WO2014120309A2

WO2014120309A2 - Fluid-cooled seal arrangement for a gas turbine engine

Info

Publication number: WO2014120309A2
Application number: PCT/US2013/068053
Authority: WO
Inventors: Stephen J. Lyle; Fungayi Mutsengi; Ernest Boratgis; Santosh Ranganath; Christopher J. Larson; M Rifat Ullah
Original assignee: United Technologies Corporation
Priority date: 2012-11-01
Filing date: 2013-11-01
Publication date: 2014-08-07
Also published as: WO2014120309A3; EP2914819A4; EP2914819A2; US20140119887A1

Abstract

A rotational component for a gas turbine engine includes at least one passage with a semi-spherical end.

Description

FLUID-COOLED SEAL ARRANGEMENT FOR A GAS TURBINE ENGINE

Applicant hereby claims priority to U.S. Patent Application No. 13/666,591 filed November 1, 2012, the disclosure of which is herein incorporated by reference.

BACKGROUND

[0001] The present disclosure relates to a gas turbine engine and, more particularly, to fluid-cooled seal arrangements therefor.

[0002] Certain sections of gas turbine engines may be subjected to high temperatures and pressures. Some engine components may be sensitive to the harsh environment thereof and are shielded therefrom. These components are typically vented by ambient or cooling bleed-off air or have cooling oil flowing therethrough. In order to maintain cool air in the cavities housing these components, the cavities must be shielded from engine pressure and temperature differentials. Typically, seal systems are positioned to prevent high temperature and pressure air from flowing downstream into the areas with lower temperature and pressure air. One such seal system includes arcuate carbon material segments arranged to form a stationary carbon ring that forms a rubbing interface with a rotating seal runner.

[0003] The rubbing interface between the rotating seal runner and the carbon ring minimizes or prevents leakage, however if the heat generated by the rubbing interface is not adequately dissipated, the rotating seal runner may thermally distort. This may degrade performance manifested by excessive fluid leakage.

[0004] One approach to minimize overheating of the seal interface includes the delivery of a cooling fluid onto the underside of the rotating seal runner sprayed from a stationary nozzle positioned proximate the rotating seal runner. The relative motion between the rotating seal runner and the stationary nozzle causes a uniform film of cooling fluid to be deposited on the seal runner to extract thermal energy. Stationary nozzles provide a consistently even film of cooling fluid on the underside of the rotating seal runner, however their applicability on many gas turbine engines is limited by physical constraints that prevent the nozzle from being located proximate the seal runner.

[0005] Another approach utilizes a rotating distributor to deliver the cooling fluid onto the underside of the rotating seal runner. The rotating distributor is typically affixed to the seal runner, and a steady stream of cooling fluid is delivered through a central passageway in the rotating distributor to the underside of the seal runner. A series of openings in the rotating distributor dispense the cooling fluid onto the seal runner. Inherently, because of the absence of relative motion between the distributor and the seal runner there may be an uneven distribution of cooling fluid on the underside of the seal runner.

[0006] Carbon seal systems with a rotating distributor include a significant quantity of openings to deliver an even film of cooling fluid. If this design parameter is not satisfied, an uneven film of cooling fluid is distributed across the seal runner, which may cause an uneven extraction of heat. Although effective, carbon seal system designs with large quantities of cooling fluid dispensing openings may, under centrifugal loads, concentrate stresses which may lead to fatigue life shortfalls. In some cases, expensive rotor grade material may be used to meet fatigue life requirements. In other cases even rotor grade material is insufficient to meet desired fatigue life requirements. SUMMARY

[0007] A rotational component for a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes at least one passage with a semi- spherical end.

[0008] In a further embodiment of the foregoing embodiment, the at least one passage is an oil communication passage.

[0009] In a further embodiment of any of the foregoing embodiments, the at least one passage is a blind hole.

[0010] A gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes a rotational component having at least one passage with a semi-spherical end.

