MAIN SHAFT OF A GAS TURBINE ENGINE AND GAS TURBINE ENGINE

Abstract

 The invention concerns a gas turbine assembly (20) with a central turbine shaft (28) that is connected to at least a gas turbine stage (26) of the gas turbine assembly (20) wherein the turbine shaft (28) extends in a longitudinal direction within a shaft tunnel (28.1) and is supported by bearings (28.2) arranged at or near the longitudinal ends of the shaft tunnel (28.1). The turbine shaft (28) has a longitudinal shaft section (28.3) that extends within the shaft tunnel (28.1). The longitudinal shaft section (28.3) that extends within the shaft tunnel (28.1) is provided with a plurality of grooves (28.4) extending in the longitudinal direction of the shaft (28) and being arranged around the circumference of the shaft (28).

Description

[0001] The invention relates to a gas turbine engine with a gas turbine. The invention relates in particular to a main shaft of a gas turbine engine and gas turbine engine.

[0002] Gas turbine engines in particular turboprop engines are generally known. A turboprop engine comprises a gas turbine that drives a propeller via a reduction gear. The gas turbine engine comprises a compressor, a combustion chamber and a turbine. The compressor is fed with air from the environment, compresses the air and feeds the compressed air to the combustion chamber. In the combustion chamber, the compressed air is mixed with fuel and the mixture is ignited and combusts. The hot combustion gases then drive a gas turbine that in turn drives the compressor and the propeller.

[0003] A turboprop engine is typically used as an aero engine that creates thrust by means of the propeller. The thrust of the hot combustion gases at the exhaust of the turbine does not significantly contribute to the engine's overall thrust.

[0004] In general, gas turbine engines can comprise an axial compressor or a centrifugal (i.e. radial) compressor. The gas turbine of the gas turbine engine in most cases is an axial turbine with one or more turbine stages. Turboprop gas turbine engines can have separate turbines for driving the propeller and for driving the compressor. Alternatively, the compressor can be fixed to the same shaft as the turbine for driving the propeller. Then, only one turbine is needed.

[0005] In a gas turbine engine, the compressor feeds high-pressure air to the combustion chamber. The combustion chamber is also known as combustor. In the combustor the compressed air is mixed with fuel and ignited. Thus, the compressed air is heated at constant pressure as the fuel/air mix burns. As it burns the fuel/air mix heats and rapidly expands. The burned mix, i.e. the combustion gases are exhausted from the combustor through the nozzle guide vanes to a turbine wheel.

[0006] The gas turbine can be an axial gas turbine comprising one or more turbine stages with nozzle guide vanes and turbine wheels (also known as turbine vanes). A nozzle guide vane is a stationary component located at the entry of a turbine stage in a gas turbine. The nozzle guide vane's function is to direct and guide the flow of high-velocity hot gases from the combustion chamber onto the turbine blades. The nozzle guide vane also helps to optimize the velocity and pressure of the gases entering the turbine stage, which can improve the efficiency and performance of the engine.

[0007] For larger airplanes, gas turbine engines and turboprop engines are the most secure and efficient aero engines. Gas turbine engines and turboprop engines having a size, weight and power suiting smaller airplanes, however, suffer from a very high consumption and thus are economically not feasible.

[0008] While turboprop engines typically are used for larger airplanes, small airplanes typically are powered by piston engines, for instance air cooled boxer engines providing a power between 100 and 400 kW. While turboprop engines have proven to be very reliable, replacing less reliable piston engines with turboprop engines typically is not feasible, because the efficiency of turboprop engines suffers from downscaling. This results in higher fuel consumption.

[0009] Gas turbine engine typically rotate at high rotation speeds. Therefore, reliable and sufficient lubrication of the main shaft of a gas turbine engine is important.

[0010] It is an object of the invention to provide an improved gas turbine engine of small airplanes.

