US20160115966A1

US20160115966A1 - Guide vane ring for a turbomachine and turbomachine

Info

Publication number: US20160115966A1
Application number: US14/886,659
Authority: US
Inventors: Joachim Wulf
Original assignee: MTU Aero Engines AG
Current assignee: MTU Aero Engines AG
Priority date: 2014-10-27
Filing date: 2015-10-19
Publication date: 2016-04-28
Also published as: EP3015715A1

Abstract

The present invention relates to a guide vane ring (1) for a turbomachine, having a plurality of rotatable guide vanes (3), and having an inner ring (5), wherein the inner ring (5) has a seal (17) for sealing a radial gap between the inner ring (5) and an opposite-lying rotor segment (7), and wherein the inner ring (5) comprises at least two inner ring segments (11). The inner ring (5) is produced from a material or has a material that has a heat expansion coefficient α of less than 6*10⁻⁶per Kelvin in a temperature range between at least 20 degrees Celsius and 90 degrees Celsius. In addition, the present invention relates to a turbomachine having at least one guide vane ring (1) according to the invention.

Description

BACKGROUND OF THE INVENTION

The studies that have led to this invention were supported according to the Financial Aid Agreement No. CSJU-GAM-SAGE-2008-001 under the Seventh Framework Program of the European Union (FP7/2007-2013) for Clean Sky Joint Technology Initiative.
The present invention relates to a turbomachine and a guide vane ring for a turbomachine.
Guide vane rings for turbomachines are subjected to intense temperature fluctuations during operation. For example, when starting up an aircraft engine, temporarily high temperature gradients form in the inner rings of guide vane rings, which can lead to deformations. These deformations that regress again in the further operation of the engine after the startup procedure may influence secondary flows, such as, for example, leakage flows between inner rings on guide vanes and rotors. These influences reduce the overall efficiency of the engine.

