DE102022121803A1 - Gas turbine engine for an aircraft - Google Patents

Abstract

A gas turbine engine (10) for an aircraft with an engine core (11) is described. Oil can be introduced into a collecting channel (41) from a radially inner region (42). In addition, oil can be discharged from the collecting channel (41) via a discharge opening (46) in the direction of at least one consumer of a planetary gear (30). A radially outer region (58) of the discharge opening (46) is arranged on a circle around a rotation axis (9) of the rotatable component (34). The radius (rSA) of the circle is larger than the radius (riR) of an inner region (42) of the collecting trough (41) and less than or equal to the radially outer radius (raR) of the collecting trough (41). Furthermore, the radius (rSA) of the circle is smaller than the radius (rGSFmax or rGSFmin) of outsides (53B) of fan blade roots (53A). Additionally or alternatively, the radius (raVS) of an outside (56) of a compressor disk (57) of the compressor (14) is larger than the radius (rSA) of the circle.

Description

The present disclosure relates to a gas turbine engine for an aircraft having an engine core. The engine core comprises at least one compressor and at least one turbine as well as a core shaft connecting the compressor to the turbine and limits a flow cross section for a core air flow. A fan is provided upstream of the engine core and is connected to the core shaft via a planetary gear.

A gas turbine engine of an aircraft with a planetary gear is known from practice. The planetary gear includes a sun gear, a ring gear fixed to the housing and a rotatable planet carrier through which a fan can be driven. Several planet gears mesh with the sun gear and the ring gear. Oil is guided via a collecting channel connected to the planetary carrier of the gearbox in the direction of the meshing between the planetary gears and the ring gear as well as with the sun gear. The collecting trough extends in the circumferential direction of the planet carrier and is open radially on the inside. Oil is fed into the collecting trough via the radially inner opening.

In addition, the gas turbine engine comprises an engine core which is formed with at least one compressor and at least one turbine as well as with a core shaft connecting the compressor to the turbine and which limits a flow cross section for a core air flow. A fan is provided upstream of the engine core and is connected to the core shaft via the planetary gear. The blower has a so-called blower spinner with a blower spinner cone and blower blades connected to it in the area of blower blade roots.

Outer sides of the fan blade roots limit the flow cross section of the air in this axial region of a gas turbine engine. The outer sides of the fan blade feet form a closed ring area in the circumferential direction that widens at least approximately in the manner of a truncated cone, the diameter of which is smaller in the area of the front edges of the fan blades than in the area of the rear edges of the fan blades.

Air flows into the flow area of the engine core along the outside of the fan spinner cone and the outside of the fan blade roots. The flow cross section of the engine core is limited radially on the inside, among other things, by a radial outside of a compressor disk of the compressor, the diameter of the outside of the compressor disk increasing depending on the respective structural design of the gas turbine engine from a side facing the fan towards a side facing away from the fan decreases. The gearbox is arranged axially between the fan and the compressor and radially within the flow cross section of the engine core.

In order to avoid the risk of undersupply of planetary wheel bearings and in the area of tooth meshing between the components of the transmission, it is necessary that the oil introduced into the collecting trough is accelerated within the collecting trough to the extent necessary for forwarding in the direction of the planetary wheel bearings and the tooth meshing. In addition, an air flow path between the fan spinner cone and the compressor of the engine core should be designed in such a way that the air can be introduced into the flow cross section of the engine core with the lowest possible flow losses in order to be able to operate the gas turbine engine with high efficiency.

The present disclosure is based on the object of providing a gas turbine engine with a planetary gear, in which a sufficient oil supply to the planetary gear is provided and which can be operated with high efficiency.

This task is achieved with a gas turbine engine with the features of claim 1. Advantageous further developments are the subject of the subclaims and the following description.

