DE102011102359A1 - Gas turbine for use in gas turbine engine i.e. turbo-engine, of aircraft, has guide tracks extending in axial direction with respect to axis, and blades moved along curved tracks by piston-cylinder that has front piston and rear piston - Google Patents

Abstract

The turbine has a fan (12) arranged in an inlet area of the turbine and comprising a hub (29). Fan blades (30) are arranged at the hub, and guide tracks (31) are formed at outer circumference of the hub. The guide tracks extend in an axial direction with respect to an engine axis (1) and comprise a curvature, where blade roots (36) of the blades are supported at the guide tracks. The blades are moved along the curved guide paths by a piston-cylinder that has a front piston and a rear piston, where the pistons are supplied with hydraulic fluid by hydraulic channels.

Description

The invention relates to an aircraft gas turbine with a fan whose blades / blades are finely adjustable and in particular to an aircraft gas turbine with a fan which is arranged in the inflow region, wherein the fan has a hub on which a plurality of fan blades are mounted.

The fan of an aircraft gas turbine forms a first compressor area, through which incoming air is compressed, which is subsequently fed to a core engine and a bypass duct.

In the constructions known from the prior art, the blades of the fan are optimized for a particular engine state (flight mode) (pitch angle). In other engine situations, however, the efficiency of the fan is not optimal. These engine states may be, for example, the takeoff phase, the cruise phase, or the descent phase.

The optimization of the leaves of the fan is done by optimizing the angle of attack. In a non-optimized operating condition, the angle of attack is usually larger, inducing more turbulent flow at certain areas of the fan blade. This results in unfavorable force effects on the fan blade, which lead to unwanted vibrations (flutter). These vibrations have a negative influence on the power transmission capacity of the material of the fan blades and cause unwanted noise. The prior art has attempted to remedy this issue with heavier and stiffer fan blades, which in turn leads to the need to make the hub of the fan harder and more stable.

The ability to rotate a fan blade similar to a propeller blade or a stator blade of a compressor about a radial axis, is only partially possible because the fan blade is located very close to the periphery and therefore is not sufficient space available.

The invention has for its object to provide an aircraft gas turbine with a fan, which while avoiding the disadvantages of the prior art, an optimization of the flow conditions of the fan is possible.

According to the invention the object is achieved by the combination of features of claim 1, the dependent claims show further advantageous embodiments of the invention.

According to the invention it is thus provided that guideways are formed on the circumference of the hub, which extend substantially in the axial direction, relative to an engine axis, and have a curvature, each fan blade is mounted by means of its blade root on a guideway and slidable along the curved guideway is.

Due to the curvature of the guide track, which serves to support the blade root of the fan blade, a displacement of the fan blades relative to the hub leads to a change in the angle of attack of the fan blade. It has certainly been shown according to the invention that even a slight longitudinal displacement along the guideway leads to an effective adjustment of the fan blades. Thus, it is possible to achieve optimization of the fan with simple means for different flight conditions. By optimizing the pitch of the fan blades, it is possible to eliminate the vibration problems associated with prior art designs. Furthermore, it is possible to make both the fan blades and the hubs easier. This reduces the total weight. It is thus possible to operate the fan in an optimum manner over the entire operating states of the aircraft gas turbine (flight envelope). This also results in a significant noise reduction, since unwanted vibration and flow conditions are avoided.

In a particularly favorable development of the invention, the fan blades are displaced relative to the hub by means of at least one double-acting piston-cylinder unit. The piston-cylinder unit can be mounted directly on the hub and operated by present in the gas turbine oil supply systems. It is thus a simple, inexpensive operability, which does not require significant modifications to the rest of the aircraft gas turbine.

Furthermore, it is advantageous if the piston-cylinder unit is provided with a front arm and a rear arm for engagement with a front and rear portion of the fan blade. Thus, pressure forces are exerted on the fan blade, in particular on its blade root, by the piston-cylinder unit. This leads to a reliable adjustment.

Furthermore, it is advantageous if the contact areas between the piston-cylinder unit and the fan blade blade root) are spherical. This simplifies the geometric design, as always a good pressure transfer is ensured by the piston-cylinder unit on the blade root.

In order to minimize the forces required to move the fan blades, it may be particularly advantageous if the fan blades are each mounted with their blade root on the hub by means of roller bearings. In particular slide guides with rolling bearings prove to be particularly advantageous, as they are already known from linear guide systems, for example in machine tools.

In the following the invention will be described by means of an embodiment in conjunction with the drawing. Showing:

1 a schematic representation of a gas turbine engine according to the present invention,

2 1 is a partial perspective sectional view of an embodiment of a hub of a fan according to the invention;

3 an enlarged sectional view analog 2 .

4 a perspective view to illustrate the displacement movements, and

5 a partial perspective view of a storage area.

1 shows a schematic representation of a gas turbine engine according to the present invention.

