EP3510255B1 - Method for clear position determination and identification of a turbine blade - Google Patents

Description

The invention relates to a method for the unambiguous position assignment and identification of a turbine blade of a high-pressure turbine stage of a gas turbine, in particular of an aircraft engine.

The turbine blades of the high pressure turbine are among the most stressed components of a gas turbine or an aircraft engine. At the same time, even small damage or deviations from their respective original shape of the high-pressure turbine turbine blades have a not inconsiderable influence on the efficiency of the gas turbine, which ultimately has an effect on fuel consumption.

Due to the particular load, the high-pressure turbine blades are regularly overhauled or, if this is not possible, replaced with new parts. For this purpose, the gas turbine must be dismantled at least to such an extent that the turbine blades of the high-pressure turbine can be removed. The previously removed turbine blades can be reassembled after a possibly time-consuming overhaul. Alternatively, other turbine blades or new parts that have already been otherwise overhauled can be installed in order to keep the required maintenance time for the gas turbine as short as possible. The latter is particularly important in aircraft engines.

Turbine blades regularly have a unique serial number. For turbine blades of aircraft engines is this is even mandatory due to approval regulations. The serial numbers are arranged on the base of the turbine blades, with which the blades are attached to the blade carrier of the engine, in order to avoid that they come into direct contact with the hot gas flow and are made unrecognizable by it.

The disadvantage of positioning the serial number at the base of a turbine blade is that the serial numbers can no longer be viewed after the turbine blades have been inserted. In the case of a completely assembled gas turbine or a completely assembled aircraft engine, for example, it is not possible to check the serial number of an individual engine blade. In order to determine the serial number of a specific engine blade, according to the prior art, it is rather necessary to almost completely dismantle the entire gas turbine.

Particularly when using overhauled turbine blades, even if the turbine blades have been checked before and during assembly, shortly after the reassembled gas turbine has been put into operation, in some cases already after a first test run, an at least slightly reduced degree of efficiency can occur, which ultimately results in increased fuel consumption Use of the gas turbine results. Even if a correspondingly reduced efficiency often does not justify a renewed dismantling of the gas turbine, it can be desirable for various reasons that the component of the gas turbine causing the loss of efficiency is clearly identified. For example, a corresponding identification is desirable in order to be able to make possible recourse claims against the manufacturer of a clearly identifiable component or against the maintenance company that overhauled the component in question.

If a turbine blade of the high-pressure turbine of the aircraft engine is identified as the probable cause of the loss of efficiency, however, in the prior art it is not possible to clearly identify the turbine blade in question, e.g. by means of its serial number, without an almost complete dismantling of the gas turbine. As a result, the origin of the turbine blade from a manufacturer or repair shop cannot be easily clarified.

From the prior art - for example DE 10 2011 114 541 A1 , EP 1 609 957 A2 , US 2013/0 113 915 A1 , US 2014/0 188 423 A1 or the WO 2015/130 870 A1 - Various, also partially automated methods for boroscopic inspection of turbine blades or other components in a fully assembled aircraft engine are known. However, a way of identifying the turbine blades visible through the borescope is not described.

The documents DE 10 2011 103 003 A1 and US 9 016 560 B2 disclose automated optical methods for checking dismantled engine blades, wherein the serial number provided on the blade root can be designed to be machine-readable.

In the EP 2 597 273 A2 A method is disclosed for inspecting a turbo-machine, an image of the rotor blade taking place by means of borescopes in a first and second borescope opening upstream and downstream of a rotor blade and the images being forwarded to a processor for further analysis.

It is an object of the present invention to create methods with which an unambiguous position assignment and identification of a turbine blade of a high-pressure turbine stage of a gas turbine is possible without having to dismantle the gas turbine for identification. This object is achieved by the method according to claims 1 and 5. Advantageous developments are the subject matter of the dependent claims.

