Gas turbine having at least two shafts designed as hollow shafts at least in some areas and arranged coaxially relative to one another

Abstract

The present invention describes a gas turbine having two shafts, which are rotatably mounted using bearing devices in the area of a casing. An intermediate shaft sealing device is arranged between the bearing devices, which includes a sealing element connected to a first shaft, in which a further sealing element connected to the second shaft radially engages. An operating pressure is applied, in the area limited by the shafts in the axial direction of the shafts, to facing effective areas of the sealing elements, while the pressure prevailing in the area outside the shafts in the axial direction acts on effective areas of the sealing elements facing away from one another. A ratio between the outer diameters limiting the facing effective areas and the inner diameters likewise limiting these effective areas is in each case greater than or equal to 1.25.

Claims

1. A gas turbine comprising: a casing; bearing devices supported by the casing; two shafts, each of the two shafts being hollow in some areas, the two shafts being arranged coaxially relative to one another, the two shafts being rotatably mounted in the casing by the bearing devices, a hydraulic intermediate shaft sealing device arranged between the bearing devices and which separates a first area inside at least one of the two shafts and limited by the two shafts from a second area outside the two shafts, where, in the first area, a first operating pressure prevails that differs from a second operating pressure acting in the second area, the hydraulic intermediate shaft sealing device including a first sealing element connected to a first shaft of the two shafts, the first sealing element extending over a circumference of the first shaft, and including an inner groove area with a defined groove depth, a second sealing element connected to a second shaft of the two shafts and radially engaging the inner groove area, where the first operating pressure is applied in the first area in an axial direction of the two shafts to facing effective areas of the first and second sealing elements, with each facing effective area being radially limited by an outer diameter and an inner diameter, each outer diameter and each inner diameter positioned on a respective one of the first and second sealing elements, while the second operating pressure prevailing in the second area acts in the axial direction on effective areas of the first and second sealing elements facing away from one another, where a ratio between the outer diameters of the facing effective areas of the first and second sealing elements and the inner diameters of the facing effective areas of the first and second sealing elements is in each case greater than or equal to 1.25.

2. The gas turbine in accordance with claim 1, wherein the inner groove area includes a groove bottom, in a radial direction, at a distance from a radially outer area of the second sealing element.

3. The gas turbine in accordance with claim 2, wherein during operation of the gas turbine, the radially outer area of the second sealing element dips into a defined oil volume in the inner groove area of the first sealing element.

4. The gas turbine in accordance with claim 3, wherein the second sealing element is shaped in some areas as an annular disk.

5. The gas turbine in accordance with claim 4, and further comprising a bearing chamber in which the bearing devices are positioned.

6. The gas turbine in accordance with claim 5, and further comprising a piston ring positioned between the two shafts, which separates, while the two shafts are stationary, the second area and the first area on a side of the hydraulic intermediate shaft sealing device facing the first area.

7. The gas turbine in accordance with claim 6, wherein the two shafts rotate in a same rotational direction.

8. The gas turbine in accordance with claim 2, wherein the second sealing element is shaped in some areas as an annular disk.

9. The gas turbine in accordance with claim 8, and further comprising a bearing chamber in which the bearing devices are positioned.

10. The gas turbine in accordance with claim 9, and further comprising a piston ring positioned between the two shafts, which separates, while the two shafts are stationary, the second area and the first area on a side of the hydraulic intermediate shaft sealing device facing the first area.

11. The gas turbine in accordance with claim 10, wherein the two shafts rotate in a same rotational direction.

12. The gas turbine in accordance with claim 1, wherein the second sealing element is shaped in some areas as an annular disk.

13. The gas turbine in accordance with claim 12, and further comprising a bearing chamber in which the bearing devices are positioned.

14. The gas turbine in accordance with claim 13, and further comprising a piston ring positioned between the two shafts, which separates, while the two shafts are stationary, the second area and the first area on a side of the hydraulic intermediate shaft sealing device facing the first area.

15. The gas turbine in accordance with claim 14, wherein the two shafts rotate in a same rotational direction.

16. The gas turbine in accordance with claim 1, and further comprising a bearing chamber in which the bearing devices are positioned.

17. The gas turbine in accordance with claim 1, and further comprising a piston ring positioned between the two shafts, which separates, while the two shafts are stationary, the second area and the first area on a side of the hydraulic intermediate shaft sealing device facing the first area.

18. The gas turbine in accordance with claim 1, wherein the two shafts rotate in a same rotational direction.

Description

(1) The sole FIGURE of the drawing shows a schematized longitudinal sectional view in partial representation of an exemplary embodiment of the gas turbine in accordance with the present invention.

