JPS59103947A - Reversely rotating power turbine - Google Patents

Abstract

(57) [Summary] This bulletin contains application data before electronic filing, so abstract data is not recorded.

Description

DETAILED DESCRIPTION OF THE INVENTION Field of the Invention This invention relates to gas turbine engines, and more particularly to counter-rotating rotors or The present invention relates to a new and improved gas turbine engine including a power turbine having counter-rotating rotors.

[Blue Application of the Invention] This invention has particular but not limited use in gas turbines such as those used for aircraft propulsion.

Several types of gas turbine engines are currently available to power aircraft. Turbofans and turboprops are two examples of such engines.

A turbofan engine includes a core engine, or gas generator, that powers the fan. Turboprop engines, on the other hand, include a gas generator that powers a propeller.

Because these engines drive a propeller or fan to generate thrust, they are more fuel efficient at subsonic speeds than pure turbojet engines that generate thrust only through the exhaust jet. is typical.

Medium-sized transport aircraft, e.g. 100 to 180 passengers
Aircraft used to transport people typically use turbofan engines for propulsion. Turbofans generate a relatively large amount of thrust to provide the necessary power for these aircraft at relatively high altitudes and cruise speeds of about 0.6 to about 0.8. Aircraft designed for lower cruising speeds may use conventional turboprops (in-type) because they offer superior performance and efficiency. For example, turbofans By using turboprops, which are aerodynamically more efficient, the fuel burn m, or fuel consumption per passenger mile, can be significantly reduced.

Therefore, by combining the advantages of a turbofan and a turbob [TI hub,
It would be desirable to improve efficiency at within-type aircraft cruise speeds.

However, it is not possible to simply oversize an ordinary turboprop engine suitable for powering a medium-sized transport aircraft at the cruise speeds and altitudes typical of turbofan-powered aircraft. , one propeller would need to be about 16 feet in diameter. About 15.0 more
It is required to be able to generate 00 shaft horsepower, which is several times the power output of a normal turboprop engine.

Configuring a conventional turbo-bobbed engine for these conditions would result in the development of a relatively large and undesirably heavy reduction gearing system to transmit the required power and torque to the propeller at relatively low speeds. It becomes necessary. The rotational speed of the large diameter propeller becomes the limiting factor in keeping the helical speed of the propeller tip, ie, the speed of the aircraft plus the tangential speed of the propeller tip, below supersonic speed. This is desirable because propeller tips operating at supersonic speeds generate significant undesirable noise and result in reduced aerodynamic efficiency.

Gas turbine engines that operate to drive propellers or fans without reduction gearing are also known in the art. Typically, this includes a relatively low speed counter-rotating spindle rotor with a relatively small number of stages of rotor blades driving a pair of counter-rotating fans or propellers. These engines come in various forms that utilize fans or propellers simply to augment the thrust generated by the exhaust jets.

However, to propel today's medium-sized aircraft, which require relatively large power outputs, they require more performance than conventional turbofan and turboprop engines, or the counter-rotating turbine-rotor engines just mentioned. What is needed is a new 111-generation practical model with significantly improved fuel efficiency and relatively good fuel efficiency.

Accordingly, one object of the present invention is to provide a new and improved gas turbine engine.

Another object of the invention is to provide a new and improved gas turbine engine that includes a power turbine with counter-rotating rotors.

Another object of the invention includes a power turbine having stages of a plurality of counter-rotating turbine rotor blades, wherein substantially all of the output power is obtained by expanding combustion gases in the stages. It is an object of the present invention to provide a new and improved gas turbine engine in which substantially little energy remains in the exhaust gases leaving the engine.

Another object of the invention is to provide a new and improved gas turbine engine which provides a power horn without the use of reduction gearing.

Another object of the invention is to provide a new and improved system comprising a gas generator and a power turbine having counter-rotating rotors, the power turbine being fixedly supported aft of the gas generator. Its purpose is to provide gas turbine engines.

Another object of this invention is to provide a new and improved gas turbine engine for powering counter-rotating blade members such as propellers and fan blades. It is to offer.

SUMMARY OF THE INVENTION The present invention provides a new and improved gas turbine engine that includes a gas generator and a power turbine.

