EP3144540B1 - Gas turbine compressor stage - Google Patents

Description

The present invention relates to a compressor stage for a gas turbine, a gas turbine with at least one such compressor stage, an aircraft engine with such a gas turbine and a method for designing such a compressor stage and a method for designing a compressor of such a gas turbine, in particular an aircraft engine.

So far, compressor stages of gas turbines have been designed in such a way that their throttling factor σ is always less than 5.16 minus 1.33 times the aspect ratio AR _ax defined by the quotient of mean duct height h and mean chord length _lax (σ ≤ -1.33 _ARax + 5.16).

H. GRIEB: "Project planning for turbo aircraft engines" shows the state of the art.

In particular, the desire to reduce fuel consumption is increasingly leading to geometrically small compressors with high efficiency and high aerodynamic and mechanical loads with a small overall length.

An object of an embodiment of the present invention is to improve a gas turbine.

This object is achieved by a compressor stage having the features of claim 1 and a method having the features of claim 7 . Claims 3, 5 and 8 protect a gas turbine with a compressor stage described here, an aircraft engine with a gas turbine described here or a method for designing a compressor of a gas turbine described here, in particular an aircraft engine gas turbine. Advantageous embodiments of the invention are the subject matter of the dependent claims.

genügt.According to one aspect of the present invention, one or more compressor stages of a compressor or one or more compressor stages of several compressors of a gas turbine, in particular an aircraft engine gas turbine, which (each) have a rotor cascade and a guide cascade, are aerodynamically designed such that the throttling factor σ and the aspect ratio AR _ax (in each _case ) of the condition defined by the quotient of mean channel height h and mean chord length lax $$\mathrm{\sigma }>-1.33\cdot {\mathrm{AR}}_{\mathrm{ax}}+\mathrm{5:16}$$

enough.

mit der Drosselziffer σ und dem durch den Quotienten aus mittlerer Kanalhöhe h und mittlerer Sehnenlänge l_ax definierte Seitenverhältnis AR_ax.According to one aspect of the present invention, one or more compressor stages for a compressor or one or more compressor stages for several compressors of a gas turbine, in particular an aircraft engine gas turbine, in particular one or more compressor stages of a compressor or one or more compressor stages of several compressors of a gas turbine, in particular suffice an aero engine gas turbine, each having a rotor blade and a vane blade, (each) of the condition $$\mathrm{\sigma }>-1.33\cdot {\mathrm{AR}}_{\mathrm{ax}}+\mathrm{5:16}$$

with the throttle factor σ and the aspect ratio AR _ax defined by the quotient of the mean channel height h and the mean chord length _lax .

It has surprisingly been found that such compressor stages or compressor stages designed in this way, in contrast to previously known compressor stages or designs with the same aerodynamic load and number of stages, reduce a compressor overall length and weight or, with the same overall length, increase the efficiency of the compressor and thus the specific fuel consumption can be reduced.

In one embodiment, a rotor cascade has a plurality of rotor blades spaced apart in the circumferential direction, which are arranged on a rotor which is rotatable (bearing) about a main or machine axis, in particular by a turbine of the gas turbine. The blades can be detachably or integrally attached to the rotor or formed integrally with it. In one embodiment, they can be without a shroud or have a closed outer shroud.

In one embodiment, a guide vane has a plurality of guide vanes which are spaced apart in the circumferential direction and are arranged in a fixed or adjustable manner on a housing which surrounds the rotor. In one embodiment, they can be without a shroud or have a closed inner shroud.

In one embodiment, the guide vane is arranged adjacent to a guide vane downstream or to the moving vane downstream. In particular, it can be a so-called guide vane for converting kinetic energy generated by the rotating rotor cascade into pressure energy from the air flowing through the gas turbine.

In one embodiment, the compression stage in the sense of the present invention consists of the moving cascade and the guide cascade.