1. a. Ermitteln der Position eines vorgegebenen Schaufelschlosses einer vorgegebenen Stufe des Hockdruckkompressors, wobei die Stufe des Hockdruckkompressors derart gewählt ist, dass das Schaufelschloss im vollständig montierten Zustand der Gasturbine durch Drehen der Welle in das Sichtfeld eines durch eine erste Boroskopöffnung an der Gasturbine geführten ersten Boroskops bringbar und eindeutig identifizierbar ist; und
2. b. Markierung derjenigen Turbinenschaufel auf der Satzkarte, die im vollständig montierten Zustand der Gasturbine im Sichtfeld eines durch eine zweite Boroskopöffnung an der Gasturbine geführten zweiten Boroskops ist, wenn das vorgegebene Schaufelschloss im Sichtfeld des ersten Boroskops ist.

Accordingly, the invention relates to a method for the unambiguous position assignment of a turbine blade of a high-pressure turbine stage of a gas turbine connected via a shaft to a high-pressure compressor, each turbine blade of the high-pressure turbine being clearly identifiable with its relative position in a record map, with the following steps:

1. a. Determination of the position of a predetermined blade lock of a predetermined level of the high pressure compressor, the level of the high pressure compressor being selected such that the blade lock can be brought into the field of view of a first boroscope guided through a first boroscope opening on the gas turbine by rotating the shaft in the fully assembled state of the gas turbine is clearly identifiable; and
2. b. Marking of the turbine blade on the record card that is in the fully assembled state of the gas turbine in the field of view of a second boroscope guided through a second borescope opening on the gas turbine when the specified blade lock is in the field of view of the first boroscope.

1. a. Einführen eines ersten Boroskops durch eine erste Boroskopöffnung an der Gasturbine und eines zweiten Boroskops durch eine zweite Boroskopöffnung an der Gasturbine; und
2. b. Drehen der den Hochdruckkompressor und die Hochdruckturbinenstufe verbindenden Welle, bis das vorgegebene Schaufelschloss im Sichtfeld des ersten Boroskops ist, womit die auf der Satzkarte markierte Turbinenschaufel im Sichtfeld des zweiten Boroskops ist.

Furthermore, the invention relates to a method for the unambiguous identification of the turbine blades of the high pressure turbine of a gas turbine marked on the record card with the above method according to the invention, with the steps:

1. a. Inserting a first borescope through a first borescope opening on the gas turbine and a second borescope through a second borescope opening on the gas turbine; and
2. b. Rotate the shaft connecting the high-pressure compressor and the high-pressure turbine stage until the specified blade lock is in the field of view of the first boroscope, so that the turbine blade marked on the set card is in the field of view of the second boroscope.

First, some terms used in the context of the invention are explained.

Known in gas turbines and common in aircraft engines are "record cards" on which the "relative position" of the individual turbine or compressor blades of a gas turbine stage are recorded. All the blades used in the gas turbine stage are listed on the record card based on their serial number. In addition, the sequence of the individual blades in the circumferential direction emerges from the set card. In other words, the set map can be used to determine for each blade of the gas turbine stage which two blades are respectively arranged adjacent to it. However, the record card only provides information about the relative position of the blades, but not the absolute position, e.g. with respect to the shaft.

Gas turbine and aircraft engine blades include an airfoil and a root. The blade root regularly has a shape that enables the blade root to be pushed into a blade groove of the blade carrier of a turbine stage. To secure one or more blades in the blade groove, one or more "blade locks" are provided with which the blades are fixed in their position relative to the blade carrier.

The invention is based on the knowledge that the areas of the high-pressure turbine that can be seen through a second borescope in a fully assembled gas turbine are subject to the particular load caused by the hot gas coming from the combustion chamber, which is why all markings made in this area are removed or removed at the latest with a first test run of the gas turbine would be burned. Due to the particular load in this area, the high-pressure turbines are also regularly designed in such a way that the surfaces of the high-pressure turbine, which can be viewed through a second borescope when the gas turbine is fully assembled, cannot draw any conclusions about their angular position or the like. allow.