SUMMARY OF THE INVENTION

An object of the present invention is to propose a guide vane ring for a turbomachine that reduces the leakage flow between a rotor segment and an inner ring. In addition, an object of the present invention is to propose a turbomachine with a guide vane ring according to the invention.
The object according to the invention is achieved by a guide vane ring, which is discussed in detail below. It is further achieved by a turbomachine according to the present invention.
Thus, according to the invention, a guide vane ring for a turbomachine, in particular for a compressor, is proposed, which comprises a plurality of rotatable guide vanes and an inner ring. The inner ring has a seal for sealing a radial gap between the inner ring and an opposite-lying rotor segment. The inner ring comprises at least two inner ring segments.
The inner ring is produced from a material or has a material that has a heat expansion coefficient α of less than 6*10⁻⁶per Kelvin in a temperature range between at least 20 degrees Celsius (° C.) and 90 degrees Celsius (° C.). The lower and upper temperature values may vary, each time depending on the application of the guide vane ring according to the invention. For example, aircraft engines may have different temperatures in different operating states. The upper temperature value can be slightly or clearly greater than 90° C., for example 150° C., 300° C., 500° C., 800° C., or another value. For example, the lower temperature value can be less than 20° C., for example 0° C., −10° C., −20° C., or another value, each time depending on the location of an aircraft with an aircraft engine that has a guide vane ring according to the invention.
The heat expansion coefficient α can be called a coefficient of linear thermal expansion. The value and the unit of the coefficient of thermal linear expansion α of 6*10⁻⁶per Kelvin can be represented as 6*10⁻⁶/K or as 6 ppm/K (ppm=parts per million).
The heat expansion coefficient α of the material of the inner ring can be less than 6*10⁻⁶/K, for example, the heat expansion coefficient α can have a value of 5*10⁻⁶/K, 2*10⁻⁶/K, 1.7*10⁻⁶/K, 1.2*10⁻⁶/K, 0.55*10⁻/K, or another value. An inner ring having these values for the heat expansion coefficient α can advantageously have a linear expansion that is reduced by approximately 60% to 65% in comparison to the usually employed stainless iron-nickel-chromium alloys (the parts are indicated in weight percent for these alloys in the overall discussion of advantages given below) with clearly higher values of the heat expansion coefficient α (for example a heat expansion coefficient α of 15*10⁻⁶/K, 20*10⁻⁶/K, 25*10⁻⁶/K, or 30*10⁻⁶/K). By means of this reduced linear expansion, it can be achieved advantageously that, in particular, the end regions are deformed to a lesser extent in the peripheral direction of inner ring segments when compared to inner ring segments of the usually employed stainless iron-nickel-chromium alloys. This smaller deformation of the inner ring segments can have the consequence that sealing fins produce incisions that are less pronounced in inlet seals at the radially inner ends of the guide vane ring, and thus leakage flows can be reduced in this region.
The guide vane ring according to the invention may have adjustable guide vanes and/or inlet seals. Upon first-time startup, sealing gaps can be forged or cut in between the inlet seals and sealing fins on radially opposite-lying rotors by means of these inlet seals. The sealing gaps, in which leakage flows usually form during operation of the turbomachine, can be reduced or minimized in this way.
Advantageous enhancements of the present invention are the subject of each of the dependent claims and embodiments.
Exemplary embodiments according to the invention may have one or more of the features named in the following.
In particular, gas turbines are described in the following as turbomachines purely by way of example. but without wanting to limit turbomachines to gas turbines. The turbomachine can be an axial turbomachine, in particular. The gas turbine can be an axial gas turbine, in particular, for example an aircraft gas turbine.
In specific embodiments according to the invention, the material of the inner ring has a heat conductivity λ of more than 10 watts per meter and per Kelvin (10 W/(m*K)) at a temperature between 20° C. and 25° C., in particular at 23° C. The heat conductivity X may be, for example, 13 W/(m*K), 15 W/(m*K), 30 W/(m*K), 50 W/(m*K), or another value. Advantageously, the heat from the inner ring segments can be rapidly conducted further or discharged by means of a high heat conductivity value λ, and thus a local deformation of the material can be avoided. In this way, for example, the formation of a large sealing gap between a sealing fin and an inlet seal can be reduced or avoided, and advantageously a leakage flow is minimized.
In certain embodiments according to the invention, the inner ring comprises at least two divided inner ring segments on the periphery of the inner ring. The ring segments can each have a peripheral angle of 180 degrees (180°) as so-called half rings. The ring segments can also have other peripheral angles, for example, 120° and 240°. The inner ring may have more than two ring segments, for example three ring segments, each having a 120° peripheral angle, four ring segments each having a 90° peripheral angle, or other values.
In several embodiments according to the invention, the material of the inner ring is a nickel-alloyed steel. The nickel fraction in the material may comprise at least 25 weight percent.
In many embodiments according to the invention, the material of the inner ring has an iron fraction of at least 50 weight percent.
In specific embodiments according to the invention, the material of the inner ring has a cobalt fraction of at least 10 weight percent.
In several embodiments according to the invention, the material of the inner ring has an iron fraction between 62 weight percent and 66 weight percent, in particular 64 weight percent, and a nickel fraction between 34 weight percent and 38 weight percent, in particular 36 weight percent.
In some embodiments according to the invention, the material of the inner ring has an iron fraction between 52 weight percent and 56 weight percent, in particular 54 weight percent, a nickel fraction between 27 weight percent and 31 weight percent, in particular 29 weight percent, and a cobalt fraction between 15 weight percent and 19 weight percent, in particular 17 weight percent.
Some or all embodiments according to the invention may have one, several, or all of the advantages named above and/or in the following.
By means of an inner ring according to the invention, designed in particular as the inner ring of a compressor, which is produced from a material having an iron fraction of approximately 54 weight percent and having a nickel fraction of approximately 36 weight percent, in a temperature range between 20° C. and 500° C., a linear expansion reduced by approximately 60% to 65% can be achieved advantageously in comparison to one of the following materials (the percentage data refer to weight percents):
1) stainless steel (iron-nickel-chromium alloy) having the following components in weight percent: 0.03 to 0.08% carbon (C), less than or equal to 1% silicon (Si), 1 to 2% manganese (Mn), less than or equal to 0.025% phosphorus (P), less than or equal to 0.015% sulfur (S), 13.5 to 16% chromium, 1 to 1.5% molybdenum (Mo), 24 to 27% nickel (Ni), 0.1 to 0.5% vanadium (V), 1.9 to 2.3% titanium (Ti), 0.003 to 0.01% boron (B), less than 0.35% aluminum (Al), with the remaining fraction: iron (Fe).
2) stainless steel (iron-nickel-chromium alloy) having the following components in weight percent: less than 0.08% carbon (C), less than 0.35% silicon (Si), less than 0.35% manganese (Mn), less than 0.015% phosphorus (P), 0.2 to 0.8% aluminum (Al), less than 0.6% boron (B), less than 1% cobalt (Co), 17 to 21% chromium, less than 0.3% copper (Cu), 2.8 to 3.3% molybdenum (Mo), 4.75 to 5.5% niobium (Nb), 50 to 55% nickel (Ni), 0.65 to 1.15% titanium (Ti), with the remaining fraction: iron (Fe).
The thermal deformation can be considerably reduced by means of the inner ring according to the invention (see above). Thus, the rubbing of sealing tips of the rotor into inlet seals of the inner ring can be reduced, in particular, in temporary (transient) operating states, such as, for example, during the startup of an aircraft engine. This reduced rubbing-in advantageously can lead to a permanent reduction of leakage flows between the inner ring and the rotor.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