According to a first aspect of the present disclosure, a gas turbine engine for an aircraft having an engine core is provided. The engine core comprises at least one compressor and at least one turbine as well as a core shaft connecting the compressor to the turbine and limits a flow cross section for a core air flow. A fan is provided upstream of the engine core, which is connected to the core shaft via a planetary gear. The blower has a blower spinner, which includes a blower spinner cone and blower blades connected thereto in the area of blower blade roots. Air flows into the flow cross section of the engine core along an outer region of the fan spinner cone and along the outer sides of the fan blade roots. The fan spinner cone and the fan blade bases can be designed to optimize flow technology and the outer sides of which at least partially form the shape of an ellipsoid.

The flow cross section of the engine core is limited radially on the inside, among other things, by a radial outside of a compressor disk of the compressor. In the event that the compressor is designed in multiple stages and has a plurality of compressor disks arranged one behind the other in the axial direction of the engine core, in the present case it is the outside of the compressor disk of a first compressor stage of the compressor, which faces the fan and represents a compressor disk of a so-called low-pressure compressor.

In addition, the gear is arranged in the axial direction between the fan and the compressor. The transmission comprises a rotatably mounted component which is designed with an at least approximately rotationally symmetrical collecting channel. Oil can be introduced into the collecting channel starting from its radially inner area. In addition, oil can be discharged from the collecting trough via a discharge opening of a flow branch in the direction of at least one consumer of the planetary gear.

A radially outer region of the discharge opening of the flow branch is arranged on a circle around a rotation axis of the rotatable component, the radius of the circle being larger than the radius of the inner region of the collecting trough and less than or equal to the radially outer radius of the collecting trough.

Furthermore, the radius on which the radially outer region of the discharge opening lies is smaller than the radius of the outside of the fan blade roots and/or smaller than the radius of the outside of the compressor disk of the compressor.

The radius of the outer sides of the fan blade roots corresponds to the radius that the outer sides of the fan blade roots have in the areas that face the engine core and which is preferably the largest radius of the outer sides of the fan blade roots.

Furthermore, in the present case, the radius of the outside of the compressor disk of the compressor corresponds to the largest radius of the outside of the compressor disk, which, depending on the design of the compressor, can be present in the area of leading edges of compressor blades or in the area of trailing edges of the compressor blades of the compressor in the direction of flow of air through the compressor .

The advantageous coordination of the radii of the gas turbine engine described in more detail above ensures, on the one hand, that the oil that is introduced into the collecting trough is, during operation of the gas turbine engine, in the collecting trough to the required extent from the centrifugal force then acting on the oil, starting from the radially inner region of the collecting trough is accelerated in the direction of the discharge opening of the flow branch and from there forwarded in the direction of a consumer of the planetary gear. This ensures a desired high supply of oil to the planetary gear and avoids an undersupply of a planet wheel bearing and/or tooth meshing between components of the gear.

In addition, the advantageous coordination of the radii of the gas turbine engine described in more detail above ensures that the air flow, which flows from the fan blade roots in the direction of the compressor, flows with the lowest possible flow losses due to the arrangement of the gear in the axial direction of the gas turbine engine between the fan and the compressor enters the flow cross section of the engine core.

For this purpose, a ratio between the radius on which the radially outer region of the discharge opening lies and the radius of the outside of the compressor disk of the compressor can be in a range of 0.7 to 0.9.

Furthermore, a ratio between the inner radius of the collecting trough and the radius of a radially inner region of the annular disk-shaped compressor disk of the compressor, which delimits a central opening of the compressor disk, can be in a range of 0.8 to 1.2, on the one hand to ensure an oil supply to the To ensure planetary gearing and also to be able to introduce an air flow with the lowest possible flow losses into the flow cross section of the engine core.

A ratio between the radius of the circle on which the radially outer region of the discharge opening lies and the largest radius of the outer sides of the fan blade roots can be in a range of 0.4 to 0.7, on the one hand to ensure an oil supply to the planetary gear and additionally to be able to introduce an air flow into the flow cross section of the engine core with the lowest possible flow losses.