A power turbine includes a first rotor and a plurality of first turbine rotor rows extending radially outwardly therefrom, a second rotor and a plurality of second rows extending radially inwardly therefrom.
It includes a row of turbine rotor blades. A power turbine is supported behind the gas generator, receives combustion gas from the gas generator, and directs the waste to the first and second turbines.
Inflate it in the row and almost all of it comes out - 1! C1 acts to drive the first and second rotors in opposite directions.

In some embodiments of the invention, the power turbine acts to drive counter-rotating fans or propellers located at either the front or rear end of the engine.

The structure, other objects, and advantages of the present invention will be better understood from the following detailed description of the drawings.

[Detailed Description of the Invention] A gas turbine engine 10 according to an embodiment of the invention is not shown in FIG. The engine 10 includes a longitudinal axis 12 serving as a centerline, and an annular casing 14 is disposed coaxially around the axis 12. The engine 1o is, for example, a booster compression 1fi1
8. Compressor 2o, combustor 22, high pressure turbine (IP)
T) 2'4, and a conventional gas generator 16, which may consist of an intermediate pressure turbine (IPT) 26, all of which are coaxial in series axial flow around the longitudinal axis 12 of the engine 10. It is located at -C. A first annular drive shaft 28 connects the compressor 20 and the shaft 24 . A second annular drive @23 couples booster compression 118 and IP-C26.

In operation, the gas generator 16 supplies pressurized air from the booster compressor 18 and compressor ts2'o to the combustor 22, where it is mixed with fuel and suitably ignited to produce carbon dioxide.
It acts to generate combustion gas.

This combustion gas is 1-11) 1' 24 and IP-r
'26, which drive compressor 20 and booster compressor 18, respectively. The combustion gas is discharged from the gas generators 16 to 11' at an average discharge radius R1 from the longitudinal axis 12.
It is discharged via T26.

An annular support member 30 is attached to the rearmost end of the casing 14 behind the gas generator 16 .

Support portion 30 extends radially inwardly and rearwardly from the rear end of casing 14 . The support member 30 is attached to the casing 1
4 includes a plurality of circumferentially spaced apart strut members 32 extending radially inwardly from the rear ends of the strut members 32, and an annular hub member 34 secured to the radially inner ends of the strut members 32 and extending rearwardly.

The branch ttA32 supports the hub member 34 and has a passageway for supplying combustion gases from the gas generator 16 to a power turbine 36 constructed in accordance with one embodiment of the present invention.
0 works. Power turbine 36 or low pressure turbine (
L l]1') 36 is rotatably mounted on the hub member 34.

L I) T 36 is attached to the hub member 34 at its forward end 42.
and a first annular drum-shaped rotor 38 rotatably mounted at a rear end 44 by suitable bearings 40 . Extending radially outward from the first rotor 38 are a plurality of first rows of turbine rotor blades 46 that are axially spaced apart on the rotor.

LPT 36 also includes a first rotor 38 and a second annular drum-shaped rotor 48 disposed radially outwardly from first rotor 38 and first row of rotor blades 46 . Extending radially inwardly from the second rotor 48 are a plurality of second rows of turbine rotor blades 50 that are axially spaced apart on the rotor. The second rotor 48 is located at the radially inner end of the frontmost rotor blade row 50a of the second rotor blade row 50 and at the radially inner end of the rearmost rotor blade row 50b. It is rotatably mounted to the hub member 34 by suitable bearings 52 located therein. The rearmost rotor blade row 5ob is a hub member 34
The first rotor 38 is rotatably arranged on the IC first rotor 38.

Each of the first and second turbine rotor blade rows 46,50 is composed of a plurality of turbine rotor blades spaced apart in the circumferential direction, and the first rotor blade row 46 and the second rotor blade row 50 are They are arranged alternately and separated from each other. Combustion gas flowing along the rotor blade row 46.50 flows along the average flow path radius R2. This average flow path radius, by definition, represents the rotor blade radius around which the resultant work load of the LPT 36 is assumed to be concentrated. for example,
The radius R2 can be defined as the average pitch line radius of all rotor blade rows of the LPT 36.