However, the invention has recognized that the high-pressure compressor, which is non-rotatably connected to the high-pressure turbine via a shaft, has clearly identifiable vane locks through a first borescope, even when the gas turbine is fully installed, which can serve as an identification feature for the angular position of the high-pressure compressor and thus the high-pressure turbine. With knowledge of the position of the borescope openings for the first and second borescopes, it is possible to determine precisely which turbine blade is in the field of view of the second boroscope when the specified blade lock is in the field of view of the first boroscope.

Since this relationship is recorded according to the invention on the record card during the assembly of the high-pressure turbine, a first turbine blade can initially be clearly identified on the basis of this information on the record card when the gas turbine is fully assembled. By subsequently rotating the high-pressure compressor by a predetermined angle or by simultaneous observation through the second borescope, the other turbine blades can also be clearly identified using the relative position of the turbine blades recorded on the record map.

In order to facilitate the determination of the turbine blade to be marked on the record card, it is preferred if the angular position of the specified blade lock of the high pressure compressor relative to the shaft is marked on the shaft itself and, based on this marking, the turbine blade to be marked on the record card is determined. The marking can be brought up to the high pressure turbine along the shaft, which simplifies the identification of the turbine blade in question. The marking can be carried out, for example, using a marking pen in accordance with the DIN LN 9051 standard.

The determination of the turbine blades to be marked on the record card can alternatively or additionally be made easier if the turbine blades have a temporary identification feature that is visible after assembly. E.g. the serial number of the individual turbine blades must be applied directly to the blade, i.e. at a point that is also visible after the turbine blade has been installed. Such an identification feature, which is also visible in the installed state, is basically a temporary identification feature which, after a first test run of the gas turbine, has already been burned without residue due to the particular loads in the area of the high-pressure turbine.

If there are more than two vane locks in the specified stage of the high-pressure compressor, a first and a second vane lock can usually be clearly identified. The distance in the circumferential direction opposite to a predetermined direction of rotation between the first shovel lock and the second shovel lock is usually less than the corresponding distance between the last shovel lock - which can also be the second shovel lock - and the first Shovel lock. Proceeding from the first shovel lock, which can be clearly identified in this way, the second shovel lock is the shovel lock following the first shovel lock in the specified direction of rotation in the circumferential direction. It is preferred if the specified shovel lock is the second shovel lock. As explained below, this can facilitate the later identification of the turbine marked on the record card.

If a turbine blade is marked according to the invention on the record card, this turbine blade can be clearly identified even when the gas turbine is fully assembled.

For this purpose, a first borescope is guided through a first borescope opening on the gas turbine and a second borescope is guided through a second borescope opening on the gas turbine. The first borescope opening is arranged so that the given blade lock of the given stage of the high pressure compressor can be seen through a passed borescope, while the second borescope opening allows the turbine blades of the relevant high pressure turbine stage to be viewed with the aid of a boroscope.

After introducing at least the first boroscope - the second borescope can also be inserted later - the shaft is turned on both the high pressure compressor and the high pressure turbine stage until the specified blade lock is in the field of view of the first boroscope. The turbine blade, which is then visible through the second borescope, which has now been introduced at the latest, is then the turbine blade marked on the record card.

If the specified vane lock is the second vane lock of the specified stage of the high-pressure compressor, the rotational speed can be determined after the first shovel lock can preferably be reduced by the first borescope in order to be able to stop the wave without delay as soon as the second shovel lock is in the field of view of the first borescope. In other words, the rotational speed can initially be comparatively high in order to quickly turn the shaft into the area in which the second shovel lock is located, in order then to be able to stop the shaft immediately due to a comparatively low rotational speed as soon as the second shovel lock is in the field of vision of the first Boroscope is coming.

