DE7629587U1 - Gas turbine engine, in particular small gas turbine engine - Google Patents

Description

Munich, September 16, 1976

Gas turbine engine, in particular small gas turbine engine

The invention relates to a (gas turbine engine, in particular a small gas turbine engine, which has a radial compressor or a combined axial-radial compressor, a reversing ring combustion chamber, a compressor drive turbine and, if necessary, sin © of; the latter inech ~ sni5t; h

Has utility turbo, the gas generator rotor via a Fixed bearing and at least one other floating bearing is supported.

Smaller gas turbine engines, especially gas turbine engines! works with centrifugal compressors or combined axial-radial ί Compressors are usually designed so that the FeSt- |

I bearing of the gas generator rotor is arranged in the area of the compressor inlet. This design brings advantages for the Design of the rotor bearing in the hot area also clearly has disadvantages for efficiency, operating behavior and safety the machine. These disadvantages are particularly evident if the engine is without an entry bearing star (e.g. with a radial or side air inlet).

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The cause of these disadvantages is the thermal expansions of the rotor and stator, which vary in direction, size and time, and which have an effect in particular on the efficiency and the pumping behavior of the centrifugal compressor. The gaps between the impeller and the cover plate have the same negative influence as the axial offsets between the impeller and the radial part of the diffuser. Furthermore, the efficiency and the operating behavior of the axial part of the compressor, possibly also the supply of cooling air to the turbine blades to be cooled, is adversely affected because in the former case, depending on the design of the compressor, the radial clearances between the blades and the housing or between the guide blades and the rotor need to be enlarged unnecessarily, in the second case the labyrinths Ίτη area, the cool 1 lu-f i ^ f iiiur iina to there tnrbinenlaufSehaufslH impaired in their function, which easily leads to overheating of the blading.

The invention is based on the object of avoiding the disadvantages mentioned and a gas turbine engine as well Known to create mainly cheaper storage.

According to the invention, this is done in order to solve the problem posed proposed to arrange the fixed bearing of the gas generator rotor behind the radial compressor and in front of the compressor turbine, whereby the bearing star required in this case as against axial forces

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rigid card construction is formed. Here, the radial and the axial diffuser of the radial compressor should be two separate Be assigned to components that are arranged with one another by at least two on different diameters Flanges are bolted. The front component contains the radial flow section of the radial compressor diffuser and the bend, which redirects the flow 90 into the axial. The rear component, expediently designed as a cast part, contains the section of the radial compressor diffuser through which there is axial flow as well as the complete storage chamber of the fixed bearing with all associated seals and supply lines.

In a further embodiment of the invention, the axial part of the

trad you stock cannon through a roughly

double-conical shell connected, which is also the combustion chamber area separated from the compressor area. All supply lines for the camp can be used as hollow ribs be arranged on the compressor side of the double-conical shell and as additional stiffening against axial deformations this component act.

Overall, the solution according to the invention offers the following advantages:

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- Negligibly small axial deformations between the hot bearing and the machine housing, even with large ones Pressure differences between the combustion chamber and compressor side and / or in the case of high axial forces from the rotor bearing

- Negligible operational axial misalignments between centrifugal compressor impeller and diffuser

- Negligible operational axial misalignments between the centrifugal compressor impeller and its cover disk

- The compressor rotor and stator "grow" with thermal expansion in the same direction

- Negligible differential axial expansions in the area of Labyrinth seals on the front of the turbine rotor

- Small operational axial differential expansions between Rotor and stator of the compressor turbine.

In addition, the flanges of the two shells of the bearing star can be replaced by soldered or welded connections, which, if necessary, waiving the interchangeability of the radially flowed-through part of the guide apparatus of the radial compressor can lead to lower production costs.

The invention is explained further with reference to the drawings, for example; show it

1 shows a gas turbine engine shown schematically along the upper half as a central longitudinal section and

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FIG. 2 as a central longitudinal section compared to FIG. 1 enlarged details of the gas turbine engine, including the radial compressor, the fixed bearing, the compressor drive turbine as well as illustrating part of the annular combustion chamber.