A ratio between the radius on which the radially outer region of the discharge opening lies and the smallest radius of the outer sides of the fan blade roots can be in a range from 0.7 to 0.9 in order to ensure an oil supply to the planetary gear and also to be able to introduce an air flow with the lowest possible flow losses into the flow cross section of the engine core.

in Abhängigkeit des äußeren Radius der Auffangrinne festgelegt sein. Dabei entspricht die Variable x einem Verhältnis, das gemäß dem formelmäßigen Zusammenhang $$x=\sqrt{\frac{r_{aR}^2-r_{iR}^2}{r_{aR}^2}}$$

in Abhängigkeit des inneren Radius und des äußeren Radius der Auffangrinne festgelegt sein kann.The inner radius of the collecting channel can be determined according to the formula $$r_{iR}=\sqrt{r_{aR}^2-x^2\cdot r_{ar}^2}$$

be determined depending on the outer radius of the collecting channel. The variable x corresponds to a ratio that corresponds to the formula $$x=\sqrt{\frac{r_{aR}^2-r_{iR}^2}{r_{aR}^2}}$$

can be determined depending on the inner radius and the outer radius of the collecting channel.

The ratio x can have values of 0.5 to 0.8, preferably 0.65 to 0.75, in order to ensure, on the one hand, a desired high oil supply to the planetary gear and, on the other hand, to ensure that the air flow between the fan spinner cone and the compressor is as flow-optimized as possible achieve.

In an embodiment of the gas turbine engine according to the present disclosure, in which a desired high oil supply to the planetary gear and at the same time a flow-optimized introduction of an air volume flow into the flow cross section of the engine core is achieved, the rotatable component is designed as a planetary carrier of the planetary gear.

A ratio between the inner radius of the collecting trough and a radially outer radius of a bearing of the planet carrier can be in a range of 1.1 to 1.7 in order to ensure, on the one hand, a desired high oil supply to the planetary gear and, on the other hand, to ensure that the air flow between the blower spinner is as flow-optimized as possible -Cone and the compressor to achieve.

Furthermore, there is the possibility that oil can be introduced into the collecting trough from an oil supply fixed to the housing, which is arranged in the axial direction of the planetary gear next to the collecting trough and in the radial direction between the inner radius of the collecting trough and the outer radius of the bearing of the planet carrier.

In an embodiment of the gas turbine engine characterized by small component dimensions in the axial direction, the transmission is arranged in the axial direction between the fan and the compressor and radially within the flow cross section of the engine core.

The present disclosure is not limited to the specified combinations of features of the independent claims or the claims dependent thereon. There are also possibilities to combine individual features with one another, including those that emerge from the claims, the following description of embodiments and directly from the drawing. The reference of the claims to the drawings through the use of reference numerals is not intended to limit the scope of the claims.

• 1 eine schematisierte Längsschnittansicht eines Gasturbinentriebwerkes;
• 2 eine vergrößerte Teillängsschnittansicht eines stromaufwärtigen Abschnitts eines Gasturbinentriebwerks;
• 3 eine Alleindarstellung eines Getriebes für ein Gasturbinentriebwerk; und
• 4 eine Schnittansicht des Getriebes entlang einer in 3 näher gekennzeichneten Schnittlinie IV-IV.

Preferred further developments result from the subclaims and the following description. Embodiments of the subject matter according to the present disclosure are, without being limited to this, explained in more detail with reference to the drawing. This shows:

• 1 a schematic longitudinal sectional view of a gas turbine engine;
• 2 an enlarged partial longitudinal sectional view of an upstream portion of a gas turbine engine;
• 3 a sole representation of a gearbox for a gas turbine engine; and
• 4 a sectional view of the gearbox along an in 3 Section line IV-IV marked in more detail.