The present invention will be explained in the following by an example based on the appended drawings, in which identical reference numbers designate identical or similar components. In the schematically simplified figures:

FIG. 1 shows a guide vane ring according to the invention with a rotatable guide vane, an inner ring, and a rotor segment;

FIG. 2 shows an inner ring segment in perspective representation; and

FIG. 3 shows a gas turbine with a guide vane ring according to the invention in a schematically greatly simplified manner.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 shows a guide vane ring 1 according to the invention, having a rotatable guide vane 3, an inner ring 5, and a rotor segment 7. The guide vane 3 is disposed rotatably in the inner ring 5 by means of an inner journal 9. The inner ring 5 comprises divided inner ring segments 11 in the peripheral direction u.
The rotor segment 7 is joined to a rotating blade 13. Another upstream rotor segment 15 is flanged to the rotor segment 7.
The inner ring segments 11 are joined to seals, in particular to inlet seals 17, at their radially inner ends. A leakage flow 21 can form during the operation of the turbomachine between the inlet seals 17 and sealing tips or sealing fins 19, which, in particular, are joined integrally to rotor segment 7. The leakage flow 2 usually runs counter to the primary flow direction 23 of the turbomachine (dependent on the pressure ratios upstream and downstream of the guide vane ring 1).
In turbomachines, in particular in compressors of aircraft engines, the rotor segments 7 and 15, the sealing fins 19, the inner ring segments 11, the rotating blades 13, as well as the guide vane ring 1 are often subjected to high temperature fluctuations approximately between 20° C. and 500° C. Both the lower temperature range as well as the upper temperature range can be shifted still further, each time depending on the application and the operating conditions. The named components can expand, bend, or change their shape in another way, each time depending on the materials employed. In particular, the linear thermal expansion of the inner ring segments 11 can influence the gap width between the inlet seals 17 and the sealing fins 19 and thus change the leakage flow 21.
The inner ring 5 can be divided or segmented in the axial direction a and/or in the radial direction r and/or in the peripheral direction u (as inner ring segments 11 in FIG. 1). For reasons of assembly, the inner ring 5 is often designed in the shape of two half-ring segments, each of which may have a peripheral angle of 180 degrees. The inner ring 5 can also be segmented differently in other embodiments, for example into three segments, each with a 120 degree peripheral angle, into four segments, each with a 90 degree peripheral angle, etc.
In one possible embodiment of a turbomachine as an aircraft engine, a radial temperature gradient can build up temporarily in the inner ring segments 11, in particular upon startup of the engine. The inner ring segments 11 then have a higher temperature radially outside (on the outer radius) than radially inside (on the inner radius). Based on this temperature gradient, the ends of the inner ring segments 11 can bend radially inward temporarily (during the engine startup process), considered in the peripheral direction. Based on this temporary bending of the inner ring 5, an increased rubbing of the sealing fins 19 into the inlet seals 17 can result. After the inner ring 5 has completely heated throughout (after the engine startup procedure), the temperature gradient of the inner ring segments 11 can be reduced again in the radial direction r, and the deformation can regress again, for example, approximately back to the initial state. The gap which is formed between the sealing fins 19 and the inlet seal 17 due to the temporary deformation remains or exists, however, after the regression of the inner ring 5. This gap can effect or generate an elevated, possibly permanent leakage flow 21. The efficiency of the engine can be permanently reduced in this way.
The described effect of the temporary deformation of the inner ring 5 based on temperature gradients can be designated as the so-called “cording effect” or “cording”. The “cording effect” is a thermal effect primarily in the case of inner rings 5, which can lead to a three-dimensional deformation of the inner ring segments 11 at the dividing planes (in the peripheral direction u). These deformations can lead to a greater run-in of sealing fins 19 into the inlet seals 17, whereby the sealing gaps and leakages can increase. Greater leakages can reduce the efficiency.
The run-in (elevated rubbing of the sealing fins 19 into the inlet seals 17) on the inner rings 5 can be advantageously reduced by means of the guide vane ring 1 according to the invention with the material properties named in the claims. Possible leakage losses due to elevated leakage flows 21 can at least be reduced advantageously.
FIG. 2 shows an inner ring segment 11 with a separating plane 25 in perspective representation. The inner journals 9 of the adjustable guide vanes 3 are inserted into the depressions 27 on the radial outer side of the inner ring segment 11.
FIG. 3 shows in schematically very simplified manner a gas turbine 29 as an embodiment of a turbomachine according to the invention with a guide vane ring 1 according to the invention, which is disposed, for example, in the high-pressure compressor section of a gas turbine 29.