Combustion gas discharged from the gas generator 16 at a point with an average flow path radius (1) is sent to the LPT 36 via a flow path formed by the pillar portion U32. -C1 along the radius [2] expands this combustion gas in the first and second rows of turbine blades 46, 50 so as to extract substantially all of the output energy of this gas;
The first and second rotors 38,48 are driven in opposite directions at rotational speeds that are relatively lower than the speed of the first drive shaft 28.

Gas light generator 16 and LPT 3 configured as described above
6 provides a new and improved gas turbine engine with counter-rotating rotors that act to provide output shaft power at relatively low rotational speeds. An important feature of the invention is the complementary arrangement of the engine elements.

Specifically, Hl]T 24 is located aft of combustor 22 to initially receive relatively high pressure combustion gases after they are discharged. HP'T 24 is the most powerful when this turbine and first drive shaft 28 are set to rotate at approximately io, ooo to 15,000 RF) M in a 15,000 shaft horsepower engine. and efficient,
The high pressure combustion gas from the combustor 22 is utilized most efficiently.

After the combustion gases pass through the IPT 24, they are at a lower intermediate pressure. This intermediate pressure saw gas passes through IP 126, which further reduces the pressure of the gas to a relatively low pressure, while the second drive shaft 23 and booster compression To extract the power for rotating ti18 most efficiently.

Most importantly, the low pressure combustion gases are sent to 1 F) T 36 where they are further expanded and substantially all of their remaining energy is extracted to rotate the first and second rotors 38.48; Provides output shaft power. 11) Very little energy remains in the exhaust jet discharged from T2O, thus reducing its utilization efficiency. Furthermore, LPT 36 is engine 1
Since it is the last element in 0, the combustion gases it experiences are the lowest in temperature and therefore the thermally induced stresses are also reduced, allowing Ll) T'36 to become less complex. .

In order to more efficiently extract 1 energy from the combustion gas with LPI' 36, the average flow path radius R2 is the same as that of the gas generator 16.
It is preferable that the radius is larger than the average discharge radius R1. In the embodiment shown in FIG. 1, the average flow path radius R2 is the average discharge radius R
The human body 218 has a size of 1. This arrangement places the turbine rotor blade rows 46,50 at a large radius from the longitudinal axis 12 to increase their relative tangential velocity and more efficiently extract energy from the gas flowing therethrough. It is effective for

The LPT 36 is a power turbine that acts to provide substantially all of the output power to the rotor 38, 48, and is preferably located aft of the gas generator 16, thus allowing for effective and suitable mounting. equipment is required. The support member extending from the rear end of the casing 14, as discussed above, is therefore also an important feature of the invention.

In the embodiment shown in FIG. 1, the LPT 36 acts to drive oppositely rotating forward propellers 54 and aft propellers 56 with opposite pitches.

More specifically, an aft rotor blade row 46a extends from the aft-most end of the first rotor 38 and extends radially outward to approximately the radial position of the second rotor 48. It extends in the direction. An annular shroud member 58 is attached to the radially oriented end of the rear rotor blade row 4.6a. The rear propeller 56 is suitably mounted on the shroud member 58. Similarly, the front propeller 54 is suitably mounted on the front end of the second rotor 48.
o to independently control the pitch of the front and rear propellers 54,56.

The most important feature of the invention is that the gas turbine engine 10 operates to provide relatively large output power and torque at relatively low rotational speeds without the use of reduction gearing.
It is to have PT 36. Reduction gearing and associated accessories add considerable weight and complexity to engines that can generate the relatively large thrusts required to operate a transport aircraft, such as a 150-passenger transport aircraft. will join.

When using a gas turbine engine to drive an engine member such as a propeller or fan, deceleration is necessary. A common low pressure turbine (not shown) typically has approximately 10.0
Contains one rotor that rotates from 'OO to 15.00 ORPM. In order to drive the W member, this rotational speed must be approximately 1
．． 000 to a relatively low speed of about 2,000 RPM. Propellers and fans are designed to move relatively large amounts of air at relatively low shaft speeds to generate thrust, and are more efficient when operating at relatively low rotational speeds. Additionally, low rotational speeds are required to limit the propeller tip helical speed to below supersonic speeds.

