GEARED TURBOMACHINE ARCHITECTURE HAVING A LOW PROFILE CORE FLOW PATH CONTOUR

Abstract

An exemplary geared turbomachine assembly includes a core inlet having a radially inner boundary that is spaced a first radial distance from a rotational axis of a turbomachine, and a compressor section inlet having a radially inner boundary that is spaced a second radial distance from the rotational axis. A ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS

This disclosure claims priority to U.S. Provisional Application No. 61/593230, filed Jan. 31, 2012, and is incorporated herein by reference.

BACKGROUND

This disclosure relates to a geared turbomachine having a core inlet radially spaced from a compressor inlet.

Turbomachines, such as gas turbine engines, typically include a fan section, a turbine section, a compressor section, and a combustor section. The fan section drives air along a core flow path into the compressor section. The compressed air is mixed with fuel and combusted in the combustor section. The products of combustion are expanded in the turbine section.

A core inlet controls flow of air into the core flow path. The flow of air moves from the core inlet to a compressor section inlet. The relative radial positions of the core inlet and the compressor section inlet influence flow through the core and a profile of the turbomachine.

SUMMARY

A turbomachine assembly having a geared architecture according to an exemplary aspect of the present disclosure includes, among other things, a core inlet having a radially inner boundary that is spaced a first radial distance from a rotational axis of a turbomachine, and a compressor section inlet having a radially inner boundary that is spaced a second radial distance from the rotational axis. The ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9.

In a further non-limiting embodiment of the foregoing turbomachine assembly having a geared architecture, the assembly may include a radially inner boundary of the core inlet at the location of a core inlet stator.

In a further non-limiting embodiment of either the foregoing turbomachine assemblies having a geared architecture, the assembly may include a radially inner boundary of the compressor section inlet at the location of a compressor rotor.

In a further non-limiting embodiment of any of the foregoing turbomachine assemblies having a geared architecture, the assembly may include a compressor rotor that is a first stage rotor of a low-pressure compressor.

In a further non-limiting embodiment of any of the foregoing turbomachine assemblies having a geared architecture, the core inlet may be an inlet to a core section of the turbomachine.

In a further non-limiting embodiment of any of the foregoing turbomachine assemblies having a geared architecture, the assembly may have an inlet flow of the compressor section that is from about 30 lb/sec/ft² to 37 lb/sec/ft² when the turbomachine is operating at a cruise speed.

In a further non-limiting embodiment of any of the foregoing turbomachine assemblies having a geared architecture, the assembly may have a turbine inlet temperature of a high-pressure turbine within the turbomachine that is from about 2,000 F to 2,600 F when the turbomachine is operating at a cruise speed.

In a further non-limiting embodiment of any of the foregoing turbomachine assemblies having a geared architecture, the assembly may have a blade in the compressor section having a tip speed during operation that is from about 1,050 fps to 1,350 fps.

In a further non-limiting embodiment of any of the foregoing turbomachine assemblies having a geared architecture, the assembly may include a fan section of the turbomachine driven by a geared architecture that is driven by a shaft that rotates a compressor rotor within the compressor section.

In a further non-limiting embodiment of any of the foregoing turbomachine assemblies having a geared architecture, the assembly may include a geared architecture that has a gear reduction ratio from about 2.2 to 4.

A gas turbine engine assembly according to an exemplary aspect of the present disclosure includes, among other things, a core inlet stator having a stator root that is spaced a first radial distance from a rotational axis of a gas turbine engine, and a compressor blade within a first stage of a compressor section. The compressor blade has a blade root that is spaced a second radial distance from the rotational axis. A ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9.

In a further non-limiting embodiment of the foregoing gas turbine engine assembly, the assembly may include a stator root radially aligned with a radially inner boundary of a core flow path through the gas turbine engine.

In a further non-limiting embodiment of either of the foregoing gas turbine engine assemblies, the assembly may include a blade root radially aligned with a radially inner boundary of a core flow path through the turbine engine.

In a further non-limiting embodiment of any of the foregoing gas turbine engine assemblies, the assembly may include a core inlet stator positioned within an inlet to a core section of the gas turbine engine.

In a further non-limiting embodiment of any of the foregoing gas turbine engine assemblies, the assembly may have an inlet flow of the compressor section that is from 30 lb/sec/ft² to 37 lb/sec/ft² when the gas turbine engine is operating at a cruise speed.