1 illustrates a gas turbine engine 10 with a main axis of rotation 9. The gas turbine engine 10 includes an air inlet 12 and a thrust fan or fan 23 that generates two air streams: a core air stream A and a bypass air stream B. In addition, the gas turbine engine 10 includes an engine core 11, through through which the core air flow A flows. The engine core 11 is designed in axial flow order with a low-pressure compressor 14, with a high-pressure compressor 15, with a combustion device 16, with a high-pressure turbine 17, with a low-pressure turbine 19 and with a core thrust nozzle 20. An engine nacelle 21 surrounds the gas turbine engine 10 and defines a bypass thrust nozzle 18 and a bypass duct 22 through which the bypass air stream B flows. The blower 23 is operatively connected to the low-pressure compressor 14 and to the low-pressure turbine 19 via a core shaft 26 and a planetary gear 30 and is driven by the low-pressure turbine 19. The planetary gear 30 is in the axial direction X between the Blower 23 and the low-pressure compressor 14 arranged and designed as a reduction gear.

During operation of the gas turbine engine 10, the core air flow A is accelerated and compressed by the low-pressure compressor 14 and passed into the high-pressure compressor 15, where further compression takes place. The compressed air expelled from the high pressure compressor 15 is directed into the combustion device 16 where it is mixed with fuel and the mixture is burned. The resulting hot combustion products then propagate through and thereby drive the high pressure and low pressure turbines 17, 19 before being expelled through the nozzle 20 to provide some thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 through a suitable connecting shaft 27, which is also referred to as a core shaft. The fan 23 generally provides the majority of the thrust force.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in 2 shown. The core shaft 26 is coupled to a sun gear 28 of the planetary gear 30. Several planet gears 32 mesh with the sun gear 28 and are each rotatable 3 Planet bolts 42 shown in more detail are arranged, which are firmly connected to the planet carrier 34. The transmission 30 includes four planetary gears 32 and can be designed with 3 to 7 planetary gears 32 depending on the respective application.

The planet carrier 34 restricts the planet gears 32 from rotating synchronously about the sun gear 28 while allowing each planet gear 32 to rotate about its own axis on the planet pins 42, which are static axes. The planet carrier 34 is coupled to the fan 23 via linkage 36 in order to drive the fan 23 in rotation about the engine axis 9. An external gear or ring gear 38, which is coupled to a stationary support structure 24 via linkage 40, meshes with the planet gears 32.

At the in 2 and 3 Planetary gear 30 shown as an example, the planet carrier 34 is in operative connection with the blower 23 via the linkage 36. In principle, there is also the possibility that the planetary gear 30 has a so-called star arrangement, in which the planet carrier 34 is designed to be non-rotatable and the ring gear 38 is designed to be rotatable. With such an arrangement, the fan 23 is driven by the ring gear 38. As a further alternative example, the planetary gear 30 can also be designed as a differential gear, in which, in addition to the sun gear 28, both the ring gear 38 and the planet carrier 34 are designed to be rotatable.

The geometry of the gas turbine engine 10 and components thereof are defined by a conventional orthogonal axis system having an axial direction X (aligned with the axis of rotation 9) and a radial direction Y (in the bottom-up direction). 1 ). In addition, a circumferential direction U (perpendicular to the view in 1 ) of the gas turbine engine 10 is used in the description in order to be able to clearly explain the aspects described in more detail below.

3 and 4 show an orientation of the planetary gear 30 in its installed position in the gas turbine engine 10 during horizontal flight of an aircraft equipped with the gas turbine engine 10. The rotatable planet carrier 34 of the transmission 30 is in the in 3 and 4 shown in more detail on its side facing the shaft 26 with a rotationally symmetrical collecting channel 41. The collecting trough 41 is arranged coaxially to the axis of rotation 9 of the planet carrier 34 and the sun gear 28 and is designed in a radially inner region 42 with an opening 43 extending in the circumferential direction U of the collecting trough 41. Starting from the radially inner diameter of the collecting trough 41, oil can be introduced into the collecting trough 41 from an oil supply 44 fixed to the housing via the opening 43.