[0011] In a further embodiment of foregoing embodiment, the rotational component is a seal runner. In the alternative or additionally thereto, in the foregoing embodiment the at least one passage is an oil communication passage. In the alternative or additionally thereto, in the foregoing embodiment the at least one passage is a blind hole. In the alternative or additionally thereto, in the foregoing embodiment the oil communication passage is an outlet passage.

[0012] In a further embodiment of any of the foregoing embodiments, the rotational component includes a cantilevered section. In the alternative or additionally thereto, in the foregoing embodiment the cantilevered section collects and directs a lubricant toward an internal inlet passage. In the alternative or additionally thereto, in the foregoing embodiment the cantilevered section include a hook-shaped end section. [0013] A method of forming a passage in a rotation component according to another disclosed non-limiting embodiment of the present disclosure includes forming a semi-spherical end at an end of the passage.

[0014] In a further embodiment of the foregoing embodiment, the method comprises drilling the passage and inserting a ball-endmill into the passage to form the end.

[0015] In a further embodiment of any of the foregoing embodiments, the method includes drilling the passage with a drill tool having spherical shaped flutes.

BRIEF DESCRIPTION OF THE DRAWINGS [0016] Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows:

[0017] Figure 1 is a schematic cross-section of a gas turbine engine;

[0018] Figure 2 is a partial longitudinal schematic sectional view of a bearing compartment that may be used with the gas turbine engine shown in Figure 1 ;

[0019] Figure 3 is a sectional view of a seal runner with a passage with a semi- spherical end according to one disclosed non-limiting embodiment; and

[0020] Figure 4 is an expanded view of a RELATED ART passage. DETAILED DESCRIPTION

[0021] Figure 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flowpath while the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a 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 as the teachings may be applied to other types of turbine engines such as a three-spool (plus fan) engine wherein an intermediate spool includes an intermediate pressure compressor (IPC) between the LPC and HPC and an intermediate pressure turbine (IPT) between the HPT and LPT.

[0022] The engine 20 generally includes a low spool 30 and a high spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing structures 38. The low spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 ("LPC") and a low pressure turbine 46 ("LPT"). The inner shaft 40 drives the fan 42 directly or through a geared architecture 48 to drive the fan 42 at a lower speed than the low spool 30. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.

[0023] The high spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 ("HPC") and high pressure turbine 54 ("HPT"). A combustor 56 is arranged 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 about the engine central longitudinal axis A which is collinear with their longitudinal axes.

[0024] Core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed with the fuel and burned in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The turbines 54, 46 rotationally drive the respective low spool 30 and high spool 32 in response to the expansion.

[0025] The main engine shafts 40, 50 are supported at a plurality of points by bearing structures 38 within the static structure 36. It should be understood that various bearing structures 38 at various locations may alternatively or additionally be provided.

[0026] In one non-limiting example, the gas turbine engine 20 is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 bypass ratio is greater than about six (6:1). The geared architecture 48 can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1. The geared turbofan enables operation of the low spool 30 at higher speeds which can increase the operational efficiency of the low pressure compressor 44 and low pressure turbine 46 and render increased pressure in a fewer number of stages.

[0027] A pressure ratio associated with the low pressure turbine 46 is pressure measured prior to the inlet of the 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 gas turbine engine 20. In one non- limiting embodiment, the bypass ratio of the gas turbine 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). 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. [0028] In one embodiment, a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio. The fan section 22 of the gas turbine 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 gas turbine 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.

[0029] 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 gas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of "T" / 518.7^0'5 in which "T" represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine 20 is less than about 1150 fps (351 m/s).

[0030] With reference to Figure 2, a bearing structure 38A according to one disclosed non-limiting embodiment includes a bearing 58 and fluid-cooled seal arrangements 60. The fluid-cooled seal arrangements 60A, 60B may each be, in the disclosed non-limiting embodiment a carbon seal system to seal a "wet" zone from a "dry" zone. For ease of reference hereinafter regions or volumes containing lubricant will be referred to as a "wet" zone and a lubricant free region will be referred to as a "dry" zone. So, for example, the interior of a bearing compartment may be referred to as a wet zone which ultimately communicates with a sump (not shown) of a lubricant system (illustrated schematically) while the region external thereto will be referred to as a dry zone. The fluid-cooled seal arrangement 60 generally includes a stationary component 62 and a rotational component 64. [0031] The stationary component 62 is coupled to a rotationally fixed structure such as the static structure 36 while the rotational component 64 is mechanically connected to a rotating structure such as the outer shaft 50. It should be appreciated, however, that any rotating structure such as a rotor hub may alternatively mount the rotational component 64. The rotational component 64 is referred to herein as a seal runner.