In this invention, the second part of L l) 1' 36 in FIG.
The two output shafts, namely the first rotor 38 and the second rotor 48, are covered so that they can rotate in the opposite direction to the first rotor 38. , a conventional rotor with one rotor similar to this +) approximately 1 of the speed of H
It rotates at a speed of /4 and decelerates in this way.

Furthermore, the number of first and second turbine rotor blade rows 46.50;
That is, by increasing the number of stages, additional deceleration can be achieved. This is because when the rotational speed of the rotor 38.48 is low, less energy is extracted from the combustion gases per stage of the LPT 36. In order to achieve the desired reduced speed and extract substantially all of the remaining energy from the combustion gases, it is necessary to increase the number of stages.

However, by increasing the value of the ratio R2/R1, the combustion gas is supplied to the LPT 36 at a point with a larger average flow path radius R2.
By doing so, fewer stages can be used to achieve the above objectives. It is undesirable to increase the number of stages due to the increase in complexity, weight and weight. LPT36, which has a small number of stages and a relatively large ratio R2/R1, is also desirable because it increases the frontal area and weight. do not have.

In this invention, as mentioned above, a ratio of about 2.0 [<2/
It has been found that R1 is preferred.

Furthermore, in the embodiment shown in FIG.
4.56, the first and second rotors 38.4
To obtain an output shaft speed of approximately 1.200 RPM (7) to 8,
Ll) 136 having approximately 14 stages is preferred. This speed is
Much lower than the rotational speed of the first and second drives @23.28.

In the embodiment, which is not shown in FIG.
6 is mounted at the rear of the engine 1o, radially outward from 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. Example aircraft pA10 having approximately 141 stages and having an LPT36 driving the propeller.
Then, it is preferable that Be1/R4, R2/R4 and R3/R4 are equal to 0°18, 0.35 and 0.45, respectively. However, the number of stages of LPT36 is about 10 stages to about 18 stages.
R1/F<4. R2/v<4 and R
3/R4 is 0.2~0.16.0°4~0.
3 and can range from 0.5 to 0.4.

This relationship is preferred in order to obtain an engine 10 suitable for most efficiently driving counter-rotating propellers 54,56 at rotational speeds of approximately 1:2oo RPM.

As the speed of the rotor 38,48 of the LPT 36 decreases, a quadratic decrease in the centrifugal force generated stress occurs. for example,
When the speed decreases to 1/4, the stress due to centrifugal force decreases by 1/4.
It has decreased to 16. This is important because L P T2C requires less material to withstand the stress of centrifugal force.
This results in a lighter LPT36. Opposite LPT
The overall effect of using 36 is that the engine's gravity is significantly reduced compared to an engine containing a conventional L P l' and reduction gearing.

The engine 10 of the embodiment shown in FIG. 1 also has other advantages. For example, by placing the propellers 54,56 at the rear end of the engine 10, the annular inlet region 62 of the engine 10 is relatively free of flow obstructions. Therefore, the annular nacelle 64 surrounding the inlet area hfi 62 as well as the engine 10
Is it appropriately set so as to improve the aerodynamic performance of the air entering the engine 10 and the air flowing there? I can do it.

By using two propellers instead of one propeller, an equivalent 1 at a rotational speed of 1.20 ORPM and 900 Rl) M at an aircraft cruise speed of about 0.7 to about 0.8, respectively. To generate power, it can be attached to a smaller diameter propeller. For example, instead of 16 feet for one piece, -C is about 12 feet, that is, R4
-Can climb 6 feet. As a result of the smaller diameter, the tip speed of the blade is reduced and its noise is also reduced.

By mounting the propellers 54 and 56 radially outward of the second rotor 48, the propeller's hub-to-tip ratio R3/R4 is increased, which improves its aerodynamic performance. Furthermore, B[lpera is L l) T 3
．． Do not obstruct the flow of combustion gases discharged from propeller 54;
A cooling system would be required to prevent damage from +56 heat.

In Figure 2, the device attached to the rearmost end of Aviation 166 is shown.
As shown in Figure 1, the oppositely rotating propeller is
An aircraft 66 is shown including two engines 10 . The counter-rotating propeller engine 10 of this invention mounted on the rear side is as follows:
It acts to improve performance as well as fuel combustion to create an Ic aircraft 66. Additionally, engine 10 has a reduced weight when compared to a conventional one-way turboprop engine producing the same thrust power. It is possible to reduce propeller noise, thereby reducing the scale of noise attenuating modifications to the aircraft, and therefore -
This would further reduce the overall weight of the C aircraft.