Claims

What is claimed is:

1. A guide vane ring (1) for a turbomachine, having a plurality of rotatable guide vanes (3), and having an inner ring (5), wherein the inner ring (5) has a seal (17) for sealing a radial gap between the inner ring (5) and an opposite-lying rotor segment (7), and wherein the inner ring (5) comprises at least two inner ring segments (11), wherein the inner ring (5) is produced from a material or has a material that has a heat expansion coefficient α of less than 6*10⁻⁶per Kelvin in a temperature range between at least 20 degrees Celsius and 90 degrees Celsius.

2. The guide vane ring (1) according to claim 1, wherein the material of the inner ring (5) has a heat conductivity λ of greater than 10 watts per meter and per Kelvin at a temperature between 20 degrees Celsius and 25 degrees Celsius.

3. The guide vane ring (1) according to claim 1, wherein the inner ring (5) comprises at least two divided inner ring segments (11) on the periphery of the inner ring (5).

4. The guide vane ring (1) according to claim 1, wherein the material of the inner ring (5) is a nickel-alloyed steel.

5. The guide vane ring (1) according to claim 1, wherein the material of the inner ring (5) comprises a nickel fraction of at least 25 weight percent.

6. The guide vane ring (1) according to claim 1, wherein the material of the inner ring (5) comprises an iron fraction of at least 50 weight percent.

7. The guide vane ring (1) according to claim 1, wherein the material of the inner ring (5) comprises a cobalt fraction of at least 10 weight percent.

8. The guide vane ring (1) according to claim 1, wherein the material of the inner ring (5) has a heat expansion coefficient α of less than 6*10⁻⁶per Kelvin in a temperature range between at least 20 degrees Celsius and 500 degrees Celsius.

9. The guide vane ring (1) according to claim 1, wherein the guide vane ring (1) is configured and arranged in a turbomachine.

10. The guide vane ring (1) according to claim 9, wherein the turbomachine is an axial high-pressure compressor.