In a further non-limiting embodiment of any of the foregoing gas turbine engine assemblies, the assembly may have a turbine inlet temperature of a high-pressure turbine within the gas turbine engine that is from 2,000 F to 2,600 F when the gas turbine engine is operating at a cruise speed.

In a further non-limiting embodiment of any of the foregoing gas turbine engine assemblies, the assembly may include the compressor blade having tip speed during operation that is from 1,050 fps to 1,350 fps when the gas turbine engine is operating at cruise speed.

In a further non-limiting embodiment of any of the foregoing gas turbine engine assemblies, the assembly may include a fan section of the turbomachine driven by a geared architecture that is driven by a shaft. The geared architecture having a gear reduction ratio from 2.2 to 4.

A compressor module of a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a compressor module a first inner boundary facing radially outward and defining in-part a core inlet, the first inner boundary located at a first radial distance from the rotational axis. The compressor module includes a second inner boundary facing radially outward and located axially downstream of the first inner boundary. The second inner boundary is located at a second radial distance from the rotational axis. A ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9.

In a further non-limiting embodiment of the foregoing compressor module, the radially inner boundary of the compressor section inlet is at a location of a compressor rotor and the radially inner boundary of the core inlet is at a location of a core inlet.

DESCRIPTION OF THE FIGURES

The various features and advantages of the disclosed examples will become apparent to those skilled in the art from the detailed description. The figures that accompany the detailed description can be briefly described as follows:

FIG. 1 shows a partial section view of an example turbomachine.

FIG. 2 shows a close-up view of a core inlet portion of the FIG. 1 turbomachine.

DETAILED DESCRIPTION

Referring to FIG. 1, an example turbomachine, which is a two-spool gas turbine engine 20, generally includes a fan section 22, a compressor section or module 24, a combustion section 26, and a turbine section 28. Other examples may include an augmentor section (not shown) among other systems or features.

Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans. That is, the teachings may be applied to other types of turbomachines and turbine engines including three-spool architectures.

In the example engine 20, air moves from the fan section 22 to a bypass flow path B or a core flow path C. The core flow path C is within a core section of the engine 20. Within the core flow path, compressed air from the compressor section 24 communicates through the combustion section 26. The products of combustion expand through the turbine section 28.

The example engine 20 generally includes a low-speed spool 30 and a high-speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36. The low-speed spool 30 and the high-speed spool 32 are rotatably supported by several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively, or additionally, be provided.

The low-speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low-pressure compressor 44, and a low-pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a geared architecture 48 to drive the fan 42 at a lower speed than the low-speed spool 30.

The high-speed spool 32 includes an outer shaft 50 that interconnects a high-pressure compressor 52 and high-pressure turbine 54.

The combustion section 26 includes a circumferentially distributed array of combustors 56 generally arranged axially between the high-pressure compressor 52 and the high-pressure turbine 54.

The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A, which is collinear with the longitudinal axes of the inner shaft 40 and the outer shaft 50.

A mid-turbine frame 58 of the engine static structure 36 is generally arranged axially between the high-pressure turbine 54 and the low-pressure turbine 46. The mid-turbine frame 58 supports bearing systems 38 in the turbine section 28.

In the example engine 20, the core airflow C is compressed by the low-pressure compressor 44 then the high-pressure compressor 52, mixed and burned with fuel in the combustors 56, then expanded over the high-pressure turbine 54 and low-pressure turbine 46. The high-pressure turbine 54 and the low-pressure turbine 46 rotatably drive the respective high-speed spool 32 and low-speed spool 30 in response to the expansion. The high-pressure turbine 54 and the low-pressure turbine 46 drive a compressor rotor of the high-pressure compressor 52 and a compressor rotor of the low-pressure compressor 44, respectively.

In some non-limiting examples, the engine 20 is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6:1).