(2) The FIGURE shows a longitudinal sectional view in partial representation of a gas turbine 1 designed as aircraft engine having two shafts 2, 3 designed as hollow shafts at least in some areas and arranged coaxially relative to one another, which shafts are rotatably mounted using bearing devices 4, 5 in the area of a casing 6. A hydraulic intermediate shaft sealing device 7 is arranged between the bearing devices 4, 5 of the shafts 2, 3, which separates a first area 8 provided inside at least one of the shafts 2 and limited by the shafts 2, 3 from an area 9 provided outside said shafts, which area in the present case represents a bearing chamber of the bearing devices 4 and 5.

(3) The operating pressure acting in the area of the bearing chamber 9 during operation is lower than the operating pressure acting in the area 8 limited by the shafts 2, 3, thus ensuring that even in the event of a failure of the sealing effect in the area of the intermediate shaft sealing device 7 any entry of oil emanating from the bearing chamber 9 in the direction of the area 8 is prevented. The area 8 is connected to the high pressure area of the jet engine 1, in the area of which such high operating temperatures prevail during operation of the jet engine 1 that oil present in the area 8 is ignited, which is however undesirable.

(4) The intermediate shaft sealing device 7 includes a sealing element 10 connected to the shaft 2 or to the high-pressure shaft, respectively, and extending over the circumference of the first shaft 2 as well as an inner groove area 11 with a defined groove depth t. A further sealing element 12 of the intermediate shaft sealing device 7 connected to the second shaft 3 radially engages the inner groove area 11. The operating pressure is applied, in the area 8 limited by the shafts 2, 3 in the axial direction of the shafts 2, 3, to facing effective areas 13, 14 of the sealing elements 10, 12, while the pressure prevailing in the area 9 outside the shafts 2, 3 in the axial direction acts on effective areas of the sealing elements 10, 12 facing away from one another.

(5) A groove bottom 17 of the inner groove area 11 is arranged in the radial direction, at a distance from the radially outer area of the further sealing element.

(6) To permit, during operation of the aircraft engine 1, a sealing separation to the required extent of the bearing chamber 9 from the area 8 by means of the intermediate shaft sealing device 7, an oil volume flow is introduced into the inner groove area 11 during operation of the jet engine 1. Depending on the pressure drop between the bearing chamber 9 and the area 8 and on the speeds of the unidirectionally rotating shafts 2 and 3, a defined oil volume 21 is present in the inner groove area 11 without any further measures and is maintained there by the centrifugal force acting on the oil due to the rotation, where due to a continuous supply of oil into the inner groove area 11 the oil volume present there is continually replaced, hence preventing any unwelcome heating up of the oil in the inner groove area 11.

(7) Depending on the pressure drop between the bearing chamber 9 and the area 8, a different oil level is obtained in the inner groove area 11 on that side of the sealing element 12 facing the bearing device 5 than on that side of the sealing element 12 facing the bearing device 4. This results from the fact that the pressure in the area 8 is higher than in the area of the bearing chamber 9.

(8) To ensure an appropriately high sealing effect, the further sealing element 12, in its outer rim area 19 arranged in the inner groove area 11, is designed in annular disk form and substantially parallel to the areas of the first sealing element 10 limiting the inner groove area 11. In addition, during operation of the gas turbine 1, the further sealing element 12 dips into the defined oil volume 21 with a radially outer area 18 of the rim area 19 inside the inner groove area 11 of the sealing element 10. As a result, any leakage between the area 8 limited by the shafts 2, 3 and the area 9 outside the shafts 2, 3 is substantially equal to 0, without causing high performance losses or causing undesirably high wear in the area of the intermediate shaft sealing device 7.

(9) In addition, a piston ring 20 is provided between the shafts 2 and 3, which separates, while the shafts 2, 3 are stationary, the area or the bearing chamber 9 outside the shafts 2, 3 and the area 8 limited by the shafts 2, 3 on that side of the intermediate shaft sealing device facing the area 8 limited by the shafts 2, 3, in order to prevent any ingress of oil emanating from the bearing chamber 9 into the area 8 while the shafts 2, 3 are stationary and when the aircraft engine 1 is switched off. During operation, i.e. while the shafts 2, 3 are rotating, the piston ring 20 lifts off from the shaft 3, so that the piston ring 20 is then no longer contacting the outside of the shaft 3, and undesirable performance losses in this area are prevented during operation of the aircraft engine 1.

