EP2527743A2 - Segment component comprising high temperature cast material for an annular combustion chamber, annular combustion chamber for an aircraft engine, aircraft engine and method for producing an annular combustion chamber - Google Patents

Abstract

The component has inner and outer combustion chamber walls (11, 12) shielded from an environment when a fuel flame extends along burner axes (21) during operation of the component. The combustion chamber walls comprise projections in a direction away from the burner axes. The fuel flame is arranged between the combustion chamber walls. The projections of the combustion chamber walls are adapted to a contour of the fuel flame during operation of the component. Length and/or width of the projections correspond to length and/or width of the fuel flame. Independent claims are also included for the following: (1) an annular combustion chamber (2) a method for manufacturing an annular combustion chamber.

Description

The invention relates to a segment component of high-temperature casting material for an annular combustion chamber, an annular combustion chamber for an aircraft engine, an aircraft engine and a method for producing an annular combustion chamber.

Modern aircraft engines usually have annular combustion chambers, which are arranged axially between the compressor and the turbine. An annular combustion chamber has coaxially with the engine longitudinal axis a combustion chamber walls bounded annular space, which is also referred to as a flame tube. Along the annular cross section of the annular space, the injectors for the fuel are arranged. In operation, the fuel flames extend from these injectors into the annulus.

Due to the high thermal loads, the combustion chamber walls must be designed to be thermally stable accordingly. So it is known to equip the combustion chamber walls with thermally particularly resilient plates. From the EP 1 106 927 For example, a method is known in which the annulus of an annular combustor is assembled from individual segments of cast material using high temperature cast materials.

It is the object of the present invention to provide segment components for annular combustion chambers, which are improved thermally and fluidically.

This object is achieved by a segment component with the features of claim 1.

In this case, a combustion chamber wall, which during operation shields a fuel flame extending along a burner axis from the environment, has a bulge, the bulge pointing in a direction pointing away from the burner axis. A part of a segmental component for an outer combustion chamber wall of an annular combustion chamber has e.g. a bulge that points radially outward. A portion of a segmental component for an internal combustion chamber wall includes e.g. a bulge that points outward. By the bulges a larger space is created in the immediate area around the burner flame by the distance of the combustion chamber walls is at least partially increased by the burner flame.

It is advantageous if an inner combustion chamber wall and an outer combustion chamber wall, between which in operation a fuel flame is arranged along a burner axis, are used, e.g. U-shaped are arranged. In this case, then the inner and / or the outer combustion chamber wall on a bulge in the direction pointing away from the burner axis.

It is particularly advantageous if the at least one bulge of the combustion chamber wall is adapted substantially to the contour of the fuel flame during operation. In this case, advantageously, the length and / or width of the write-off substantially correspond to the length and / or width of the fuel flame during operation.

Advantageous high-temperature cast materials are a superalloy containing nickel, chromium, cobalt and / or nickel-iron, especially Inconel 738 / Inconel 738 LC, Inconel 939 / Inconel 939 LC, Inconel 713 / Inconel 713 LC, C1023, Mar M 002 and / or CM 274LC. These materials have a sufficient temperature resistance.

In an advantageous embodiment, the inner combustion chamber wall and the outer combustion chamber wall are integrally connected as a casting via a combustion head or the inner combustion chamber wall and outer combustion chamber wall are connected to a combustion chamber head. In the first variant, one-piece segment components are present; in the second variant, there are two segment components which are connected to one another.

In this case, an advantageous embodiment is provided if at least one mounting flange is arranged on the combustion chamber head. Furthermore, it is advantageous if a device for arranging an injector for fuel is provided on the combustion chamber head. Also, advantageously, at least one integrally formed on a combustion chamber wall nozzle for cooling air can be provided.

Advantageously, in one embodiment, the combustion chamber wall has an average thickness of between 1 and 4 mm, in particular 1.4 to 3 mm.

The object is also achieved by an annular combustion chamber for an aircraft engine with the features of claim 9. In this case, at least two segment components according to at least one of claims 1 to 8 are used.