[0011] According to a first aspect, a gas turbine assembly with a central turbine shaft that is connected to at least a gas turbine stage of the gas turbine assembly is provided, wherein the turbine shaft extends in a longitudinal direction within a shaft tunnel and is supported by bearings arranged at or near the longitudinal ends of the shaft tunnel. The turbine shaft has a longitudinal shaft section that extends within the shaft tunnel. The longitudinal shaft section that extends within the shaft tunnel is provided with a plurality of grooves extending in the longitudinal direction of the shaft and being arranged around the circumference of the shaft.

[0012] The inventor found that the slots or grooves cause - in combination with the small radial distance between the turbine shaft and the shaft tunnel - a suction effect, even almost a vacuum effect that sucks the lubricating and cooling oil into the shaft tunnel and allows controlling the flow of the lubricant, for instance the flow of lubricating oil. This lubrication system is self-supported. Thus, the need for an oil pump is avoided.

[0013] To maximize this effect, it is preferred if at least one longitudinal sub-section of the turbine shaft section has an outer diameter that is less than 1 mm smaller than an inner diameter of the shaft tunnel, resulting in a radial distance - i.e. a clearance - between the turbine shaft and the shaft tunnel that at average is smaller than 0,5 mm.

[0014] Preferably, the bearings are ball bearings.

[0015] According to a preferred embodiment, 6 to 24 longitudinally extending grooves are equally distributed around the circumference of the shaft. The grooves preferably have a width between 2 mm and 4 mm and a depth between 1 mm and 3 mm.

[0016] According to a further preferred embodiment, the longitudinal section of the turbine shaft that is provided with grooves tapers towards the longitudinal ends of the longitudinal section. This results in a turbine shaft having a longitudinal section that is characterized by longitudinally extending rips that taper towards the longitudinal ends and which are defined by the grooves separating the rips from each other.

[0017] According to a further preferred embodiment, the longitudinal section has a middle sub-section and tapering longitudinal end sections that taper towards the longitudinal ends of the longitudinal section. The depth of grooves preferably corresponds to half of the increase in diameter of the tapering longitudinal end sections. Due to the turbine shaft's middle sub-section and tapering longitudinal end sections that taper towards the longitudinal ends of the longitudinal section, the grooves are defining rips that extend in the longitudinal direction of the turbine shaft's longitudinal section and also extend in the radial direction.

[0018] The depth of grooves and thus the radial extension of the rips is between 1 mm and 3 mm, in particular 2 mm.

[0019] The number of grooves and thus the number of rips preferably is between 12 and 24, in particular about 18.

[0020] Preferably rips are defined by the grooves and the width of the grooves preferably corresponds to the width of the rips. For instance, the width of the grooves and/or the rips may be between 2 mm and 4 mm.

[0021] The gas turbine assembly with the turbine shaft as disclosed herein is particularly suitable for a small and light turboprop engine.

[0022] Accordingly, it is a second aspect of the invention to provide a turboprop engine with a gas turbine engine as defined by claim 1.

[0023] According to a third aspect, a gas turbine engine, in particular a turboprop engine, with an axial gas turbine is provided, wherein the gas turbine comprises a nozzle guide vane and a turbine wheel. The nozzle guide vane is arranged between an exhaust of a combustion chamber and the turbine wheel. The turbine wheel is connected to a turbine shaft. The nozzle guide vane is arranged to guide hot combustion gases from the combustion chamberto turbine blades of the turbine wheel. According to the invention, the nozzle guide vane comprises nozzle guide blades radially extending between an inner ring of the nozzle guide vane and an outer ring of the nozzle guide vane. The outer ring has an inner diameter that narrows in the direction of gas flow and thus provides a Venturi effect to accelerate the gas flow between the inner ring and the outer ring of the nozzle guide vane.

[0024] Thus, the nozzle guide vane provides a specific geometry of the walls provided by the inner ring and the outer ring and the guide blades extending there between in a radial direction to thus maximize the efficiency of the gas flow towards the turbine blades of the turbine wheel.

[0025] It is noted that the nozzle guide vane is a stator in the hot section of the gas turbine engine and is fundamental for directing the flow of hot combustion gases from the combustion chamber to the turbine wheel.