The gas turbine engine 10 according to 1 is an example of a turbomachine to which the invention may find application. However, it will be understood from the following that the invention can be used with other turbomachinery as well. The engine 10 is formed in a conventional manner and comprises one air inlet in succession in the flow direction 11 , a circulating in a housing fan 12 , an intermediate pressure compressor 13 , a high pressure compressor 14 , Combustion chambers 15 , a high-pressure turbine 16 , an intermediate-pressure turbine 17 and a low-pressure turbine 18 and an exhaust nozzle 19 all around a central engine axis 1 are arranged.

The intermediate pressure compressor 13 and the high pressure compressor 14 each comprise a plurality of stages, each of which comprises a circumferentially extending array of fixed stationary vanes 20 commonly referred to as stator blades and radially inward of the engine casing 21 in an annular flow channel through the compressors 13 . 14 protrude. The compressors further include an array of compressor blades 22 on, which is radially outward from a rotatable drum or disc 26 project with hubs 27 the high-pressure turbine 16 or the intermediate-pressure turbine 17 are coupled.

The turbine sections 16 . 17 . 18 have similar steps, comprising an array of fixed vanes 23 extending radially inward from the housing 21 into the annular flow channel through the turbines 16 . 17 . 18 project, and a subsequent arrangement of turbine blades 24 facing outward from a rotatable hub 27 protrude. The compressor drum or compressor disk 26 and the blades arranged thereon 22 as well as the turbine rotor hub 27 and the turbine blades disposed thereon 24 rotate around the engine axis during operation 1 ,

The 2 and 3 each show in partial sectional view of a hub 29 , This has distributed over the circumference a plurality of guideways 31 on which (see 4 ) are arcuately curved and substantially in the axial direction, relative to the engine axis 1 , extend.

At the individual guideways 31 are fan blades 30 stored, which has a blade foot 36 exhibit. This is contoured in cross section and matching the cross section of the guideway 31 formed so that the blade foot 36 in the guideway 31 is insertable.

According to the invention is a double-acting piston-cylinder unit 32 provided, which has a front piston 37 and a rear piston 38 includes. These are via hydraulic channels 39 and 40 ( 3 ) supplied with hydraulic fluid. In particular from the representation of 3 it follows that a pressurization of a front working chamber 41 or a rear working chamber 42 leads to a shift of the double piston construction (the two pistons 37 and 38 are firmly connected).

The piston-cylinder unit 32 also includes a front arm 33 and a back arm 34 , which are offered U-shaped and in contact with the blade foot 36 are located. In each case there is a spherical contact 43 which is formed, for example, by a part-spherical surface.

The 4 shows by the dotted line the direction of movement of the blade root 36 or the fan blade 30 along the guideway 31 , It can be seen that by the arcuate design of the guideway 31 during the shift of the fan blade 30 the angle of attack changes. For example, a displacement of 20 mm results in a change in the angle of attack of 5 °.

The 5 schematically shows a rolling bearing 35 , which is similar to a carriage guide and includes a plurality of received in a cage member rolling bearing elements. This results in a considerable reduction in friction, which results in lower locking forces.

LIST OF REFERENCE NUMBERS

1

Engine axis

10

Gas turbine engine

11

air intake

12

in the housing revolving fan

13

Intermediate pressure compressor

14

High pressure compressor

15

combustors

16

High-pressure turbine

17

Intermediate pressure turbine

18

Low-pressure turbine

19

exhaust nozzle

20

vanes

21

Engine casing

22

Compressor blades

23

vanes

24

turbine blades

26

Compressor drum or disc

27

Turbinenrotornabe

28

outlet cone

29

hub

30

fan blade

31

guideway

32

Piston Cylinder

33

Front arm

34

Back arm

35

Rolling bearings / carriage guide

36

blade

37

Front piston

38

Rear piston

39

hydraulic channel

40

hydraulic channel

41

Front working chamber

42

Rear working chamber

43

Spherical contact

44

Nose cone

Claims (5)

Aircraft gas turbine with a fan ( 12 ) in the inflow region of the aircraft gas turbine, wherein the fan ( 12 ) a hub ( 29 ), at the plurality of fan blades ( 30 ) Are arranged, characterized in that (at the outer periphery of the hub 29 ) Guideways ( 31 ) are formed, which are substantially in the axial direction, relative to the engine axis ( 1 ), and have a curvature, each fan blade ( 30 ) by means of her blade foot ( 36 ) on a guideway ( 31 ) and along the curved guideway ( 31 ) is displaceable. Aircraft gas turbine according to claim 1, characterized in that the fan blades ( 30 ) by means of at least one double-acting piston-cylinder unit ( 32 ) are displaceable. Aircraft gas turbine according to claim 2, characterized in that the piston-cylinder unit ( 32 ) each have a front arm ( 33 ) and a rear arm ( 34 ) for engagement with a front and a rear portion of the fan blade ( 30 ) or a blade root ( 36 ). Aircraft gas turbine according to claim 3, characterized in that the contact areas between the front arm ( 33 ) and the rear arm ( 34 ) and the fan blade ( 30 ) are each spherically formed. Aircraft gas turbine according to one of claims 1 to 4, characterized in that the fan blades ( 30 ) with their shovel feet ( 36 ) each by means of rolling bearings ( 35 ), in particular a slide guide with rolling bearings on the hub ( 29 ) are stored.

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