Figure 3 shows a counter-rotating propeller engine 1 as shown in Figure 1.
An alternative configuration is shown for attaching G10 to a wing 68 of an aircraft (not shown). In this embodiment, the hub member 34 of the engine 10 is extended rearwardly and the 9! 68 (t G). A stationary annular exhaust duct 70 is suitably secured to the hub member 34 and carries exhaust gases from the engine 10, e.g.
Squeeze it appropriately to the bottom of 68. Engine 1 of the first embodiment shown in FIG.
0 clearly shows the important advantages of the support member 30 of the engine 10. Specifically, the support member 30 is L P
1-36 in the engine 10
``Instead of taking the whole engine 10 to the aircraft development 68'']
It also has an effect.

FIG. 4 shows a gas turbine engine 7 according to another embodiment of the present invention.
2 is shown -C. Engine 72 includes gas generator 16 . This gas generator is substantially identical to the gas generator 16 of the I warmer 10 of FIG. However, in this example, L
PT 74 drives oppositely rotating front and rear fans 76 and 78, respectively, mounted at the rear end of engine 72. Fans 76, 78 include a plurality of circumferentially spaced and radially outwardly extending fan blades. Annular fan duct 8
0 is disposed radially outward from the fans 76 and 78, and is appropriately attached to the casing 14 and the nacelle 64 of the engine 72 by a plurality of struts 82. Hub member 34 may be fitted with suitable thrust reversal means (not shown).

Fan blades have a different action from propeller blades, so L
PT 74 is basically the same as LP-136 of FIG. 1, but is preferably designed to drive the fan blades. More specifically, the total number of stages δ1 of the first and second turbine rotor blade rows 45.50 is about 6 stages to about 1 stage.
A range of two stages is preferred, with about eight stages (as shown in FIG. 4) being preferred. Correspondingly, R
1/1 dog 4 and I' (2/R4 preferably have values of 0.35 to 0°25 and 0.65 to about 0.45, respectively. However, in the case of 8 stages, R1/R4 and r<
2/R4 preferably have values of 0.3 and 0.58, respectively. As in the embodiment shown in FIG.
has a larger value than R1, preferably twice as large.

FIG. 5 shows a gas turbine engine 84 according to another embodiment of the invention. Engine 84 includes gas generator 16, which is substantially identical to that shown in FIG. Engine 84 also includes an LPI 86, which is substantially identical to LPT 74 shown in FIG. However, in this embodiment, it is preferred that L P T2C include an additional aftmost rotor blade row 50c, for a total of eight stages. These stages are configured to drive opposing front and rear fans 88, 90 which are rotatably mounted at the forward-most end of engine 84. Fan 88°90
A partial fan duct 92 is disposed radially outwardly of -C and is suitably secured to engine 84 by struts 94.

In contrast to L P T 74 shown in FIG. It is fixed. The drive shaft 98 extends to the front end of the engine 84 and is appropriately attached to the rear fan 90.The rearmost moving image train 50c extends radially inward from the second rotor 48. The radially inner end 100 of the rearmost rotor blade row 500 is fixed to a fourth drive shaft 102, and this drive shaft extends to the front end of the engine 84 and is fixed to the front fan 88. For this reason, the engine 84 has four
one drive shaft 23.28.98. '102 included, LPT
86 act to drive the front and rear [7] pins 88 and 90 in opposite directions, respectively. The engine 84 thus constructed is capable of very high shunting ratios of greater than about 6 to 1.

FIG. 6 shows a gas turbine engine 1 according to another embodiment of the present invention.
04 is shown. This embodiment is substantially the same as the embodiment shown in FIG. The drive shaft 98 is driven by the first rotor 38 of the LPT 86 and the IPI-
It is fixed to a circular rotor of 26.