If the turbine blade marked on the rate card is clearly identified, at least one turbine blade can be identified on the basis of the rate card, preferably by subsequently turning the shaft connecting the high pressure compressor and the high pressure turbine stage while observing the turbine blades through the second borescope or while observing the change in the angular position of the shaft. Since the relative position of the individual turbine blades can be seen on the record card, starting from the clearly identified, marked turbine blade, the respective adjacent turbine blades and successively the adjacent turbine blades can be identified. Since the basic geometric configuration of the turbine blades of a gas turbine is known, a turbine blade that is not adjacent to the highlighted turbine blade can also be identified directly via the angular position of the shaft, starting from the angular position that the shaft assumes when the marked turbine blade is identified.

Figur 1:

eine Hochdruckturbinenstufe mit einem über eine Welle verbundenen Hochdruckkompressor;

Figur 2:

eine Detaildarstellung des Hochdruckkompressors aus Figur 1; und

Figur 3:

ein Flugzeugtriebwerk mit der Hochdruckturbinenstufe und dem Hochdruckkompressor gemäß Figur 1.

The invention will now be described by way of example on the basis of an embodiment with reference to the accompanying drawings. Show it:

Figure 1:

a high pressure turbine stage with a high pressure compressor connected by a shaft;

Figure 2:

a detailed representation of the high pressure compressor Figure 1 ; and

Figure 3:

an aircraft engine with the high pressure turbine stage and the high pressure compressor according to FIG Figure 1 .

Figure 1 is a schematic representation of a high pressure turbine stage 10 of a gas turbine 1 or an aircraft engine 1 (see Figure 3 ). The high-pressure turbine stage 10 has a multiplicity of turbine blades 11, the respective serial numbers of which are engraved on the blade root and their relative position is recorded on a record card. A record card can have the form of a list of serial numbers, the order of the listing corresponding to the arrangement of the turbine blades in the circumferential direction of the high-pressure turbine stage.

The high-pressure turbine stage 10 is non-rotatably connected to a two-stage high-pressure compressor 30 via a shaft 20. Each of the two stages of the high-pressure compressor 30 comprises a plurality of compressor blades 31, which are each connected via two blade locks 32 (in the illustration according to FIG Figure 1 only a blade lock 32 is visible) are secured in their position relative to the blade carrier of the respective stage of the high-pressure compressor 30.

In Figure 2 the high pressure compressor 30 is off Figure 1 shown enlarged. Two blade locks 32, 32 'can be seen between the blades 31 of the one stage of the high-pressure compressor 30. The distance in the circumferential direction between the first shovel lock 32 'and the second shovel lock 32 is opposite to the predetermined direction of rotation 90 - the does not have to coincide with the direction of rotation of the aircraft engine 1 during operation - less than the distance between the second and last blade lock 32 in the direction of rotation and the first blade lock 32 '.

The shovel lock 32 identified in this way as the second shovel lock 32 serves as the basis for the marking 21 along the shaft 20, which is applied with an aviation-certified marking pen. The marking 21 indicates the angular position of the shovel lock 32 relative to the axis of the shaft 20. Starting from this marking, that turbine blade 11 'can be marked on the record card which - as below with the aid of Figure 3 is explained - in the fully assembled state of the aircraft engine 1 in the field of view of a second borescope 42 guided through a second borescope opening 5 on the aircraft engine 1, when the specified blade lock 32 is in the field of view of a first borescope 41 guided through a first borescope opening 4 of the aircraft engine 1.

In Figure 3 the aircraft engine 1 is shown in the fully assembled state. Only the turbine cover 2 is open (and therefore shown in dashed lines) in order to have access to the core engine 3.

For unambiguous identification of the turbine blade 11 ′ of the high-pressure turbine 10 marked on the record card, the shaft 20 is rotated in the direction of rotation 90 at a first rotational speed. As soon as the first shovel lock 32 ′ can be seen through the first borescope 41, the rotational speed is reduced and the shaft 20 is stopped as soon as the second shovel lock 32 is in the field of view of the first borescope 41.