The geared architecture 48 of the example engine 20 includes an epicyclic gear train, such as a planetary, star, or differential gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.2 (i.e., 2.2:1). In some examples, the gear reduction ratio is from about 2.2 to 4. In such examples, these parameters result in superior engine fuel consumption characteristics. These ranges of ratios help define the inner diameter of the core flow path in these examples. The low-pressure turbine 46 pressure ratio is pressure measured prior to inlet of low-pressure turbine 46 as related to the pressure at the outlet of the low-pressure turbine 46 prior to an exhaust nozzle of the engine 20. In one non-limiting embodiment, the bypass ratio of the engine 20 is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low-pressure turbine 46 has a pressure ratio that is greater than about 5 (5:1). In prior art designs, significantly larger sized, lower-speed turbines are required to achieve this same range of ratios. In this embodiment, the low-pressure turbine is smaller and runs faster than the prior art designs. The geared architecture 48 of such embodiments is an epicyclic gear train with a gear reduction ratio of greater than about 2.4 (i.e., 2.4:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

In the example engine 20, a significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. Although described above as about 0.8 Mach, cruise may range from about 0.7 to 0.9 Mach.

Fan Pressure Ratio is the pressure ratio across a blade of the fan section 22 without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example engine 20 is less than 1.45. This ratio is a characteristic of many of the disclosed examples.

Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of “T”/518.7^0.5. T represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example engine 20 is less than about 1150 fps (351 m/s). Low fan tip speed are often desirable for good fan efficiency and low noise.

In some examples, the tip speed of the low-pressure compressor 44 is from about 1,050 fps (320 m/s) to 1350 fps (411 m/s). Speeds within this range may balance fuel consumption, with low stage number and parts count.

Referring now to FIG. 2 with continued reference to FIG. 1, the core flow path of the example engine 20 begins at a core inlet 60 and extends through and past the low-pressure compressor 44. The core inlet 60 has a radially inner boundary 62 and a radially outer boundary 66.

A core inlet stator 70 is located at or near the core inlet 60. The core inlet stator 70 attaches to a core case 74 at the radially inner boundary 62. The core inlet stator 70 attaches to an inlet case 78 at the radially outer boundary 66. The core inlet stator 70 extends radially across the core flow path C.

In this example, the radially inner boundary 62 is positioned a radial distance D₁ from the axis A. The distance Di, in this example, also corresponds to the radial distance between a root 64 of the core inlet stator 70 and the axis A. In this example, the root 64 of the core inlet stator 70 is radially aligned with the radially inner boundary 62 of the core flow path C.

After flow moves through the core inlet 60, the flow moves through a compressor inlet 82 into the compressor section 24. In this example, the compressor section inlet 82 is an inlet to the low-pressure compressor 44 of the compressor section 24. The compressor inlet 82 extends from a radially inner boundary 86 to a radially outer boundary 90.

Notably, a blade 98 of a rotor within the low-pressure compressor 44 extends from a root 102 to a tip 106. The blade 98 is located at or near the compressor inlet 82. The blade 98 part of a compressor rotor within a first stage of the compressor section 24. The blade 98 is thus part of a first stage rotor, or a leading blade of the compressor section 24 relative to a direction of flow along the core flow path C.

In some examples, the blade 98 represents the axial position where air enters the compressor section 24 of the core flow path C. The blade 98 extends radially across the core flow path C.

The radially inner boundary 86 is positioned a radial distance D₂ from the axis A. The distance D₂, in this example, also corresponds to the radial distance between the root 102 of the blade 98 and the axis A. In this example, the root 102 is radially aligned with the radially inner boundary 86 of the core flow path C.

In the example engine 20, a preferred ratio range of the distance D₂ to the distance D₁ spans from about 0.65 to about 0.9, which provides a relatively low profile core flow path contour. High profile flow path contours have greater differences between D₂ and D₁, and thus larger “humps” between the core inlet 60 and the compressor inlet 82. High profile flow path contours introduce discontinuities that undesirably disrupt the airflow and undesirably add weight to the engine 20. The ratio range of about 0.65 to 0.9 is made possible, in part, by the incorporation of the geared architecture 48 into the engine 20. The “hump” in this example is generally area 100.

Other characteristics of the engine having this ratio may include the engine 20 having a specific inlet flow of the low pressure compressor at cruising speeds to be between 30 lb/sec/ft² to 37 lb/sec/ft². The specific inlet flow is the amount of flow moving into the compressor section 24 and specifically, in this example, into a compressor inlet 82 and through the compressor section 24.

Another characteristic of the example engine 20 is that the cruise speeds of the example engine are generally Mach numbers of about 0.7 to 0.9.