An introduction direction E of the oil into the collecting channel 41 starting from the oil supply 44 is in the figure 4 in a manner shown in more detail in a cross-sectional plane QS and encloses an angle α with the axial direction X, which is equal to 90 °. Furthermore, the introduction direction E of the oil into the collecting channel 41 includes an angle β with the radial direction Y, depending on the respective application, which is greater than or equal to 0° and less than 90°. The oil is introduced into the collecting trough 41 tangentially and in the direction of rotation of the collecting trough 41 at an angle value of the angle β equal to 90 °. In contrast to this, the introduction direction E is equal to the Y direction when the angle β is equal to 0°.

Alternatively, there is also the possibility that the oil is introduced into the collecting channel 41 at an angle α 'with an introduction direction E' with respect to the axial direction X. The angle α' can have values between 60° and 90°, preferably between 75° and 90° or between 80° and 90°, with the oil supply mounted with such an orientation in 4 is identified in more detail under the reference number 44 '.

In order to be able to guide the oil introduced into the collecting trough 41 from the collecting trough 41 into the area of the bearings of the planet gears 32, the collecting trough 41 has, in a radially outer region 45, a plurality of outlet openings 46 of a so-called flow branch 46A, which are distributed over the circumference of the collecting trough 41 the oil on. Via the outlet openings 46, the oil introduced into the collecting trough 41 with a desired impulse via the oil supply 44 is accelerated outwards in the radial direction y in addition to the imposed impulse when the planet carrier 34 is rotating due to the centrifugal force then acting on the oil in the collecting trough 41 , can first be diverted from the collecting trough 41 and is forwarded from there via line areas 47 of the planet carrier 34 in the axial direction X of the planetary gear 30. The line areas 47 are in operative connection with several branch lines 48 which run inwards from the line areas 47 in the radial direction Y, the mouth areas 49 of which each lie in the area of the bearings of the planet gears 32.

The oil supply 44 comprises an oil nozzle 50. An outlet opening 51 of the oil nozzle 50 is arranged at a distance in the radial direction Y from the opening 43 of the collecting channel 41 for the oil. The oil is discharged from the oil nozzle 50 with a defined supply pressure and, depending on the design of the outlet opening 51 of the oil nozzle 50, injected or sprayed into the collecting trough 41 with such a pulse that the oil in the collecting trough 41 within short operating times starting from the opening 43 of the Collecting trough 41 flows essentially in the radial direction Y to the outlet openings 46 of the collecting trough 41. The aim is for the oil to be introduced into the line areas 47 via the outlet openings 46 of the collecting channel 41 at a flow speed that ensures a desired oil supply to the bearings of the planet gears 32.

The oil nozzle 50 is on the in 3 and 4 shown embodiment of the transmission 30 is arranged above the horizontally extending longitudinal center plane WLM and next to the vertically extending longitudinal center plane VLM. Alternatively, there is also the possibility of positioning the oil nozzle in the vertical direction of the gas turbine engine 10 at the level of the horizontal longitudinal center plane WLM between the sun gear 28 and the collecting trough 41. Furthermore, the oil supply 44 can also include a plurality of oil nozzles 50, wherein at least one oil nozzle can be arranged at the level of the horizontal longitudinal center plane WLM on one side of the sun gear 28 and essentially between the sun gear 28 and the collecting trough 41, while another oil nozzle is positioned at the level of the horizontally extending longitudinal center plane WLM on the opposite side of the sun gear 28 essentially between the sun gear 28 and the collecting trough 41.

Any further oil nozzles of the oil supply are then arranged at the same distance from one another in the area between the sun gear 28 and the collecting trough 41 in relation to the two oil nozzles arranged at the level of the horizontal longitudinal center plane WLM in the circumferential direction U of the sun gear 28 or the collecting trough 41.

The oil nozzles arranged radially within the collecting trough 41 can be positioned in the axial direction of extension of the collecting trough 41 centrally between the areas delimiting the collecting trough 41 in the axial direction. The introduced oil is then introduced as uniformly as possible across the axial width of the collecting channel 41.

Alternatively or additionally, the sun gear, the planetary gears and/or the ring gear can also be designed in the manner described above with a collecting channel into which oil can be introduced via a corresponding oil supply in order to be able to supply consumers of the planetary gear 30 with oil.