FIG. 7 shows a gas turbine engine 1 according to another embodiment of the present invention.
06 is shown. The LP'l' 108 in this embodiment is substantially identical to the LPI-36 of FIG. 1 with fourteen stages. However, L P T' 108 is an additional video sequence 50
Including C, the structure is similar to L P T2C in FIG. 5, and there are 15 stages in total, and the third and fourth drive shafts 98
．． Including +02. Drive shaft 98, 102 HaPs section 10
6 (7) Frontmost (7) GmkJljl Rotatably mounted front and rear variable pitch wheels act in opposite directions to drive the propellers 110 and 112, respectively.

In this embodiment, one or more gas generators 114 are provided to drive L'PT 108. Gas generator 114 is substantially identical to gas generator 16 of FIG. 1 and includes a longitudinal axis 116 which is the C1 centerline. However, in contrast to that shown in FIG. 1, gas generator 114 is mounted such that its longitudinal axis 116 is parallel to, but spaced from, the longitudinal axis of engine 106. Appropriate annular duct +-
118 fluidly connects gas generator 114 to LPT 108 to provide combustion gas to C1LPH. In this embodiment, one or more gas generators 114 are connected to the engine 1.
06, parallel to and circumferentially around the longitudinal axis 12, Ll)T 108 can be supplied with combustion gases to drive counter-rotating propellers 110, 112.

Although what is presently considered to be the preferred embodiment of the invention has been described, other embodiments will be readily apparent to those skilled in the art from the foregoing description.

For example, combustion gas can be generated using the gas generator 16 of FIG. 1 without the booster compressor 18 and IPT 26. Furthermore, since the counter-rotating L P l' 36 acts to provide relatively large output power and torque at low speeds, gas turbine engines using this L P T are suitable for use in ships, generators, and large pumps, for example. Such devices may be arranged in stages with oppositely rotating input shafts suitably mounted on the first and second rotors 38, 48 of the LPT 36. I can do it.

Furthermore, although the case where this invention is applied to a 15,000-shaft horsepower engine has been described, this invention can also be scaled up to a 15,000 shaft horsepower engine. For example, a smaller 1,500
Shaft horsepower engine, shorter propeller 54. ．． 56, the HPT 24 is approximately 30. OOORP
Designed to work with M. The first part of LPT36 in FIG.
rotor 38 and second rotor 48 (operating at a correspondingly reduced speed of about 10:1, i.e., about 300 ORPM).
-Design so that Propeller 54.56 is approximately 3.00
OR1) M operates, but the tip radius (<4) decreases, so the helical tip speed can be kept lower than supersonic speed.

[Brief explanation of drawings]

FIG. 1 is a sectional view of a gas turbine engine according to one embodiment of the present invention, which includes a power turbine mounted on the rear side and having a rotor rotating in opposite directions and acting to drive and cover a propeller rotating in opposite directions. , FIG. 2 is a schematic diagram of an aircraft equipped with two gas turbine engines as shown in FIG.
Figure 4 shows an alternative configuration of a similar gas turbine engine mounted on an aircraft wing, and is a 91 degree schematic illustration of this engine, which includes a power turbine acting to drive a counter-rotating fan mounted on the rear side. A cross-sectional view of a gas turbine engine according to another embodiment of the invention,
FIG. 5 is a cross-sectional view of a gas turbine engine according to another embodiment of the invention including a power turbine operating to drive a counter-rotating fan mounted on the front side; FIG. FIG. 7 is a cross-sectional view of a gas turbine engine according to another embodiment of the present invention having a common drive shaft with a front-mounted fan and a rotating shaft of a power turbine, with an annular gas generator parallel to the longitudinal axis of the engine. FIG. 2 is a cross-sectional view of another embodiment of a cast turbine engine of the present invention, which includes a power turbine acting to drive a counter-rotating propeller mounted on the front side, spaced apart from the engine; . Explanation of main symbols 14: Annular casing, 16: Cas generator, 18: Booster compressor, 20: Compressor, 22: Combustor, 2
3.28: drive shaft, 24: high pressure turbine, 26: intermediate pressure turbine, 30: annular support member, 32: strut member,
34: Hub member, 36: Power turbine, 38.4
8: Rotor, 46, 50: Rotating blade row, 54.56 + propeller, Patent applicant

Claims (1)