Yet another characteristic of the engine 20 is that a temperature at an inlet to the high-pressure turbine 54 may be from 2,000° F. (1093.33° C.) to 2,600° F. (1426.66° C.). Maintaining temperatures within this range balance of good fuel consumption, low engine weight, and low engine maintenance costs.

Yet another characteristic of the engine 20 is that a tip speed of blades in a rotor of the low-pressure compressor 44 (a compressor rotor) may be from about 1,050 fps (320 m/s) to 1,350 fps (411 m/s).

In this example, the geared architecture 48 of the engine 20 may have a gear ratio of from about 2.2 to 4.

The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. Thus, the scope of legal protection given to this disclosure can only be determined by studying the following claims.

Claims

1. A turbomachine having a geared architecture, comprising:

a core inlet having a radially inner boundary that is spaced a first radial distance from a rotational axis of a turbomachine assembly; and

a compressor section inlet having a radially inner boundary that is spaced a second radial distance from the rotational axis, wherein a ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9.

2. The turbomachine having a geared architecture of claim 1, wherein the radially inner boundary of the core inlet is at a location of a core inlet stator.

3. The turbomachine having a geared architecture of claim 1, wherein the radially inner boundary of the compressor section inlet is at a location of a compressor rotor.

4. The turbomachine having a geared architecture of claim 3, wherein the compressor rotor is a first stage rotor of a low-pressure compressor.

5. The turbomachine having a geared architecture of claim 1, wherein the core inlet is an inlet to a core section of the turbomachine.

6. The turbomachine having a geared architecture of claim 1, wherein an inlet flow of the compressor section is from about 30 lb/sec/ft2 to 37 lb/sec/ft2 when the turbomachine is operating at a cruise speed.

7. The turbomachine having a geared architecture of claim 1, wherein a turbine inlet temperature of a high-pressure turbine within the turbomachine is from about 2,000 F to 2,600 F when the turbomachine is operating at a cruise speed.

8. The turbomachine having a geared architecture of claim 1, wherein a tip speed of a blade array in the compressor section during operation is from about 1,050 fps to 1,350 fps.

9. The turbomachine having a geared architecture of claim 1, wherein a fan section of the turbomachine is driven by a geared architecture that is driven by a shaft that rotates a compressor rotor within the compressor section.

10. The turbomachine having a geared architecture of claim 9, wherein the geared architecture has a gear reduction ratio from about 2.2 to 4.

11. A gas turbine engine, comprising:

a core inlet stator having a stator root that is spaced a first radial distance from a rotational axis of a gas turbine engine; and

a compressor blade within a first stage of a compressor section, the compressor blade having a blade root that is spaced a second radial distance from the rotational axis, wherein a ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9.

12. The gas turbine engine assembly of claim 11, wherein the stator root is radially aligned with a radially inner boundary of a core flow path through the gas turbine engine.

13. The gas turbine engine assembly of claim 11, wherein the blade root is radially aligned with a radially inner boundary of a core flow path through the turbine engine.

14. The gas turbine engine assembly of claim 11, wherein the core inlet stator is positioned within an inlet to a core section of the gas turbine engine.

15. The gas turbine engine assembly of claim 11, wherein an inlet flow of the compressor section is from 30 lb/sec/ft2 to 37 lb/sec/ft2 when the gas turbine engine is operating at a cruise speed.

16. The gas turbine engine assembly of claim 11, wherein a turbine inlet temperature of a turbine within the gas turbine engine is from 2,000 F to 2,600 F of when the gas turbine engine is operating at a cruise speed.

17. The gas turbine engine of claim 11, wherein a tip speed the compressor blade during operation is from 1,050 fps to 1,350 fps when the gas turbine engine is operating at cruise speed.

18. The gas turbine engine of claim 11, wherein a fan section of the turbomachine is driven by a geared architecture that is driven by a shaft, the geared architecture having a gear reduction ratio from 2.2 to 4.

19. A compressor module of a gas turbine engine having a rotational axis, the compressor module comprising:

a first inner boundary facing radially outward and defining in-part a core inlet, the first inner boundary located at a first radial distance from the rotational axis; and

a second inner boundary facing radially outward and located axially downstream of the first inner boundary, the second inner boundary located at a second radial distance from the rotational axis, wherein a ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9.

20. The compressor module of claim 19 wherein the second inner boundary defines in-part a compressor inlet of the compressor module.