[0032] The stationary component 62 is arranged with respect to the rotational component 64 to form a rubbing interface 66 therebetween which in the disclosed non-limiting embodiment, is axially oriented with respect to the engine axis A. Alternatively, the rubbing interface 66 may be radially oriented. A first annular surface 68 is defined by the stationary component 62 and a second annular surface 70 is defined by the rotational component 64 which are maintained in rubbing contact to form a fluid tight seal at the rubbing interface 66.

[0033] The lubrication system (illustrated schematically) provides cooling fluid under pressure to lubricate and cool the moving parts of the engine 20, such as the bearing 58 and fluid-cooled seal arrangement 60 through a nozzle 72. The lubrication system discharges the fluid from the nozzle 72 with sufficient kinetic energy to spray an underside of a cantilevered section 74 of the rotational component 64. The cantilevered section 74 may include a hook- shaped end section 76 to collect and direct the lubricant toward an internal passage 78 (Figure 3).

[0034] With respect to Figure 3, the internal inlet passage 78 may be formed from one or a multiple of drill holes. The internal inlet passage 78 communicates with an internal outlet passage 82. The internal outlet passage 82 may also be formed from one or a multiple of drill holes.

[0035] The internal outlet passage 82 includes a semi-spherical end 90 as compared to a conventional conical end E (Figure 7; RELATED ART). It should be understood that although only the internal outlet passage 82 includes the semi-spherical end 90, any passage may benefit herefrom such as the internal inlet passage 78. The semi-spherical end 90 is applicable to any blind hole such as a drill hole and may be manufactured with, for example only, a ball- endmill type tool which is chased down a conventional drill hole or other special drill tool with spherical shaped flutes. The semi-spherical end 90 reduces concentrated stress in the passages and increases fatigue life of the rotational component 64. That is, the semi-spherical end 90 may be particularly applicable to any rotating structure such as the disclosed seal runner which is subject to high centrifugal forces and is potentially a less expensive alternative to usage of higher grade materials to meet fatigue life requirements.

[0036] It should be understood that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.

[0037] It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom.

[0038] Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.

[0039] The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.

Claims

CLAIMS What is claimed is:

1. A rotational component for a gas turbine engine comprising:

at least one passage with a semi-spherical end.

2. The rotational component as recited in claim 1, wherein said at least one passage is an oil communication passage.

3. The rotational component as recited in claim 1, wherein said at least one passage is a blind hole.

4. A gas turbine engine comprising:

a rotational component having at least one passage with a semi-spherical end.

5. The gas turbine engine as recited in claim 4, wherein said rotational component is a seal runner.

6. The gas turbine engine as recited in claim 5, wherein said at least one passage is an oil communication passage.

7. The rotational component as recited in claim 6, wherein said at least one passage is a blind hole.

8. The gas turbine engine as recited in claim 6, wherein said oil communication passage is an outlet passage.

9. The gas turbine engine as recited in claim 4, wherein said rotational component includes a cantilevered section.

10. The gas turbine engine as recited in claim 9, wherein said cantilevered section collects and directs a lubricant toward an internal inlet passage

11. The gas turbine engine as recited in claim 10, wherein said cantilevered section include a hook-shaped end section.

12. A method of forming a passage in a rotation component comprising:

forming a semi-spherical end at an end of the passage.

13. The method as recited in claim 12, further comprising:

drilling the passage;

inserting a ball-endmill into the passage to form the end.

14. The method as recited in claim 12, further comprising:

drilling the passage with a drill tool having spherical shaped flutes.