The blower 23 includes a blower spinner 52 with a blower spinner cone 52A and fan blades 53 connected thereto in the region of fan blade roots 53A. Air flows from the air inlet 12 along an outer region 54 of the blower spinner cone 52A and from outer sides 53B of the fan blade roots 53A into a flow cross section 55 of the engine core 11. The flow cross section 55 of the engine core 11 is delimited radially on the inside by, among other things, a radial outside 56 of a compressor disk 57 of the low-pressure compressor 14.

Radially outer regions 58 of the discharge openings 46 of the flow branch 46A for the oil from the collecting channel 41 are arranged on a circle around the main axis of rotation 9 of the transmission 30. The radius r _SA of the circle is larger than the radius r _iR of the inner region 42 of the collecting trough 41. In addition, the radius r _SA of the circle is smaller or equal to the radially outer radius r _aR of the collecting trough 41.

In addition, the radius r _SA on which the radially outer regions 58 of the discharge openings 46 lie is smaller than the largest radius r _GSFmax of the outer sides 53B of the fan blade roots 53A, with the outer sides 53B of the fan blade roots 53A in the present case having the largest radius r _GSFmax in the flow direction A in the area of trailing edges 53C of the fan blades 53. Alternatively In addition to this or in addition to this, the radius r _aVS of the outside 56 of the compressor disk 57, which is the largest radius of the outside 56 of the compressor disk 57 and is present in the area of leading edges 62 of compressor blades 63 of the low-pressure compressor 14, can be larger than the radius r _SA which the radially outer regions 58 of the discharge openings 46 lie.

Depending on the respective application, a ratio between the radius r _SA on which the radially outer regions 56 of the discharge openings 46 lie and the largest radius r _aVS of the outside 56 of the compressor disk 57 of the compressor 14 can be in a range from 0.7 to 0.9 lie.

In addition, it can be provided that a ratio between the inner radius r _iR of the collecting trough 41 and the radius r _iVS of a radially inner region 59 of the annular disk-shaped compressor disk 57 of the low-pressure compressor 14, which delimits a central opening 60 of the compressor disk 57, in a range of 0.8 to 1.2.

In addition, a ratio between the radius r _SA on which the radially outer regions 58 of the discharge openings 46 lie and the largest radius r _GSFmax of the outer sides 53B of the fan blade roots 53A can be in a range of 0.4 to 0.7.

In addition, it can also be provided that a ratio between the radius r _SA on which the radially outer regions 58 of the discharge openings 46 lie and the smallest radius r _GSFmin of the outer sides 53B of the fan blade roots 53A is in a range from 0.7 to 0 .9 lies. The outer sides 53B of the fan blade roots 53A have the smallest radius r _GSFmin in the area of the front edges 53D of the fan blades 53.

in Abhängigkeit des äußeren Radius r_aR der Auffangrinne 41 festgelegt. Die Variable x entspricht einem Verhältnis, das gemäß dem formelmäßigen Zusammenhang $$x=\sqrt{\frac{r_{aR}^2-r_{iR}^2}{r_{aR}^2}}$$

in Abhängigkeit des inneren Radius R_iR und des äußeren Radius r_aR der Auffangrinne 41 festgelegt ist. Dabei kann das Verhältnis x in Abhängigkeit des jeweils vorliegenden Anwendungsfalles Werte von 0,5 bis 0,8, vorzugsweise von 0,65 bis 0,75 aufweisen.The inner radius r _iR of the collecting channel 41 is according to the formula relationship $$r_{iR}=\sqrt{r_{aR}^2-x^2\cdot r_{aR}^2}$$

depending on the outer radius r _aR of the collecting channel 41. The variable x corresponds to a ratio that corresponds to the formula $$x=\sqrt{\frac{r_{aR}^2-r_{iR}^2}{r_{aR}^2}}$$

is determined depending on the inner radius R _iR and the outer radius r _aR of the collecting channel 41. Depending on the particular application, the ratio x can have values of 0.5 to 0.8, preferably 0.65 to 0.75.