If the second blade lock 32 is in the field of vision of the first boroscope 41, the turbine blade 11 marked on the record card is directly in the field of view of the second boroscope 42 due to the previously carried out position assignment of the turbine blade 11 in question 'can be clearly identified using the rate card.

Any markings provided on the turbine blades 11 during assembly, which can also facilitate the assignment of the position of the turbine blade 11 'to be marked, are completely removed after the first test run of the aircraft engine 1 at the latest. However, the method according to the invention makes it possible to first identify a turbine blade 11 '- namely the one previously marked on the record card - without any identification features on the turbine blade 11' or the high-pressure turbine 10 itself. Starting from this turbine blade 11 ', the other turbine blades 11 can also be clearly identified with the aid of the record card.

Claims (7)

1. Method for the clear position determination of a turbine blade (11) of a high-pressure turbine stage (10) of a gas turbine (1) that is connected to a high-pressure compressor (30) via a shaft (20), wherein each turbine blade (11) of the high-pressure turbine (10) is recorded on a record card in a clearly identifiable manner with the relative position thereof, comprising the steps:

   a. determining the position of a predefined blade lock (32) of a predefined stage of the high-pressure compressor (30), wherein the stage of the high-pressure compressor (30) is selected such that, in the completely mounted state of the gas turbine (1), the blade lock (32) can be brought into the field of view of a first boroscope (41) guided through a first boroscope opening (4) on the gas turbine (1) by rotating the shaft (20) and be identified clearly; and

   b. marking that turbine blade (11') on the record card which, in the completely mounted state of the gas turbine (1), is in the field of view of a second boroscope (42) guided through a second boroscope opening (5) on the gas turbine (1) when the predefined blade lock (32) is in the field of view of the first boroscope (41).
2. Method according to Claim 1,
   characterized in that
   when there is more than one blade lock (32, 32') in a predefined stage of the high-pressure compressor (30), the second blade lock (32) is the predefined blade lock, wherein the distance between the first blade lock (32') and the second blade lock (32) in the circumferential direction counter to a predefined direction of rotation (90) is smaller than the distance between the second blade lock (32) and the first blade lock (32') in the circumferential direction counter to a predefined direction of rotation (90) .
3. Method according to one of the preceding claims,
   characterized in that
   the angular position of the predefined blade lock (32) relative to the shaft (20) is marked on the shaft (20) and, starting from this marking (21), turbine blade (11') to be marked on the record card is determined.
4. Method according to one of the preceding claims,
   characterized in that
   the turbine blades (11) have an identification feature which is visible following the mounting in the high-pressure turbine stage (10) and which is removed by the operation of the gas turbine (1).
5. Method for the clear identification of the turbine blade (11') of the high-pressure turbine (10) of a gas turbine (1) which is marked on the record card by a method according to the preceding claims, comprising the steps:

   a. inserting a first boroscope (41) through a first boroscope opening (4) on the gas turbine (1) and a second boroscope (5) through a second boroscope opening (42) on the gas turbine (1); and

   b. rotating the shaft (20) connecting the high-pressure compressor (30) and the high-pressure turbine stage (10) until the predefined blade lock (32) is in the field of view of the first boroscope (41), whereby the turbine blade (11') marked on the record card is in the field of view of the second boroscope (42).
6. Method according to Claim 5,
   characterized in that
   with a second blade lock (32) as a predefined blade lock, the rotational speed is reduced following detection of the first blade lock (32') through the first boroscope (41) in order to stop the shaft without delay when the second blade lock (32) is in the field of view of the first boroscope (41).
7. Method according to Claim 5 or 6,
   characterized in that
   following clear identification of the turbine blades (11') marked on the record card, by means of subsequent rotation of the shaft (20) connecting the high-pressure compressor (30) and the high-pressure turbine stage (10) while observing the turbine blades (11) through the second boroscope (42) and/or while observing the change in the angular position of the shaft (20), at least one further turbine blade (11) is identified by using the record card.

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