Claims: 1) a first rotor having a gas generator effective for generating combustion gases, a first rotor having a plurality of turbine blade rows extending radially outward therefrom; a second turbine rotor blade row having a second rotor extending radially inwardly therefrom, the power turbine receiving the combustion gases and generating substantially all of the output energy therefrom; A gas turbine engine that operates to extract C1 and drive the first and second rotors in opposite directions. 2. The gas turbine engine according to claim 1) includes an annular casing disposed circumferentially around the gas generator, and extending radially inward and rearward from a rear end of the casing. an annular support member; wherein the first and second rotors of the power turbine are rotatably supported by the support member at radially inward ends thereof. 3) In a gas turbine engine having a longitudinal axis, an annular casing disposed coaxially around the axis, an annular support member extending radially inwardly rearward from a rear end of the casing, and an annular support member extending radially inwardly and rearward from a rear end of the casing; a compressor, a combustor and a high pressure turbine arranged in series with the C1 flow;
The high pressure turbine is adapted to initially receive combustion gases from the combustor and drive the compressor via a drive shaft affixed to the compressor, the high pressure turbine directing the combustion gases generally from the longitudinal axis. a gas generator operative to discharge rearwardly at a discharge radius; and a power turbine coaxially disposed about the longitudinal axis, the power turbine being rotatably mounted to the support member. a first drum-shaped rotor,
a plurality of first rows of turbine rotor blades extending radially outward from the first rotor and spaced apart from each other in the axial direction; rotatably attached to the support member; a second drum-shaped rotor disposed radially outwardly from the first row of turbine rotor blades;
a plurality of second turbine rotor blade rows extending radially inward from the rotor and arranged at alternating distances with the plurality of first turbine rotor blade rows; A turbine receives combustion gas from the gas generator and expands the gas in seven rows of moving first and second turbines along an average flow path radius to generate substantially all of the output energy. A gas turbine engine that extracts the first and second rotors to drive the first and second rotors in opposite directions at a speed relatively lower than the speed of the drive shaft. 4) The gas turbine engine according to claim 3, wherein the average flow path radius of the power turbine is larger than the average discharge radius of the gas generator. 5) The gas turbine engine according to claim 4), wherein the flow passage radius is twice as large as the average discharge radius of a human body. 6) In the gas turbine engine according to claim 3, the support member extends radially inward from the rear end of the casing, and a plurality of strut members spaced apart in the layer resistance direction, and the struts. The strut member comprises a rearwardly extending annular hub fixed to the radially inner end of the member, the strut member supporting the hub member and providing a flow path for conveying combustion residue from the gas generator to the power turbine. A gas turbine engine, wherein the hub portion acts to support the first and second rotors of the power turbine. 7) In the gas turbine engine described in patent '[Claim 3], the second drum-shaped rotor is located radially inside the frontmost rotor blade row of the second row of rotor blades. and the rearmost rotor blade row 1EIJ j, V, rotatably mounted on the support member at the radially inner end of the row, and the rearmost rotor blade V A gas turbine engine, wherein each row is rotatably disposed on a first drum-shaped rotor mounted on the support member. 8) The gas turbine engine according to claim 3, wherein the gas generator is disposed coaxially with the longitudinal axis and the power turbine. 9) A gas turbine engine according to claim 3), wherein the gas generator has a longitudinal centerline parallel to and spaced from the longitudinal axis of the gas turbine engine. 10) In the gas turbine engine according to claim 3), the first rotor and the second rotor act to drive the first and second oppositely rotating blade members. gas turbine engine. 11) The gas turbine engine according to claim 10, wherein the blade member constitutes a propeller. 12. The gas turbine engine according to claim 10, wherein the blade member constitutes a fan. 13) The gas turbine engine according to claim 10, wherein the blade member is disposed at the front end of the gas turbine 1 and in front of the gas generator thereof. 14) The gas turbine engine according to claim 10, wherein the blade member is located adjacent to the power turbine at a rear end of the gas turbine engine. 15) In the gas turbine engine according to claim 3), the first and second rotors of the power turbine act to drive first and second counter-rotating propellers, A gas turbine engine in which the total number of first and second turbine rotor blade rows is greater than about 10 and less than about 18. 16) In the gas turbine engine according to claim 3), the first and second rotors of the power turbine act to drive first and second counter-rotating propellers, The propeller has a tip radius and a hub radius, and the average discharge radius of the gas generator, the average flow path radius of the power turbine, and the hub radius of the propeller are each 0.2 with respect to the tip radius of the propeller. to about 0.16.0