Advantageous embodiments of the annular combustion chamber have a variable annular space height along the circumference of the annular space. By adapting the annulus height to, for example, burner flames and / or injectors, the thermal and / or mechanical loading of the walls can be achieved. This applies in particular when regions A of a larger annular space height H _RA alternate with regions B of a smaller annular space height H _RB along the circumference, so that the combustion chamber walls form a type of wave-like structure

It is particularly advantageous if areas are formed with a larger annulus height and areas with a smaller annulus height, which are arranged in the assembly injectors for the fuel in the areas with the larger annulus height. The areas of larger annulus height give the fuel flame more space and shield it from annoyance in the annulus.

Furthermore, in advantageous embodiments, the segment components are interconnected by welds, in particular by electron beam welding, laser welds with IN626 Filler, Polymet 972 or other ductile welding consumables.

The object is also achieved by an aircraft engine with an annular combustion chamber according to claims 11 to 14. All flow from the compressor, through the combustor to the turbine, is enhanced by the bulges around the flames.

The object is also achieved by methods for producing an annular combustion chamber.

In one embodiment, at least two segmental components having an inner combustion chamber wall, an outer combustion chamber wall and a combustion head are molded from high temperature cast material. Subsequently, the at least two segment components are joined by joining, in particular welding, to the annular combustion chamber.

Alternatively, at least two segment components are connected to form an inner full ring structure, in particular welded. At least two segment components are connected to form an outer full ring structure, in particular welded. The present solid ring structures are connected to a combustion head structure.

Fig. 1

eine schematische, perspektivische Darstellung einer an sich bekannten Ringbrennkammer;

Fig. 2

eine perspektivische Darstellung einer Ausführungsform eines Segmentbauteils mit zwei Brennkammerwänden für eine Ringbrennkammer,

Fig. 2A

eine Ansicht vom Brennkammerkopf auf die Ausführungsform gemäß Fig. 2;

Fig. 2B

eine Schnittansicht der Ausführungsform gemäß Fig. 2 in Längsrichtung;

Fig. 2C

eine Schnittansicht der Ausführungsform gemäß Fig. 2 senkrecht zur Längsrichtung;

Fig. 3

eine axiale Schnittansicht auf eine Ausführungsform für eine Ringbrennkammer gebildet aus Segmentbauteilen gemäß der Ausführungsform nach Fig. 2;

Fig. 4

eine Draufsicht auf eine weitere Ausführungsform eines Segmentbauteils mit zwei Brennkammerwänden,

Fig.5

eine weitere Ausführungsform eines Segmentbauteils mit einer Brennkammerwand;

Fig. 6A

eine perspektivische Ansicht einer ersten Stufe des Aufbaus einer Ringraumstruktur;

Fig. 6B

eine perspektivische Ansicht einer zweiten Stufe des Aufbaus einer Ringraumstruktur.

The invention will be explained in more detail with reference to the figures of the drawing with reference to several embodiments. Show it:

Fig. 1

a schematic, perspective view of a known annular combustion chamber;

Fig. 2

3 is a perspective view of an embodiment of a segment component with two combustion chamber walls for an annular combustion chamber,

Fig. 2A

a view from the combustion chamber head to the embodiment according to Fig. 2 ;

Fig. 2B

a sectional view of the embodiment according to Fig. 2 longitudinal;

Fig. 2C

a sectional view of the embodiment according to Fig. 2 perpendicular to the longitudinal direction;

Fig. 3

an axial sectional view of an embodiment for an annular combustion chamber formed from segment components according to the embodiment of FIG Fig. 2 ;

Fig. 4

a top view of a further embodiment of a segment component with two combustion chamber walls,

Figure 5

a further embodiment of a segment component with a combustion chamber wall;

Fig. 6A

a perspective view of a first stage of the construction of an annulus structure;

Fig. 6B

a perspective view of a second stage of the structure of an annulus structure.

In Fig. 1 In a perspective view, an annular combustion chamber is shown with an annular space 30, as used for example in an aircraft engine.