[0026] In general, the turboprop engine comprises a gas turbine assembly with a compressor, a combustor and an axial turbine wheel. The gas turbine comprises the at least one nozzle guide vane and the at least one turbine wheel. Preferably, the gas turbine comprises a single axial turbine stage with one nozzle guide vane and one turbine wheel.

[0027] The outer ring of the nozzle guide vane has an inner wall portion that defines a feed gas passage for feeding the hot combustion gases to the guide blades of the nozzle guide vane. A diameter of the feed gas passage as defined by the inner wall of the outer ring initially decreases in the direction of flow of the hot combustion gases causing the feeding gas passage to narrow in the direction of gas flow resulting an accelerating the gas flow through the feed gas passage due to a Venturi effect.

[0028] According to a preferred embodiment, the diameter of the inner wall portion of the outer ring increases where the guide blades contact the inner wall portion of the outer ring. Accordingly the inner wall of the outer ring has a smallest inner diameter where hot combustion gases are hitting the guide blades during operation of the gas turbine.

[0029] According to a preferred embodiment, an outer diameter of an outer wall portion of the inner ring decreases where the guide blades contact the outer wall portion of the inner ring. Thus, the distance between the outer wall of the inner ring and the inner wall of the outer ring increases in the direction of gas flow where the guide blades are arranged. This improves the efficiency of the gas turbine because it reduces losses caused by the nozzle guide vane.

[0030] Preferably, the outer ring, the guide blades and the inner ring are an integral part made of metal.

[0031] Preferably, the nozzle guide vane comprises between 16 to 24 guide blades, in particular 20 guide blades. The number of nozzle guide blades preferably is different from the number of turbine blades of the turbine wheel.

[0032] According to a preferred embodiment, the inner diameter of the inner wall portion of the outer ring at the entrance of the feed gas passage corresponds to an outer diameter of the annular exhaust nozzle of the combustion chamber of the combustor at the exit of the annular exhaust nozzle.

[0033] According to a further preferred embodiment, a longitudinal extension of the feed gas passage along a longitudinal axis of the gas turbine assembly is about 1.2 to 2.3 times of the extension of the guide blades of the nozzle guide vane along a longitudinal axis of the gas turbine assembly.

[0034] Preferably, the inner diameter of the inner wall portion of the outer ring at the entrance of the feed gas passage is between 140 mm and 170 mm, for instance 155 mm

[0035] According to a fourth aspect that can be combined with the first, the second and/or the third aspect, a combustor for a gas turbine assembly is provided. The combustor comprises a plurality of fuel injection nozzles and an annular combustion chamber. The annular combustion chamber comprises an inner space that is enclosed by a combustion chamber wall with an innerwall portion, a front wall portion and an outer wall portion. The front wall portion closes the combustion chamber at a combustion chamber front end and the inner wall portion and the outer wall portion define an open annular nozzle at the rear side of the combustion chamber. The fuel injection nozzles are circumferentially arranged around the outer wall portion and protrude into the inner space enclosed by the combustion chamber wall.

[0036] According to the fourth aspect, the inner wall portion and the outer wall portion are shaped so as to provide that the inner cross-sectional diameter of the annular inner space of the combustion chamber initially decreases (narrows) towards the open end of the combustion chamber nozzle and ultimately widens again, thus causing a Venturi effect where a radial distance between the inner wall portion and the outer wall portion is smallest.

[0037] The Venturi effect created by the annular Venturi nozzle portion of the combustion chamber improves the mass flow exiting from the combustion chamber and is fed to the axial turbine of the gas turbine assembly.

[0038] According to a preferred embodiment of the combustion chamber, the distance of the inner wall portion from a longitudinal axis of the gas turbine assembly initially increases in the direction of a combustion gas flow and then decreases again thus defining an apex that together with the outerwall portion of the combustion chamberwall defines a Venturi nozzle for accelerating the hot combustion exhaust gases and causing a lowered static pressure in the inner space close to the combustion chamber annular nozzle.