The throttle factor σ is defined in the usual way as the quotient of the pressure factor Ψ _(eff) divided by the square of the delivery number ϕ: $$\mathrm{\sigma }={\mathrm{\psi }}_{\left(\mathrm{eff}\right)}/{\mathrm{<figure-callout\; id="2"\; label="\phi "\; filenames="imgf0001.png"\; state="\{\{state\}\}">\phi </figure-callout>}}^2.$$

The pressure coefficient Ψ _(eff) is defined in the usual way as the quotient of twice the (specific) work H _(eff) of the step or the moving grid divided by the square of the peripheral speed at the step or moving grid inlet u ₁ : $${\mathrm{\psi }}_{\left(\mathrm{eff}\right)}=2\cdot {\mathrm{H}}_{\left(\mathrm{eff}\right)}/{{\mathrm{and}}_1}^2.$$

The delivery number ϕ is defined in the usual way as the quotient of the axial absolute speed c _ax , in particular at the step or grate entry (c _ax , ₁ ), divided by the peripheral speed at the step or grate entry u ₁ : $$\mathrm{\phi }={\mathrm{c}}_{\mathrm{ax}\left(,1\right)}/{\mathrm{and}}_1.$$

The throttling factor σ is thus equally defined as the quotient of twice the (specific) work H _(eff) of the stage or the moving gate divided by the Square of the axial absolute velocity c _ax , especially at the step or moving gate entrance (c _ax , ₁ ): $$\mathrm{\sigma }=2\cdot {\mathrm{H}}_{\left(\mathrm{eff}\right)}/{{\mathrm{c}}_{\mathrm{ax}\left(,1\right)}}^2.$$

The average duct height h is defined in the usual way as the geometric mean of half the difference(s) of the outer and inner diameters D _a , D _i of the flow duct of the compressor stage or the rotor cascade: $$\mathrm{H}=\left({\mathrm{D}}_{\mathrm{a}}-{\mathrm{D}}_{\mathrm{i}}\right)/2.$$

The mean chord length _lax is defined in the usual way as the geometric mean of the distance between the inlet and outlet edges of the rotor cascade or the compressor stage.

Accordingly, the aspect ratio AR _ax results in: $${\mathrm{AR}}_{\mathrm{ax}}=\mathrm{H}/{\mathrm{l}}_{\mathrm{ax}}.$$

According to one embodiment, the aspect ratio AR _ax is greater than 0.5. Additionally or alternatively, according to one embodiment, the aspect ratio AR _ax is less than 2.5. As a result, a particularly advantageous compressor stage can be made available.

According to one embodiment, a total pressure ratio Π of one or more of the compressors is at least 40, in particular at least 45. A particularly advantageous compressor can thereby be made available.

The total pressure ratio Π is defined in the usual way as the quotient of the pressure p ₂ at the compressor outlet and the pressure p ₁ at the compressor inlet: $${\displaystyle \prod ={\mathrm{<figure-callout\; id="2"\; label="p"\; filenames="imgf0001.png"\; state="\{\{state\}\}">p</figure-callout>}}_2/{\mathrm{<figure-callout\; id="1"\; label="p"\; filenames="imgf0001.png"\; state="\{\{state\}\}">p</figure-callout>}}_1}.$$

According to one embodiment, a bypass ratio BPR (by-pass ratio) of the aircraft engine is at least 10, in particular at least 12. A particularly advantageous aircraft engine can thereby be made available.

The bypass ratio BPR is defined in the usual way as the quotient of the air mass flow m _coat , which is guided past the gas turbine of the aircraft engine after a fan on the outside (bypass flow or bypass flow), divided by the air mass flow m _core , which passes inside the combustion chamber of the gas turbine and the shaft power provides (core current): $$\mathrm{BPR}={\mathrm{m}}_{\mathrm{a\; coat}}/{\mathrm{m}}_{\mathrm{core}}.$$

In particular, reference is also made to the above variables, which are defined in a manner customary in the art and are therefore known to the person skilled in the art H. Grieb: "Compressors for turbo aircraft engines", Springer-Verlag, ISBN 978-3-540-34373-8 .

Fig. 1

ein Flugtriebwerk mit einer Gasturbine mit einem Verdichter mit mehreren Verdichterstufen nach einer Ausführung der vorliegenden Erfindung; und

Fig. 2

eine Grenzkurve zur Auslegung der Verdichterstufen nach einer Ausführung der vorliegenden Erfindung.

Further advantageous developments of the present invention result from the dependent claims and the following description of preferred embodiments. This shows, partially schematized:

1

an aircraft engine including a gas turbine engine having a multistage compressor according to an embodiment of the present invention; and

2

a limit curve for the design of the compressor stages according to an embodiment of the present invention.

1 shows in a partially schematic manner an aircraft engine with a fan 1 and a gas turbine, which is only for a more compact representation and by way of example only a compressor 9, a downstream combustion chamber 5, a high-pressure turbine 6, which is coupled to the compressor 9 via a rotor 10, and a Has low-pressure turbine 7, which is coupled to the fan 1. A core flow 8 flows through the gas turbine and a bypass or bypass flow 2 flows around it.