The annular space 30 is arranged in the main flow direction of the aircraft engine behind the compressor, not shown here, and the inlet region of a turbine 40. In the presentation of the Fig. 1 two injectors 25 are visible, from which fuel flames 20 (not shown here) emerge along burner axes 21 during operation. The burner axes 21 and thus also the fuel flames 20 are thus between the inner combustion chamber wall 11 and the outer combustion chamber wall 12. This annular space 30 is also referred to as a flame tube. The combustion chamber walls 11, 12 thus shield the fuel flames 20 inwardly and outwardly from the environment.

Although the distance between the combustion chamber walls 11, 12, the annulus height H _R (also referred to as Flammraumhöhe) varies in the axial direction of the aircraft engine, but is constant along the circumference of the annular combustion chamber 10.

The invention described below with reference to various embodiments relates, inter alia, annular combustion chambers in which the ring combustion chamber height H _{R is} non-constant along the circumference.

Such an annular combustion chamber is e.g. composed of at least two segment components 10 made of high temperature casting material. In the case of two segmental components, each of the segmental components 10 would be e.g. 180 ° of the annulus 30 provide.

In Fig. 2 a segment component 10 is shown, which covers a much smaller angular range, namely 30 °, as in the view of Fig. 2A is particularly clear.

An annular combustion chamber, composed of such segmental components 10, therefore has 12 of these segmental components 10. In principle, it is possible to make the segment components 10 geometrically different, so that fewer or more than 12 segment components 10 are used. Also, it is not mandatory that an even number of segment components 10 be used to form an annulus 30.

In Fig. 2 For example, one embodiment of a segmental component 10 is shown in which parts form the inner combustion chamber wall 11 and the outer combustion chamber wall 12 when the segmental components 10 are assembled (see FIG Fig. 5 ). At the combustion chamber head 22, an opening 24 is provided for the injector 25, not shown here. The fuel flame 20 (not shown here) produced by the injector 25 extends along the burner axis 21 into the annular space 30 in the direction of the inlet region of the turbine 40 (not shown here, see FIG Fig. 1 ).

This embodiment of the segment component 10 is produced in one piece from a high-temperature casting material. For this purpose, advantageously, a superalloy may be used which contains nickel, chromium, cobalt and / or nickel-iron. Typical high temperature cast alloys are in particular Inconel 738 / Inconel 738 LC, Inconel 939 / Inconel 939 LC, Inconel 713 / Inconel 713 LC, C1023, Mar M 002 and / or CM 274LC. The casting processes (e.g., investment casting) allow segmental components 10 to be made with very thin walls and in very complex shapes.

For example, it is advantageous if the combustion chamber walls 11, 12 have an average thickness of between 1 and 4 mm. The wall of the combustion chamber head 23 can between 2 and 4 mm. In the shaping, it is possible, for example, to form nozzle 15 for air cooling during casting. Also, mounting flanges 23 can be molded integrally on the combustion chamber head 22. Basically, the possibilities of shaping are not limited to the illustrated features.

The combustion chamber walls 11, 12 of this embodiment are contoured in a particular way. The inner combustion chamber wall 11 has a bulge 13, which points downwards in the representation selected here. The bulge 13 thus points away from the burner axis 21. The outer combustion chamber wall 12 has a bulge 14, which has the same shape in a slightly upward direction. The bulge 14 thus also points away from the burner axis 21.

The bulges 13, 14 are arranged so that they correspond approximately to the contour of the fuel flame 20 when the annular combustion chamber is in operation.

These relationships are schematic in the Fig. 2B , C shown, where Fig. 2B a longitudinal section through the annular space 30 shown; Fig. 2C shows a sectional view perpendicular thereto. In the sectional view of Fig. 2B schematically the fuel flame 20 is shown, which extends from the injector 25 into the annular space 30 over a length L _B. The length of the entire annulus is denoted by L. It is advantageous if for the length L _{B of} the fuel flame 20: L _B = 0.5 - 0.9 L. This means that the fuel flame 20 extends over 50 to 90% of the axial extent of the annular space.

The bulge 13 on the inner combustion chamber wall 11 and the bulge 14 on the outer combustion chamber wall 12 extend in the axial direction approximately as far as the fuel flame 20 extends into the annular space.