[0039] Preferably, the combustion chamber is surrounded by an open space that during operation is filled with compressed air and wherein the outer wall portion narrows towards to an open end of the annular nozzle portion of the combustion chamber. Preferably, holes (orifices) are provided in the narrowing outer wall portion. The holes are placed where during operation a reduced static pressure exists in the inner space, thus allowing air entering from the surrounding open space into the inner space and increasing the mass flow of the hot combustion gases exiting the combustion chamber. This improves the efficiency of the gas turbine assembly.

[0040] According to a preferred embodiment of the combustion chamber, the holes provided in the narrowing outer wall portion have a keyhole shape.

[0041] According to a further preferred embodiment of the combustion chamber, the outer wall portion comprises a generally cylindrical sub-portion in which an annular vortex generating protrusion is arranged, that protrudes inwardly into the inner space enclosed by the combustion chamber wall. The annular vortex generating protrusion improves mixing of fuel and compressed air and provides for an equal and complete combustion, thus improving the efficiency of the gas turbine assembly.

[0042] Preferably, the annular vortex generating protrusion comprises an upstream wall portion facing towards the fuel injection nozzles and a downstream wall portion facing away from the fuel injection nozzles. The vortex generating holes preferably are arranged in the downstream wall portion. This further improves the efficiency of the gas turbine assembly.

[0043] According to a further preferred embodiment of the combustion chamber, the inner wall portion comprises a frusto-conical shaped wall sub-portion with a diameter that increases towards the annular Venturi nozzles portion of the combustion chamber and provides that the annular free space between the inner wall portion and the outer wall portion becomes narrower, thus causing an acceleration of the hot combustion gases during operation and a decreasing static pressure in the annular free space between the inner wall portion and the outer wall portion.

[0044] Preferably, the frusto-conical shaped wall sub-portion of the inner wall portion is provided with holes allowing compressed air from the surrounding open space entering into the inner space, thus increasing the mass flow of the hot combustion gases exiting the combustion chamber. The increased mass flow further improves the efficiency of the gas turbine assembly.

[0045] According to a preferred embodiment of the combustor, the combustor comprises four fuel injection nozzles that are equally spaced from each other. This arrangement of fuel injection nozzles supports producing a homogeneous stream of hot combustion gases for driving the axial gas turbine of the gas turbine assembly.

[0046] According to a fifth aspect that can be combined with the first and/or the second and/or the third and/or the fourth aspect, a fuel injection nozzle for a combustion chamber of a gas turbine assembly is provided. The fuel injection nozzle comprises a central fuel duct having a distal end provided with a fuel nozzle. The central fuel duct is configured for feeding fuel from a proximal end of the fuel duct to the fuel nozzle at the distal end of the fuel duct. The fuel injection nozzle further comprises a fuel intake connector and a fuel return connector that both are connected to the proximal end of the fuel duct for feeding pressurized fuel into the fuel duct and allowing the fuel to circulate in an external fuel line. The fuel injection nozzle further comprises a coaxial air duct coaxially surrounding the fuel duct and being configured for providing that fuel exiting at the fuel nozzle at the distal end of the fuel duct is surrounded by an air stream that prevents the fuel from sticking to parts of the fuel injection nozzle. According to the invention, the fuel injection nozzle is provided with a Venturi nozzle formed at the distal end of the air duct surrounding the fuel nozzle of the fuel duct. The Venturi nozzle has an inner diameter that varies over the length of the Venturi nozzle, i.e. in the longitudinal direction of the fuel injection nozzle, and that is larger at the beginning of the Venturi nozzle and at the end of the Venturi nozzle than in the middle of the Venturi nozzle.

[0047] The inventor found that the Venturi nozzle at the end of the air duct and in front of the fuel nozzle at the end of the fuel duct improves vaporization fuel exiting the fuel duct even if the fuel is provided with lower than usual pressure. In prior art fuel injection nozzles, the vaporization of fuel is affected if the pressure of the fuel is too low. However, a higher fuel pressure leads to a larger amount of fuel being injected and thus leads a higher fuel consumption.

