath radius of said power turbine and said hub radius of said propellers having magnitudes with respect to said tip radius of said propellers of between about 0.2 to about 0.16, 0.4 to about 0.3, 0.5 to about 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.

17. A gas turbine engine according to claim 3 wherein said first and second rotors of said power turbine are effective for driving first and second counterrotating fans and a total number of said first and second turbine blade rows is less than about 12 rows and more than about 6 rows.

18. A gas turbine engine according to claim 3 wherein: said first and second rotors of said power turbine are effective for driving first and second counterrotating fans, said fans having a tip radius; and said mean discharge radius of said gas generator and said mean flowpath radius of said power turbine having magnitudes with respect to said tip radius of said fans of between about 0.35 to about 0.25 and 0.65 to about 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, respectively.

19. A gas turbine engine according to claim 3 further including: a plurality of forward airfoil members attached to and extending radially outwardly from a forward portion of said second rotor; and a plurality of aft airfoil members attached to an annular shroud attached to radially outer ends of an aftmost blade row of said first turbine blade rows and extending radially outwardly therefrom; said power turbine being effective for driving said forward and aft airfoil members in counterrotating directions for generating thrust therefrom, said combustion gases received by said power turbine being exhaustable therefrom in an aft direction and radially inwardly of said forward and aft propellers.

20. A gas turbine engine according to claim 19 wherein said airfoil members comprise propellers.

21. A gas turbine engine according to claim 19 wherein said airfoil members comprise fans.

22. A gas turbine engine according to claim 3 wherein an aft blade row of said plurality of second turbine blade rows comprises an aftmost blade row of said power turbine and includes radially inner ends attached to a third rotor, said third rotor disposed radially inwardly of said drive shaft and being effective for driving a forward fan disposed upstream of said gas generator; and said second rotor being effective for driving an aft fan disposed between said first fan and said gas generator.

23. A gas turbine engine according to claim 3 wherein said gas generator further includes a booster compressor disposed forward of said compressor and an intermediate pressure turbine disposed downstream of said high pressure turbine, said intermediate pressure turbine being effective for receiving said combustion gases from said high pressure turbine and driving said booster compressor for providing pressurized air to said compressor.

24. A gas turbine engine including a longitudinal axis comprising: an annular casing disposed coaxially about said axis; an annular support member extending radially inwardly and in an aft direction from an aft end of said casing; 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 longitudinal axis and in aft direction; and a power turbine disposed coaxially about said longitudinal axis and including: a first drum rotor rotatably attached to said support member; a plurality of first turbine blade rows extending radially outwardly from said first rotor and spaced axially thereon, an aft blade row of said plurality of first turbine blade rows comprising an aftmost blade rowofsaid powerturbine; a second drum rotor rotatably attached to said support member and disposed radially outwardly of said first rotor and said first turbine blade rows; 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; a plurality of forward propellers attached to and extending radially outwardly from a forward portion of said second rotor; and a plurality of aft propellers attached to an annular shroud attached to radially outer ends of said aftmost blade row and extending radially outwardly therefrom; said power turbine disposed coaxially about a longitudinal axis of said gas turbine engine and 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 opposite, counterrotating directions at speeds relatively lower than those of said drive shaft for generating thrust, said combustion gases received by said powerturbine being exhaustable therefrom in an aft direction and radially inwardly of said forward and aft propellers.

25. A gas turbine engine substantially as hereinbefore described with reference to any of Figures 1 and 4 to 7, or an aircraft as hereinbefore described with reference to Figure 2 or Figure 3 of the drawings.