In advantageous embodiments, the axial extension of the bulges, 13, 14 is about 50 to 90% of the total axial extent of the annular space. Further, it is advantageous if the width B _{B of} the bookings 13, 14 is about 30 to 60% of the width B of a segment component 10, wherein the width B of the bulge is smaller on the inside than on the outside.

In Fig. 2C is the one to view the Fig. 2B also shown that the bulges 13, 14 are approximately adapted to the contour of the fuel flame.

In Fig. 2C a region A is drawn in, in which the annulus height H _{RA is increased} by the stakes 13, 14 and a region B, in which the annulus height H _{RB is} reduced.

An arc length U of the segment component 10 is thus composed of A + 2B. It is advantageous if the proportion of the range A is 50 to 80% of the arc length U and the portion of the range B is 20 to 50% of the arc length U.

Furthermore, in Fig. 2C the usual radii of the combustion chamber walls, namely R _i and R _a , wherein it can be seen that protrusions 13, 14 are partially outside of R _a or within R _i . The usual annulus _height H _konv thus corresponds to R _a - R _i .

Advantageous embodiments have bulges 13, 14, for which applies: H _RA = 1.1 - 1.5 H _conv . This means that the height of the combustion chamber in the region of the bulges 13, 14 is extended between 10 and 50% compared to the conventional design.

It is also advantageous if, in region B, ie regions without outbreaks 13, 14: H _RA = 0.7 - 0.9 H _conv . This means that the height of the combustion chamber in the area outside the bulges 13, 14 is 70 to 90% of the usual height.

If now several of these segment components 10 are connected to each other, an annular combustion chamber is formed whose annular space height H _{R is} variable in the circumferential direction. Segment components 10 are connected to each other, for example by laser or electron beam welding, whereby the introduced path energy is minimized. It can be a suitable, ductile filler used in welding (IN625 or Polymet 972).

Such a compound rivet combustion chamber is in Fig. 3 shown. For reasons of clarity, only six segment components 10 are used here to form an annular space 30. Areas A of a larger annular space height H _RA alternate with areas B of a smaller annular space height H _RB along the circumference, so that the combustion chamber walls 11, 12 form a kind of wave-like structure.

The fuel flames 20 (not shown here) are in each case in the extended areas A. Between the fuel flames 20 are narrowed areas B. This leads to the effect that each fuel flame 20 can effectively burn in its own combustion chamber. Disturbances in a region of the annular space 30 can spread more severely in the entire annular space 30 due to the constrictions in the regions B.

Also, in areas B between the injectors 25, air may be directed from the compressor to the turbine 40 with less severe deflection, thereby reducing the pressure loss on this flow path.

However, the described embodiment also has advantageous effects outside of the annular space 30, since the turbine cooling air K, which is guided outside the annular space, is also influenced by the contouring of the combustion chamber walls 11, 12.

In this case, the pressure loss during the transfer of the turbine cooling air K from the compressor outlet to the combustion chamber is determined to enter the cooling system by the flow guidance in this way. If the turbine cooling air K has to be deflected repeatedly (in particular radially) and accelerated (and then decelerated again), then the pressure loss increases. In the burner axis 21, only a small amount of turbine cooling air K flows past the burner and mixing air hole in the direction of the turbine, so the pressure loss is not so decisive there.

Between the burners in the present embodiment, the combustion chamber head 22 is designed so that the turbine cooling air K is not greatly deflected radially outwards and inwards. These are the areas B between the bulges 13, 14, but at the respective outer sides of the annular space 30. After the radial deflection then takes place a deflection in the axial direction. Thus, in area B, there is a small deflection into the much deeper annuli around the narrower combustion chamber at this point. The flow of turbine cooling air K is in Fig. 3 shown schematically.

With appropriate flow control so less pressure losses. The pressure loss is reduced by the indentation between the burners. As a result of the lower annuli, the turbine cooling air K also has less contact with the hot combustion chamber wall compared to the usual gap flow and is thus delivered colder to the turbine, which benefits the cooling effect in the turbine.

Overall, the total pressure loss can be reduced, which lowers the fuel requirement. In addition, less air flows between the injectors 25 into the region of the combustion chamber head 22 than at the position of the injectors 25, so that at this Circumferential positions sufficient air for transfer to the turbine 40 is available.