[0048] According to a preferred embodiment, the fuel injection nozzle comprises a baffle body that is arranged directly in line with the fuel duct in front of the fuel nozzle of the fuel duct. During operation, fuel exiting from the fuel nozzle at the end of the fuel duct hits the baffle body and thus is vaporized even if the pressure in the fuel duct is not high enough to cause vaporizing of the fuel alone. Thus, the baffle body further improves fuel vaporization.

[0049] Preferably, the baffle body has a cone shape with a cone tip facing away from the fuel nozzle and a baffle face facing towards the fuel nozzle. The cone shape of the baffle body supports the Venturi effect of the Venturi nozzle and helps controlling flow separation and forming of a wake downstream of the baffle body.

[0050] Preferably, the baffle body is held by a pin in the middle of the Venturi nozzle at the distal orifice of the air duct. The arrangement of the baffle body in the middle of the Venturi nozzle optimizes the Venturi effect and the acceleration of the injected fuel-air mixture.

[0051] Preferably, the fuel injection nozzle is further configured for adding hydrogen to the fuel-air mixture that is formed by the fuel injection nozzle. In particular, the fuel injection nozzle is configured for enriching the air in the coaxial air duct with 1% to 8% of hydrogen further, thus improving the mixture provided by the fuel injection nozzle.

[0052] The preferably reduced fuel pressure in the fuel duct results in less fuel being injected into the combustion chamber. While this generally improves the efficiency, it also leads to less power provided by the combusted mixture of fuel and air. Adding a small amount of hydrogen to the mixture does not increase the temperature of the combusted fuel mixture but makes the combustion 60% more clean than a prior art combustion.

[0053] The Venturi nozzle with the baffle body in front of the fuel nozzle of the fuel duct improves mixing of the fuel with air even at lower fuel pressures. Additionally adding some amount of hydrogen makes the combustion cleaner and thus reduces air pollution. In combination, a lower fuel consumption and less pollution is achieved by means of the novel fuel injection nozzle.

[0054] Further preferred features and advantages will be apparent from the disclosure of exemplary embodiments. Additional aspects of the present invention will become more readily apparent from the detailed description, particularly when taken together with the drawings.

Figure 1:

illustrates an embodiment of a turboprop engine comprising a gas turbine according to the invention;

Figure 2:

is a further, more detailed illustration of a turboprop engine comprising a gas turbine according to the invention;

Figure 3:

is a perspective view of a turbine shaft and a shaft tunnel for the turbine shaft;

Figure 4:

is an exploded perspective view of the turbine shaft, the shaft tunnel and two ball bearings that support the turbine shaft within the shaft tunnel;

Figures 5a and 5c:

are front views of an assembly comprising the turbine shaft and the shaft tunnel;

Figure 5b:

is a side elevated view of the assembly comprising the turbine shaft and the shaft tunnel;

Figure 5d:

is a longitudinal cross-sectional view of the assembly comprising the turbine shaft and the shaft tunnel as shown in figure 5b;

Figure 6a:

is a side view of the turbine shaft alone, illustrating grooves in a longitudinal section of the turbine shaft;

Figure 6b:

is a cross-sectional view perpendicular to the longitudinal axis of the turbine shaft;

Figure 6c:

is a cross-sectional view along the longitudinal axis of the turbine shaft;

Figures 7 and 8:

are schematic cross-sectional views perpendicular to the longitudinal axis of the turbine shaft illustrating the turbine shaft within the shaft tunnel;

Figures 9a and b:

are perspective views of the turbine shaft alone; and

Figure 10:

is a longitudinal cross-section of the gas turbine assembly with the combustion chamber, the gas turbine stage and the turbine shaft.

As illustrated in figures 1 and 2, the main components of a turboprop engine 10 are a gas turbine assembly 20 that can drive a main shaft 30 of a propeller 12 via a reduction gear. The reduction gear comprises a helical reduction gearing 40, a centrifugal clutch 34 and a planetary reduction gear 32.