In addition, there is also the possibility that a ratio between the inner radius r _iR of the collecting trough 41 and a radially outer radius r _aLPT of a bearing 61 of the planet carrier 34 has values in a range from 1.1 to 1.7, depending on the respective application .

The oil supply 44 fixed to the housing, via which oil can be introduced into the collecting channel 41, can also be in the axial direction X of the planetary gear 30 next to the collecting channel 41 and in the radial direction Y between the inner radius r _iR of the collecting channel 41 and the outer radius r _aLPT of the bearing 61 of the planet carrier 34 may be arranged.

It is to be understood that the invention is not limited to the embodiments described above and various modifications and improvements may be made without departing from the concepts described herein. Any of the features may be used separately or in combination with any other features unless they are mutually exclusive, and the disclosure extends to and includes all combinations and subcombinations of one or more features described herein.

Reference symbol list

9

Main axis of rotation

10

Gas turbine engine

11

engine core

12

Air intake

14

Low pressure compressor

15

High pressure compressor

16

Combustion device

17

High pressure turbine

18

Bypass thrust nozzle

19

Low pressure turbine

20

Core thruster

21

Engine nacelle

22

Bypass channel

23

Thrust blower

24

Support structure

26

Shaft, connecting shaft

27

connecting shaft

28

sun gear

30

Gears, planetary gears

32

Planetary gear

34

Planet carrier

36

linkage

38

ring gear

40

linkage

41

gutter

42

radially inner area of the collecting channel

43

inner opening of the collecting channel

44, 44'

Oil supply

45

radially outer area of the collecting channel

46

discharge opening

46A

Flow diversion

47

Management area

48

Stub line

49

Mouth area

50

oil nozzle

51

Exhaust opening

52

Blower spinner

52A

Blower spinner cone

53

Fan blades

53A

Fan blade bases

53B

Outside of the fan blade bases

53C

Rear edges of the fan blades

53D

Leading edges of the fan blades

54

outer area of the blower spinner cone

55

Flow cross section of the engine core

56

radial outside of the compressor disk

57

Compressor disk

58

radially outer area of the discharge opening

59

Radial inner area of the annular disk-shaped compressor disk

60

central opening of the compressor disk

61

Planet carrier bearing

62

Leading edges of the compressor blades of the low-pressure compressor

63

Compressor blades of the low pressure compressor

A

Core airflow

b

Bypass airflow

E, E'

Direction of introduction

QS

Cross-sectional plane

rar

outer radius of the collecting channel

riR

inner radius of the collecting channel

rSA

Radius on which the radially outer areas of the discharge openings lie

rGSFmax

largest radius of the outside of the fan blade roots

rGSFmin

smallest radius of the outside of the fan blade roots

raVS

Radius of the outside of the compressor disk

riVS

Radius of the radially inner area of the compressor disk

raLPT

outer radius of the planet carrier bearing

U

Circumferential direction

WLM

horizontal longitudinal median plane

VLM

vertical longitudinal center plane

x

Relationship

X

axial direction

Y

radial direction

α, α', β

angle

Claims (11)