Furthermore, the bulges 13, 14 cause a more uniform temperature distribution in the circumferential direction to form in the combustion chamber walls 11, 12, which has a positive influence on the service life of the annular combustion chamber. In the areas A, in which the fuel flame 20 is located, the combustion chamber wall 11, 12, due to the bulges 13, 14 relatively far away from the fuel flame 20. In the areas B, between the fuel flames 20, the combustion chamber walls 11, 12 are closer together, since the annulus height H _R is lower here. Without the bulges 13, 14, the wall regions of the combustion chamber walls 11, 12 which are closest to the fuel flame 20 would be hotter than other regions. For these reasons, not so much cooling air needs to be used in area A. The cooling air thus saved is available for measures to reduce exhaust emissions.

As in Fig. 3 As can be seen, the inner combustion chamber wall 11 and the outer combustion chamber wall 12 have a wavy structure when made of segmental components 10, for example according to FIG Fig. 2 are composed. This wavy structure allows for easier compensation of thermal and / or mechanical stresses in the combustion chamber walls 11, 12 than would be possible in annular spaces with circular cross-sections in the circumferential direction.

If deemed necessary (e.g., in larger aircraft engines), the segmental components 10 may be provided with a thermal barrier coating.

When using a ductile weld filler, in the case of subsequent laser drilling at the annular combustion chamber, the positions of the longitudinal welds between the segmental components 10 need not be considered.

In Fig. 4 a further embodiment of a segment component 10 is shown. Basically, it has the same functions and properties as the segment component 10 described above, so that reference can be made to the corresponding description.

In contrast to the substantially rectangular bulges 13, 14 in the embodiment according to FIG Fig. 2 Here are the bulges 14 in the form of

Fuel flame 20 from the combustion chamber head 23 in the direction of the turbine 40 (not shown here) arranged. The bulge 13 has a rather small width in the vicinity of the combustion chamber head 23, which widens steadily, in order then to become smaller again.

In principle, other shapes for bulges are possible with the casting process, which can be adapted to a particular application. Just by using the above-mentioned materials and the casting process, it is possible to form the bulges 13, 14 targeted.

In the Fig. 2 . 3 and 4 Embodiments were shown in which two combustion chamber walls 11, 12 are opposed. These segmental components 10 thus have a substantially U-shaped form, since the combustion chamber walls 11, 12 are connected via the combustion chamber head 23 cast integrally therewith.

In principle, however, it is also possible for a segment component 10 to have only one outer or inner part of the annular combustion chamber. In Fig. 5 an embodiment of a segment component 10 is shown, which has only one outer combustion chamber wall 12. Like the previously described embodiments, this segmental component 10 also has a bulge 14 that faces away from the burner axis 21. To illustrate the use of this segmental component 10, are in Fig. 5 dashed lines the fuel flame 20 and the burner axis 21 drawn.

Even with this embodiment and a corresponding segment component 10 for the inner combustion chamber wall 11, an annular combustion chamber can be constructed, as shown in FIG Fig. 6A , B is shown.

For this purpose, at least two segment components 10 'are connected to an inner full ring structure 31, in particular welded. Furthermore, two segment components 10 "are connected to an outer full ring structure 32, in particular welded Fig. 6A are the two full ring structures 31, 32 shown, each having only six segment components 10 for reasons of simplicity. Subsequently, the inner full ring structure 31 and the outer full ring structure 32 are connected to a combustion head structure 43, as shown in FIG Fig. 6B is shown.