[0055] The gas turbine assembly 20 comprises a compressor 22, for instance a centrifugal compressor with an impeller 22.1 that aspires air from the environment through an air intake 14. Compressed air provided by the centrifugal compressor 22 is fed through a diffuser 22.2 into an open space 23 surrounding a combustion chamber 24.1 of a combustor 24. The compressed air enters the combustion chamber 24.1 from the open space surrounding the combustion chamber 24.1. In the combustion chamber 24.1, the compressed air is mixed with fuel injected into the combustion chamber 24.1 by a fuel injection nozzle 24.2. The fuel combusts and hot exhaust gases are fed to the gas turbine 26. The combustion chamber 24.1 can for instance be an annular combustion chamber 24.1 with an outlet nozzle 24.3 for the hot exhaust that feeds the hot exhaust to the gas turbine, for instance an axial turbine 26. The axial turbine 26 drives the centrifugal compressor 22 and - via the planetary reduction gear 32 - the propeller 12.

[0056] The turboprop engine 10 as shown in figure 1 comprises a gas turbine assembly 20 with a single turbine stage 26. The gas turbine assembly 20 comprises a stationary nozzle guide vane 26.1 and a turbine wheel 26.2 that is connected to a gas turbine shaft 28. The gas turbine shaft 28 is also connected to the centrifugal compressor 22. Accordingly, gas turbine 26 can drive compressor 22 of turboprop engine 10. Turbine shaft 28 is also connected to a reduction gear 32 that in turn is connected to a propeller 12 by means of gas turbine assembly 20. For starting the gas turbine assembly 20, a starter/generator 36 is provided. The starter/generator 36 can drive the turbine shaft 28 via a starter gearbox 38.

[0057] During operation, the compressor 22 takes air from the environment, compresses the air and pushes the compressed air into the combustion chamber 24.1 wherein the compressed air is mixed with fuel and combusted by a spark. The combustion chamber nozzle 24.3 directs the resulting flow of hot combustion gases to the axial turbine 26. The axial turbine 26 comprises a single turbine stage with the nozzle guide vane 26.1 and the single turbine wheel 26.2. As mentioned above, the turbine wheel 26.2 drives the turbine shaft 28 that is connected to the centrifugal compressor 22 and the planetary reduction gear 32. The rotation speed of the turbine shaft 28 is reduced by the helical reduction gearing 40 and the planetary reduction gear 32 to a fraction in order to provide a rotation speed that is suitable for driving a standard propeller 12.

[0058] The turbine shaft 28 extends through a shaft tunnel 28.1 as can be seen in further detail in figures 3 and 5. At the longitudinal ends of shaft tunnel 28.1 ball bearings 28.2 for the turbine shaft 28 are arranged. Preferably, the turbine shaft 28 id made from steel.

[0059] Gas turbine engines need constant lubrication of the bearings 28.2 of the turbine shaft 28 due to the very high rotation speed during operation and due to the heat in the hot gas turbine parts that surround the turbine shaft 28 and the shaft tunnel 28.1, for instance the combustor 24 and the gas turbine 26. The lubricant not only lubricates the turbine shaft 28 but also cools the system down.

[0060] In prior art gas turbines, an oil pump is used to give the shaft these characteristics but like any mechanical part is subjected to possible damage due to the usage.

[0061] To avoid the need for an oil pump, the turbine shaft 28 and the shaft tunnel 28.1 are shaped to suck the necessary lubricant - for instance oil - and at the same time to generate an oil cloud in the interior of the shaft tunnel 28.1 that provides for a very good heat transfer between the parts and thus improves cooling.

[0062] A longitudinal section 28.3 of the turbine shaft 28 that extends within the shaft tunnel 28.1 between the ball bearings 28.2 is provided with slots or grooves 28.4 that are specifically shaped as a main cylindrical center 28.3.1 and two opposite side cones in the same longitudinal direction 28.3.1 taking the oil in the center and distributing it to both sides as an oil cloud due to the high speed rotation of the turbine shaft 28. The slots or grooves 28.4 are arranged around the circumference of the turbine shaft 28 and extend in the longitudinal direction of the turbine shaft 28. Preferably, 6 to 24 longitudinally extending grooves 28.4 are arranged around the circumference of the turbine shaft 28; see figure 6. The grooves 28.4 have a width between 2 mm and 4 mm and a depth between 1 mm to 3 mm.