0.5 to about 0.4, and the total number of the first and second turbine rotor blade rows is greater than about 10 rows, respectively, and less than about 18 rows. A gas turbine engine that is compatible with less. 17) In the gas turbine engine according to claim 3), the first and second rotors of the power turbine act to drive first and second counter-rotating fans. and a gas turbine engine in which the total number of the first and second turbine rotor blade rows is greater than about 6 rows and less than about 12 rows. 18) In the gas turbine engine according to claim 111103), the first and second rotors of the power turbine drive first and second counter-rotating fans.
the fan has a tip radius, and the average discharge radius of the gas generator and the average flow path radius of the power turbine are equal to the total number of the first and second rows of turbine rotor blades. 0.3 for the tip radius of the fan, corresponding to more than about 6 rows and less than about 12 rows, respectively.
5 to about 0.25 and 0.65 to about 0.45 (7). 19) In the gas turbine engine according to claim 3), a plurality of front wing members are attached to the front portion of the second rotor and extend radially outward therefrom; A plurality of shrouds are attached to the radially outer end of the aftmost rotor blade row of the first turbine rotor blade row and extend radially outwardly from the annular shroud. an aft wing section I, said power turbine operative to drive said front and aft wing members in opposite directions to generate thrust therefrom, and combustion gases received by said power turbine are directed toward said forward side and a gas turbine engine configured to be capable of discharging air to the rear of the rear propeller and radially inward. 20) In the gas turbine engine according to claim 19, the blade member is a propeller. The gas turbine engine that consists of C. 21) The gas turbine engine according to claim 19, wherein the blade member constitutes a fan. 22. In the gas turbine engine according to claim 3), the rearmost rotor blade row of the plurality of second turbine rotor blade rows is the rearmost rotor blade of the power turbine. a third rotor arranged at the radially inner end of the drive shaft, the third rotor being disposed radially inwardly of the drive shaft; The second rotor operates to drive a front fan disposed further upstream, and the second rotor operates to drive a rear fan disposed between the first fan and the gas generator. Starbin engine. 23) In the gas turbine engine according to claim 3), the gas generator is a booster compressor disposed in front of the compressor (and an intermediate pressure turbine disposed downstream of the high pressure turbine). a gas turbine engine, the intermediate pressure turbine operative to receive combustion gas from the high pressure turbine to drive a booster compressor and supply pressurized air to the compressor; 24) vertical; A gas turbine engine having an axis includes an annular casing disposed coaxially around the axis, an annular support member extending radially inward and rearward from a rear end of the casing, and an annular support member disposed within the casing. , a compressor, a combustor, and a high-pressure turbine in series with respect to the flow, the high-pressure turbine first receiving combustion gas from the combustor and compressing the gas through a drive shaft fixed to the compressor. a gas generator adapted to drive the engine and operative to discharge combustion gases rearwardly at approximately an average discharge radius from said longitudinal axis; and a power turbine disposed coaxially about said longitudinal axis. The movable horn turbine has a first drum-shaped rotor rotatably attached to the support member, extending radially outward from the first rotor and having a gap lC in the axial direction. a plurality of first turbine rotor blade rows, the rearmost rotor blade row of which constitutes the rearmost rotor blade row of the power turbine; a second drum-shaped rotor disposed radially outwardly from the first rotor and the first row of turbine rotor blades; radially inwardly from the second rotor; a plurality of second rows of turbine rotor blades extending from each other and alternately spaced apart from each other of the plurality of first rows of turbine rotors; a) a plurality of forward propellers extending radially outward therefrom; , and a plurality of aft propellers extending radially outwardly therefrom, the rotor blade turbine is disposed coaxially about the longitudinal axis of the gas turbine engine and directs the combustion gases from the gas generator. said first and second turbine motions along a receiving and average flow path radius;
expanding the gas in the train to extract substantially all of its output energy to counterrotate the first and second rotors at a speed relatively lower than the speed of the drive shaft; A gas turbine engine that is driven to generate thrust and that combustion gas received by the power turbine can be discharged rearward from the front side and the rear side 1 [1] radially inside the propeller.