LIST OF REFERENCE NUMBERS

10

segment component

11

inner combustion chamber wall

12

outer combustion chamber wall

13

Bulge of inner combustion chamber wall

14

Bulge outer combustion chamber wall

15

Nozzle for cooling air

20

fuel flame

21

Brenner

22

bulkhead

23

mounting flange

24

Device for arranging a burner

25

Injector for fuel

30

annulus

31

inner full ring structure

32

outer full ring structure

40

Inlet area turbine

K

Turbine cooling air

H _RA

Range of larger annulus height

H _RB

Area of smaller annulus height

H _R

Annulus height

H _conv

usual annulus height

R _i

Radius of the inner combustion chamber wall

R _a

Radius of the outer combustion chamber wall

B

Wide segmental component

B _B

Wide bulge

L _B

Length of fuel flame

L

Length of fuel chamber

U

Arc length of a segment component

Claims (15)

High temperature casting material segment component for an annular combustion chamber of an aircraft engine,
marked by
a combustion chamber wall (11, 12) which in operation shields a fuel flame (20) extending along a burner axis (21) from the environment, the combustion chamber wall (11, 12) having a bulge (13, 14) in one direction away from the burner axis (21). Segment component according to claim 1, characterized by an inner combustion chamber wall (11) and an outer combustion chamber wall (12), between which a fuel flame (20) is arranged in operation along a burner axis (21), the inner and / or outer combustion chamber wall (11, 12) has a bulge (13, 14) in the direction pointing away from the burner axis (21). Segment component according to claim 1 or 2, characterized in that the at least one bulge (13, 14) of the combustion chamber wall (11, 12) is substantially adapted to the contour of the fuel flame (20) during operation, in particular that the combustion chamber wall (11, 12 ) has a bulge (13, 14) whose length (L _B ) and / or width (B _B ) corresponds substantially to the length and / or width of the fuel flame (20) during operation. Segment component according to at least one of the preceding claims, characterized in that the high-temperature casting material is a superalloy containing nickel, chromium, cobalt and / or nickel-iron, Inconel 738 / Inconel 738 LC, Inconel 939 / Inconel 939 LC, Inconel 713 / Inconel 713 LC , C1023, Mar M 002 and / or CM 274LC. Segment component according to at least one of the preceding claims, characterized in that the inner combustion chamber wall (11) and the outer combustion chamber wall (12) are integrally connected as a casting via a combustion chamber head (12) or the inner Combustion chamber wall (11) and outer combustion chamber wall (12) with a combustion chamber head (13) are connected. Segment component according to claim 5, characterized in that at least one mounting flange (23) is arranged on the combustion chamber head (22) and / or a device (24) for arranging an injector (25) is provided on the combustion chamber head (13). Segment component according to at least one of the preceding claims, characterized by at least one connecting piece (15) integrally formed on a combustion chamber wall (11, 12) for cooling air. Segment component according to at least one of the preceding claims, characterized in that the combustion chamber wall (11, 12) has an average thickness between 1 and 4 mm, in particular 1.4 to 3 mm. Annular combustion chamber for an aircraft engine with at least two segment components (10) according to at least one of claims 1 to 8. Annular combustion chamber according to Claim 9, characterized by h a variable annulus height (H _R ) along the circumference of the annulus (30), in particular regions (A) of greater annulus height (H _RA ) coinciding with regions (B) of lower annulus height (H _RB ) along the circumference, so that the combustion chamber walls (11, 12) form a kind of wave-like structure Annular combustion chamber according to claim 9 or 10, characterized by regions with a larger annulus height (H _RA ) and regions with a smaller annulus height (H _RB ), wherein in the assembly injectors (25) for the fuel in the areas with the larger annulus height (H _RA ) are arranged. Annular combustion chamber according to at least one of claims 9 to 11, characterized in that the segment components (10) are interconnected by welds, in particular by electron beam welding, laser welding seams with IN626 Filler, Polymet 972 or other ductile welding consumables. An aircraft engine with an annular combustion chamber according to claims 9 to 12. Method for producing an annular combustion chamber according to at least one of claims 9 to 12,
characterized in that a) at least two segmental components (10) are cast with an inner combustion chamber wall (11), an outer combustion chamber wall (12) and a combustion chamber head (13) of high temperature casting material, and then the b) at least two segmental components (10) by joining, in particular welding to the annular combustion chamber (10) are connected. Method for producing an annular combustion chamber according to at least one of claims 9 to 12,
characterized in that a) at least two segment components (10 ') connected to an inner full ring structure (31), in particular welded, b) at least two segment components (10 ") connected to an outer full ring structure (32), in particular welded, c) the inner full ring structure (31) and the outer full ring structure (32) are connected to a combustion chamber head structure (43).

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