[0063] The number of slots or grooves 28.4 and rips 28.5 can be 18 as shown in the illustrative example. However, the number of slots or grooves 28.4 and rips 28.5 may be between 12 and 24.

[0064] The longitudinal section 28.3 of the turbine shaft 28 that is provided with grooves 28.4 may taper towards the longitudinal ends of the longitudinal section 28.3 as is apparent from figure 4 and 6. Accordingly, the diameter of middle sub-section 28.3.1 of the longitudinal section 28.3 is larger than the tapering longitudinal end sections 28.3.2 of the longitudinal section 28.3. The depth of grooves 28.4 may correspond to (half of) the increase in diameter of the tapering longitudinal end sections 28.3.2. This results in a turbine shaft 28 having a longitudinal section 28.3 that is characterized by longitudinally extending rips 28.5 that taper towards the longitudinal ends and which are defined by the grooves 28.4 separating the rips from each other.

[0065] At least the longitudinal middle sub-section 28.3.1 of the turbine shaft section 28.3 provided with slots and/or grooves 28.4 has a very small clearance C with respect to an inner wall 28.1.1 of the shaft tunnel 28.1, i.e. the radial distance - i.e. the clearance - between the turbine shaft 28 and the shaft tunnel 28.1 is very small in the respective longitudinal sub-section 28.3.1; see figure 7.

[0066] The clearance may vary over time during the operation due to dynamic effects. In an axial-symmetric static situation, the clearance may be as small as 1 mm or less. For instance, an outer diameter of the subsection 18.3.1 of shaft 28 may be less than 2 mm smaller than an inner diameter of the shaft tunnel 28.1, during the normal function and due to the rotation speed and the heat, the variation of this clearance C can be up to 1 mm without causing a significant change in the performance but this can be controlled to using special materials like Titanium and INCOLOY maintaining the deformation as low as the best performance configuration between 0,5 mm and 0,8 mm.

[0067] These slots or grooves 28.4 cause - in combination with the small radial distance between the turbine shaft 28 and the shaft tunnel 28.1 - a suction effect, even almost a vacuum effect that sucks the lubricating and cooling oil into the shaft tunnel 28.1 and allows controlling the flow of the lubricant, for instance the flow of lubricating oil. This lubrication system is self-supported.

[0068] Figures 6a, 6b and 6c illustrate two actual dimensions of an implemented, exemplary embodiment. The actual dimensions shown for the embodiment for figure 6 are, however, not binding. The concept of the invention can be applied to different shaft sizes.

[0069] Regarding the best performing distance (clearance C) between the turbine shaft 28 and the inner wall 28.1.1 of the shaft tunnel 28.1 it is noted that due to the high rotation speed of the shaft the desired effect can be achieved with different distances C.

[0070] Due to the shape of the turbine shaft 28 that is cylindrical in the middle section 28.3.1 where it receives the oil and has two conical sections 28.3.2 extending from the middle section 28.3.1 in the direction towards the ball bearings 28.2, the turbine shaft 28 carries the oil to both ball bearings 28.2 because to the turbine shaft's tapering diameter the distance between the rips 28.5 and the inner wall 28.1.1 of the shaft tunnel 28.1 becomes larger thus reducing the suction or vacuum effect and spreading the oil for all the walls generating an oil cloud.

[0071] Figure 10 illustrates how the gas turbine shaft 28 is arranged within the gas turbine assembly 24 20. The shaft is connected to and thus driven by the gas turbine wheel 26.2 of the gas turbine stage 26. The shaft tunnel 28.1 is centrally arranged within the gas turbine assembly 20 and surrounded by the combustor 24.