The compressor 9 has a plurality of compressor stages, each of which has a fixed-rotor rotor cascade 3 and a guide cascade 4 adjacent downstream.

genügt, das Seitenverhältnis AR_ax größer als 0,5 und kleiner als 2,5 ist.One or more of these compressor stages 3, 4 are or are designed in such a way that the throttle coefficient σ and the aspect ratio AR _ax defined by the quotient of the average channel height h and the average chord length 1 _ax of the condition $$\mathrm{\sigma }>-1.33\cdot {\mathrm{AR}}_{\mathrm{ax}}+\mathrm{5:16}$$

it is sufficient if the aspect ratio AR _ax is greater than 0.5 and less than 2.5.

The total pressure ratio Π of the compressor 9 is at least 45, the bypass ratio BPR of the aircraft engine is at least 12.

2 Figure 12 shows a limit curve for designing the compressor stages according to an embodiment of the present invention. These are or are designed in such a way that the throttle factor σ is above the in 2 bold limit curve σ = -1.33 AR _ax + 5.16.

Although exemplary embodiments have been explained in the preceding description, it should be pointed out that a large number of modifications are possible.

The aircraft engine or the gas turbine can have, in particular, a low-pressure compressor and a downstream high-pressure compressor, and in a further development also a medium-pressure compressor arranged between them, with at least one of these compressors being designed in the manner explained above as an example with reference to the compressor 9 can. Equally, low and high pressure compressors can also be understood as a compressor within the meaning of the present invention.

The fan 1 can be coupled to the high-pressure turbine 6 in particular via a gearbox.

In addition, it should be noted that the exemplary implementations are only examples and are not intended to limit the scope, applications, or construction in any way. Rather, the above description gives the person skilled in the art a guideline for the implementation of at least one exemplary embodiment, with various changes, in particular with regard to the function and arrangement of the components described, being able to be made without departing from the scope of protection as it emerges from the claims.

Reference List

1

fan

2

sheath flow

3

playpen

4

(Post)guide grid

5

combustion chamber

6

high pressure turbine

7

low pressure turbine

8th

core stream

9

compressor

10

rotor

ARax

aspect ratio

H

medium channel height

lax

mean chord length

σ

throttle digit

Claims (10)

1. Compressor stage for a gas turbine, in particular of an aircraft engine, comprising a rotor cascade (3), a stator cascade (4), in particular adjacent thereto downstream, and a throttle value (σ), characterized in that the throttle value (σ) and the aspect ratio AR_ax defined by the quotient of the average channel height (h) of the compressor stage or of the rotor cascade (3) and the average chord length (l_ax) of the compressor stage or of the rotor cascade (3) satisfy the condition $$\mathrm{\sigma }>-1.33\cdot {\mathrm{AR}}_{\mathrm{ax}}+5.16$$

   
2. Compressor stage according to the preceding claim,
   characterized in that the aspect ratio AR_ax is greater than 0.5 and/or less than 2.5.
3. Gas turbine comprising at least one compressor (9) having at least one compressor stage according to either of the preceding claims.
4. Gas turbine according to the preceding claim, characterized in that a total pressure ratio Π of at least one of the compressors is at least 40, in particular at least 45.
5. Aircraft engine comprising a gas turbine according to any of the preceding claims.
6. Aircraft engine according to the preceding claim, characterized in that a bypass ratio BPR of the aircraft engine is at least 10, in particular at least 12.
7. Method for designing at least one compressor stage of at least one compressor of a gas turbine, in particular of an aircraft engine, comprising a rotor cascade (3), a stator cascade (4), in particular adjacent thereto downstream, and a throttle value (σ), characterized in that the compressor stage is aerodynamically designed such that the throttle value σ and the aspect ratio AR_ax defined by the quotient of the average channel height (h) of the compressor stage or of the rotor cascade (3) and the average chord length (l_ax) of the compressor stage or of the rotor cascade (3) satisfy the condition $$\mathrm{\sigma }>-1.33\cdot {\mathrm{AR}}_{\mathrm{ax}}+5.16$$

   
8. Method for designing at least one compressor of a gas turbine, in particular of an aircraft engine, characterized in that at least one compressor stage of the compressor is designed according to the preceding claim.
9. Method for designing at least one compressor according to the preceding claim, characterized in that a total pressure ratio Π of the compressor is at least 40, in particular at least 45.
10. Method for designing a compressor according to any of the preceding claims, characterized in that a bypass ratio BPR of the aircraft engine is at least 10, in particular at least 12.

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