Gas turbine engine (10) for an aircraft with an engine core (11), which has at least one compressor (14, 15) and a turbine (17, 19) and a compressor (14 or 15) with the turbine (17 or 19) connecting core shaft (26, 27) and delimiting a flow cross section (55) for a core air flow (A), a fan (23) being provided upstream of the engine core (55), which has a planetary gear drive (30) is connected to the core shaft (26), the fan (23) having a fan spinner (52) with a fan spinner cone (52A) and fan blades (53) connected to it in the area of fan blade roots (53A), whereby Air flows into the flow cross section (55) of the engine core (11) along an outer region of the fan spinner cone (52A) and from outer sides (53B) of the fan blade roots (53A), the flow cross section (55) of the engine core (11) being radially inward a radial outside (56) of a compressor disk (57) of the compressor (14), the planetary gear (30) being arranged in the axial direction (X) between the fan (23) and the compressor (14), the planetary gear ( 30) comprises a rotatably mounted component (28, 32, 34), which is designed with an at least approximately rotationally symmetrical collecting channel (41), into which oil can be introduced starting from its radially inner region (42) and from which oil is discharged via a discharge opening ( 46) of a flow branch (46A) can be diverted in the direction of at least one consumer of the planetary gear (30), a radially outer region (58) of the discharge opening (46) of the flow branch (46A) being on a circle around an axis of rotation (9) of the rotatable component (28, 32, 34) is arranged, the radius (r _SA ) of the circle being larger than the radius (r _iR ) of the inner region (42) of the collecting channel (41) and smaller than or equal to the radially outer radius (r _aR ) of the collecting trough (41), and wherein the radius (r _SA ) of the circle is smaller than the radius (r _GSFmax or r _GSFmin ) of the outer sides (53B) of the fan blade roots (53A) and / or is smaller than the radius (r _aVS ) of the outside (56) of the compressor disk (57) of the compressor (14). Gas turbine engine Claim 1 , characterized in that a ratio between the radius (r _SA ), on which the radially outer region (58) of the discharge opening (46) lies, and the radius (r _aVS ) of the outside (56) of the compressor disk (57) of the compressor (14) is in a range from 0.7 to 0.9. Gas turbine engine Claim 1 or 2 , characterized in that a ratio between the inner radius (r _iR ) of the collecting trough (41) and the radius (nvs) of a radially inner region (59) of the annular disk-shaped compressor disk (57) of the compressor (14), which has a central opening ( 60) is limited, in a range from 0.8 to 1.2. Gas turbine engine according to at least one of the preceding claims, characterized in that a ratio between the radius (r _SA ) on which the radially outer region (58) of the discharge opening (46) lies and the largest radius (r _GSFmax ) of the outer sides (53B ) of the fan blade roots (53A) is in a range of 0.4 to 0.7. Gas turbine engine according to at least one of the preceding claims, characterized in that a ratio between the radius (r _SA ) on which the radially outer region (58) of the discharge opening (46) lies and the smallest radius (r _GSFmin ) of the outer sides (53B ) of the fan blade roots (53A) is in a range of 0.7 to 0.9. Gas turbine engine according to at least one of the preceding claims, characterized in that the inner radius (r _iR ) of the collecting trough (41) according to the formula relationship $$r_{iR}=\sqrt{r_{aR}^2-x^2\cdot r_{aR}^2}$$

is determined depending on the outer radius (r _aR ) of the collecting channel (41), where x corresponds to a ratio that corresponds to the formula $$x=\sqrt{\frac{r_{aR}^2-r_{iR}^2}{r_{aR}^2}}$$

is determined depending on the inner radius (r _iR ) and the outer radius (r _aR ) of the collecting channel (41). Gas turbine engine Claim 6 , characterized in that the ratio x has values from 0.5 to 0.8, preferably from 0.65 to 0.75. Gas turbine engine according to at least one of the preceding claims, characterized in that the rotatable component (34) is a planet carrier of the planetary gear (30). Gas turbine engine Claim 8 , characterized in that a ratio between the inner radius (r _iR ) of the collecting trough (41) and a radially outer radius (raLPT) of a bearing (61) of the planet carrier (34) is in a range from 1.1 to 1.7 . Gas turbine engine Claim 9 , characterized in that oil can be introduced from an oil supply (44; 44 ') fixed to the housing into the collecting trough (41), which is in the axial direction (X) of the planetary gear next to the collecting trough (41) and in the radial direction (Y) between the inner radius (r _iR ) of the collecting trough (41) and the outer radius (r _aLPT ) of the bearing (60) of the planet carrier (34) is arranged. Gas turbine engine according to at least one of the preceding claims, characterized in that the planetary gear (30) is arranged in the axial direction (X) between the fan (23) and the compressor (14) and radially within the flow cross section (55) of the engine core (11). .

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