Reference signs

[0072] 

10

turboprop engine

12

propeller

14

air intake

20

gas turbine assembly

22

compressor of the gas turbine assembly

22.1

impeller

22.2

diffuser

23

open space surrounding the combustor's combustion chamber

24

combustor

24.1

combustion chamber of gas turbine assembly

24.2

fuel injection nozzle

24.3

outlet nozzle / combustion chamber nozzle

26

axial gas turbine stage of the gas turbine assembly

26.1

nozzle gas vane of the axial gas turbine

26.2

turbine wheel of the axial gas turbine

28

gas turbine shaft

28.1

shaft tunnel for the gas turbine shaft

28.1.1

inner wall of the shaft tunnel 28.1

28.2

ball bearing

28.3

longitudinal shaft section within shaft tunnel 28.1

28.3.1

middle sub-section of the longitudinal shaft section 28.3

28.3.2

tapering end section of the longitudinal shaft section 28.3

28.4

slots or grooves extending in the longitudinal direction of shaft 28

28.5

longitudinally extending rips

30

main shaft, propeller shaft

32

planetary reduction gear

34

centrifugal clutch

36

starter/generator

38

starter gearbox

40

helical reduction gearing

C

minimum radial distance (clearance) between the turbine shaft and the shaft tunnel

Claims

1. Gas turbine assembly (20) with a central turbine shaft (28) that is connected to at least a gas turbine stage (26) of the gas turbine assembly (20), said turbine shaft (28) extending in a longitudinal direction within a shaft tunnel (28.1) and being supported by bearings (28.2) arranged at or near the longitudinal ends of the shaft tunnel (28.1),
characterized in that the turbine shaft (28) has a longitudinal shaft section (28.3) that extends within the shaft tunnel (28.1) and that is provided with a plurality of grooves (28.4) extending in the longitudinal direction of the shaft (28) and being arranged around the circumference of the shaft (28).

2. Gas turbine assembly (20) according to claim 1, wherein a longitudinal sub-section (28.3.1) of the turbine shaft section (28.3) has an outer diameter that is less than 1 mm smaller than an inner diameter of the shaft tunnel (28.1), resulting in a radial distance (clearance C) between the turbine shaft (28) and the shaft tunnel (28.1) that at average is smaller than 0,5 mm.

3. Gas turbine assembly (20) according to claim 1 or 2, wherein the bearings (28.2) are ball bearings.

4. Gas turbine assembly (20) according to at least one of claims 1 to 3, wherein 6 to 24 longitudinally extending grooves (28.4) are provided.

5. Gas turbine assembly (20) according to at least one of claims 1 to 4, wherein the grooves 28.4 have a width between 2 mm and 4 mm and a depth between 1 mm to 3 mm.

6. Gas turbine assembly (20) according to at least one of claims 1 to 3, wherein the longitudinal section (28.3) of the turbine shaft 28 that is provided with grooves (28.4) tapers towards the longitudinal ends of the longitudinal section (28.3)

7. Gas turbine assembly (20) according to claim 6, wherein the longitudinal section (28.3) has a middle sub-section (28.3.1) and tapering longitudinal end sections (28.3.2) that taper towards the longitudinal ends of the longitudinal section (28.3).

8. Gas turbine assembly (20) according to claim 7, wherein the depth of grooves (28.4) corresponds to half of the increase in diameter of the tapering longitudinal end sections (28.3.2).

9. Gas turbine assembly (20) according to claim 6 or 7, wherein the depth of grooves (28.4) is between 1 mm and 3 mm, in particular 2 mm.

10. Gas turbine assembly (20) according to at least one of claims 1 to 9, wherein the number of grooves (28.4) is between 12 and 24, in particular about 18.

11. Gas turbine assembly (20) according to at least one of claims 1 to 10, wherein rips (28.5) are defined by the grooves (28.4) and wherein the width of the grooves (28.4) corresponds to the width of the rips (28.5).

12. Gas turbine assembly (20) according to claim 11, wherein the width of the grooves (28.4) and/or the rips (28.5) is between 2 mm and 4 mm.

13. Turboprop engine (10) with a gas turbine assembly (20) according to at least one of claims 1 to 12.

Drawing