Gas turbine compressor stage

Abstract

The present invention relates to a compressor stage for a gas turbine, in particular, an aircraft engine, having a row of rotating blades (3) and a row of guide vanes (4), which is adjacent downstream, wherein the choke point and the aspect ratio AR.sub.ax, which is defined by the quotient between average channel height (h) and average chord length (l.sub.ax), satisfy the condition
>1.33.Math.AR.sub.ax+5.16.

Claims

1. A compressor stage for a gas turbine aircraft engine, having a row of rotating blades and a row of guide vanes, which is adjacent downstream, wherein the choke point and the aspect ratio AR.sub.ax, which is defined by the quotient between average channel height and average chord length, satisfy the condition
>1.33.Math.AR.sub.ax+5.16.

2. The compressor stage according to claim 1, wherein the aspect ratio AR.sub.ax is greater than 0.5 and less than 2.5.

3. The compressor stage according to claim 1, wherein the compressor stage is configured and arranged in a gas turbine having at least one compressor.

4. The compressor stage according to claim 1, wherein a total pressure ratio of at least one of the compressors amounts to at least 40.

5. The compressor stage according to claim 1, wherein the compressor stage is configured and arranged in an aircraft engine having a gas turbine.

6. The compressor stage according to claim 1, wherein a by-pass ratio BPR of the aircraft engine is at least 10.

7. A method for configuring at least one compressor stage of at least one compressor of a gas turbine aircraft engine, having a row of rotating blades and a row of guide vanes, which is adjacent downstream, comprising the step of: aerodynamically configuring the compressor stage so that the choke point and the aspect ratio AR.sub.ax, which is defined by the quotient between average channel height and average chord length, satisfy the condition
>1.33AR.sub.ax+5.16.

8. The method according to claim 7, wherein at least one compressor stage of the compressor is configured.

9. The method according to claim 8, wherein a total pressure ratio of the compressor amounts to at least 40.

10. The method according to claim 7, wherein a by-pass ratio BPR of the aircraft engine amounts to at least 10.

Description

BRIEF DESCRIPTION OF THE DRAWING FIGURES

(1) Additional advantageous enhancements of the present invention can be taken from the dependent claims and the following description of preferred embodiments. For this purpose and partially schematized:

(2) FIG. 1 shows an aircraft engine having a gas turbine with a compressor having several compressor stages according to an embodiment of the present invention; and

(3) FIG. 2 shows a boundary curve for designing the compressor stages according to an embodiment of the present invention.

DESCRIPTION OF THE INVENTION

(4) FIG. 1 shows in partially schematized form an aircraft engine with a fan 1 and a gas turbine, which, simply for a more compact illustration and as an example, has only one compressor 9, a downstream combustion chamber 5, a high-pressure turbine 6, which is coupled with the compressor 9 via a rotor 10, and a low-pressure turbine 7, which is coupled with the fan 1. A core flow 8 flows through the gas turbine and a by-pass flow 2 flows around the gas turbine.

(5) The compressor 9 has several compressor stages, each of which has a row of rotating blades 3 fastened to the rotor and a row of guide vanes 4 adjacent downstream.

(6) One or more of these compressor stages 3, 4 is or are designed such that the choke point and the aspect ratio AR.sub.ax, which is defined by the quotient between average channel height h and average chord length l.sub.ax, satisfy the condition
>1.33.Math.AR.sub.ax+5.16;
the aspect ratio AR.sub.ax is greater than 0.5 and less than 2.5.

(7) The total pressure ratio of the compressor 9 amounts to at least 45; the by-pass ratio BPR of the aircraft engine is at least 12.

(8) FIG. 2 shows a boundary curve for designing the compressor stages according to an embodiment of the present invention. These will be or are designed such that the choke point lies above the boundary curve =1.33.Math.AR.sub.ax+5.16, which is depicted by the bold line in FIG. 2.

(9) Although exemplary embodiments were explained in the preceding description, it shall be noted that a plurality of modifications is possible.

(10) Thus, the aircraft engine or the gas turbine may have, in particular, a low-pressure compressor and a downstream high-pressure compressor; in an enhancement, there is also an intermediate compressor disposed therebetween; whereby at least one of these compressors can be or will be able to be designed in the way explained in the preceding example with reference to compressor 9. Likewise, the low-pressure and high-pressure compressors can also be understood as a compressor in the sense of the present invention.

(11) The fan 1 can be coupled to the high-pressure turbine 6, in particular, via a gearing or drive.

(12) In addition, it shall be noted that the exemplary embodiments only involve examples that in no way shall limit the scope of protection, the applications and the construction. Rather, a guide is given to the person skilled in the art by the preceding description for implementing at least one exemplary embodiment, whereby diverse modifications, particularly with respect to the function and arrangement of the described components, can be carried out without departing from the scope of protection, as it results from the claims and combinations of features equivalent to these.