(10) The structural design of the aircraft engine 1 according to the FIGURE is such that in the zone of the bearing device 4, in the area of which the high-pressure shaft 2 is rotatably mounted, an axial force component F4A prevails at the current operating point, which must be braced by the bearing device 4 in the casing 6 and which acts in the area of the bearing device 4 in the direction of the intermediate shaft sealing device 7. At the same time, an axial force component F5A likewise acting in the direction of the intermediate shaft sealing device 7 acts in the zone of the bearing device 5, in the area of which the low-pressure shaft 3 is rotatably mounted. The two axial force components F4A and F5A are those axial forces made up of all axial forces acting on the shafts 2 and 3.

(11) In order to prevent, in the area of the bearing devices 4 and 5, resulting axial force components F4A, F5A that are inadmissibly high and that reduce a required long service life of the bearing devices 4, 5, a ratio between the outer diameters DA13/DA14 limiting the facing effective areas and the inner diameters DI13/DI14 likewise limiting these effective areas 13, 14 is in each case greater than or equal to 1.25. In this case, the inner diameters DI13 and DI14 of the effective areas 13 and 14 of the sealing elements 10, 12 of the intermediate shaft sealing device 7 in the exemplary embodiment of the jet engine 1, considered in more detail here, substantially correspond to inner diameters in the area of intermediate shaft sealing devices in conventional aircraft engines, while the outer diameters DA13 and DA14 of the effective areas 13 and 14 are enlarged to the extent in accordance with the invention when compared with conventionally designed aircraft engines. Hence the solution in accordance with the invention can be implemented to a minor extent in existing engine concepts, in order to permit application of the respectively resulting axial forces FA13 and FA14, in the zone of the bearing device 4 and also of the bearing device 5, from the pressure applying in the zone of the effective areas 13 and 14 of the area 8 limited by the shafts 2, 3, which forces counteract the resulting axial force components F4A and F5A respectively and reduce the axial bearing load in the area of the bearing devices 4 and 5 to the required extent.

(12) Depending on the application in question, however, it is also possible to design the inner diameters DI13 and DI14 in a different suitable manner in order to provide the previously described axial bearing force compensation with the diameter ratio greater than or equal to 1.25.

(13) The pressure difference between the operating pressure in the bearing chamber 9 and the operating pressure in the area 8 varies over the operating range of the jet engine 1 and, among others, also depending on the flight altitude, flight speed and the like, which also affects the height of the axial force components FA13 and FA14. Generally speaking, however, it is possible to set the axial force compensation in the area of the bearing devices 4 and 5 by appropriate design of the outer diameters DA13 and DA14 and of the inner diameters DI13 and DI14, and of the operating pressures of the bearing chamber 9 and of the area 8 applying respectively in the zone of the effective areas 13/14 and 15/16 of the intermediate shaft sealing device 7, and the sealing elements 10 and 12.

(14) With the specified procedure for axial bearing force compensation in the area of the bearing devices 4 and 5, the axial bearing forces in the area of the bearing devices 4 and 5 can be limited with minor effects on the weight of the jet engine, without increasing a fuel consumption of the jet engine 1. Furthermore, undesirable temperature increases too of oil present in the area of the bearing chamber 9 are avoided, since the hydraulic intermediate shaft sealing device 7 exhibits substantially no leakage. Generally speaking, axial bearing compensation in the area of the intermediate shaft sealing device 7 can be achieved, by means of the diameter ratio in accordance with the invention between the outer diameter DA13 and the inner diameter DI13 and between the outer diameter DA14 and the inner diameter DI14 with the pressure drop currently present in conventionally designed jet engines between the bearing chamber 9 and the area 8 limited by the shafts 2, 3.

(15) Setting and/or dimensioning in accordance with the invention of the outer diameter DA13/DA14 in the area of the intermediate shaft sealing device 7 effects an improvement of the sealing effect in the area of the intermediate shaft sealing device 7, since the centrifugal force acting on the oil introduced or sprayed into the inner groove area 11 when the shafts 2, 3 are rotating is thereby increased and the oil is kept in that position inside the inner groove area 11 which provides the sealing effect of the intermediate shaft sealing device 7.

LIST OF REFERENCE NUMERALS

(16) 1 Gas turbine 2 Shaft, high-pressure shaft 3 Shaft, low-pressure shaft 4 Bearing device of high-pressure shaft 5 Bearing device of low-pressure shaft 6 Casing 7 Intermediate shaft sealing device 8 Area limited by the shafts 9 Bearing chamber area 10 Sealing element 11 Inner groove area 12 Further sealing element 13 Effective area of sealing element 14 Effective area of further sealing element 15 Effective area of sealing element 16 Effective area of further sealing element 17 Groove bottom 18 Radially outer area 19 Rim area of further sealing element 20 Piston ring 21 Oil volume DA 13, DA 14 Outer diameter of effective areas DI 13, DI 14 Inner diameter of effective areas F4A, F5A Axial force component FA 13, FA 14 Axial force t Groove depth