The gas turbine engine according to FIG. 1 has, in sequence, essentially a multi-stage, with adjustable guide vanes provided axial compressor 1, a radial compressor 2, a Reversing annular combustion chamber 3, a compressor drive turbine 4 and a power turbine 5 that is mechanically independent of this.

The gas device rotor 6 is on the compressor side via a floating bearing and mounted behind the radial compressor 2 and in front of the compressor drive turbine 4 via a fixed bearing 8. The one for that Fixed bearing required bearing star is designed as a rigid box structure against axial forces.

This box construction (Fig. 2) consists of a first Component 9, which essentially comprises the radially flowed through section 10 of a radial compressor diffuser and a die Flow through about 90 ° in the axial deflecting bend 11 comprises and over at least two on different diameters arranged flanges 12, 13, 14, 15 with a second component

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connected is. The second component 16 comprises essentially the axially flowed through section 17 of the radial compressor diffuser adjoining the manifold, as well as the complete one Eagle chamber 18 of the fixed bearing 8 with all stationary sealing elements 19, 20, 21, 22 and supply lines 23, 24.

As can also be seen from FIG. 2, the second component 16 is between the axially flowed through section 17 of the radial compressor diffuser and the bearing chamber 18 of the fixed bearing 8 as essentially that of the contour of the combustion chamber housing the following double-conical shell 26, bulged towards the centrifugal compressor 2, is formed, which on the the centrifugal compressor 2 facing side with serving for the storage supply, here e.g. the supply line 24 of the fixed bearing forming hollow ribs 27 is provided as a simultaneous means for stiffening against axial deformation of the second component 16.

The first component 9 also has a stiffening section 28 together with the cover plate 29, the radial compressor 2 is held on a support strut 3o. With the Section 31 is the first component in the engine housing structure integrated.

At 32 there is also an external supply line for the fixed bearing 8 denotes, which e.g. for the oil supply with one or more supply lines formed by the hollow ribs, in the drawing as e.g. via the supply line 24

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the fixed bearing 8 is in connection.

The gas turbine engine according to the invention is for example for driving motor vehicles, tracked vehicles, such as tanks as well suitable as for the propulsion of airplanes, helicopters or for stationary use.

T-510

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Claims (3)

claims 1. Gas turbine engine, in particular small gas turbine engine, which is a radial compressor or a combined axial-radial compressor, a reversing ring combustion chamber, a compressor drive turbine and. possibly one that is mechanically independent of the latter Has power turbine, characterized in that the fixed bearing (8) of the gas generator rotor (6) behind the centrifugal compressor (2) and in front of the compressor drive turbine (4) and the required bearing star as a box structure that is rigid against axial forces is trained. 2. Gas turbine drive iebv / Erk according to claim 1, characterized in that that a first component (9) essentially the section (10) of a radial compressor diffuser through which the flow flows radially and one the flow by about 90 ° T-510 76295S7 05/11/78 ■ · _ yi in the axial deflecting manifold (11) includes and over at least two flanges (12, 14) arranged on different diameters with a second component (16) is connected, which essentially the adjacent to the manifold adjoining the axially flowed through section (17) of the Radial compressor diffuser as well as the complete storage chamber (18) of the fixed bearing (8) with all seals and supply lines (23, 24). 3. Gas turbine engine according to claim 1 and 2, characterized in that that the second component (16) between the axially flowed through section (17) of the radial compressor diffuser and the bearing chamber (18) of the fixed bearing (8) as essentially following the combustion chamber housing contour, against the centrifugal compressor (2) arched double-conical shell (26) is formed, which on the the radial compressor (2) facing side with serving for storage supply hollow ribs (27) as a simultaneous means for Reinforcement against axial deformation of the second component (16) is provided. Τ-5ΊΟ
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