EP2304187A1

EP2304187A1 - Axial turbo engine with low gap losses

Info

Publication number: EP2304187A1
Application number: EP09797510A
Authority: EP
Inventors: Georg Kröger; Eberhard Nicke; Christian Voss
Original assignee: Deutsches Zentrum fuer Luft und Raumfahrt eV
Current assignee: Deutsches Zentrum fuer Luft und Raumfahrt eV
Priority date: 2008-07-17
Filing date: 2009-07-16
Publication date: 2011-04-06
Also published as: CA2731092A1; US8647054B2; JP2011528081A; EP2304186B1; RU2011105788A; EP2304186A1; MX2011000649A; WO2010007137A8; EP2146053A1; JP5080689B2; RU2498084C2; CA2731092C; CN102099547A; WO2010007137A1; US20110189020A1; WO2010006975A1; MX336351B

Abstract

An axial turbo engine (1), comprises an impeller blade array, made up of impeller blades (3) each with a front edge (8), a radially outer free blade tip (15) and an annular enclosure (13) enclosing the impeller blade array with an annulus inner side (14) by means of which the annular enclosure (13) is arranged directly adjacent to the blade tips (15) to give a radial gap (16) between the enveloping ends of the blade tips (15) and the annulus inner side (14), wherein the impeller blades (3) have a radial projection (18) in the region of the front edge (8) thereof on the blade tip (15) and the annular enclosure (13) has an annular radial recess (17) in the annulus inner side (14) thereof, arranged at a radial distance (16) from the enveloping ends of the blade tips (15), such that in the main flow direction of the axial turbo engine (1) the line of the radial projections on the side thereof facing the radial gap (16) matches the line of the radial recess.

Description

Axial turbomachine with low gap losses

The invention relates to an axial turbomachine which has low gap losses.

An axial turbomachine includes a housing and a rotor surrounded by the housing. The rotor has a hub contour which, together with the inner contour of the housing, forms a flow channel through the axial turbomachine. The rotor has a plurality of rotor stages, each formed by a blade grid. The blade lattices have a plurality of rotor blades, each of which is fastened with its one end on the hub side to the rotor and with its other end radially to the rear show outside. At this other end of the blade, a blade tip is formed, which faces the inside of the housing and is disposed immediately adjacent. The distance between each blade tip and the inside of the housing is formed as a radial gap which is dimensioned such that on the one hand the blades do not touch the housing during operation of the axial turbomachine and, on the other hand, the leakage flow through the radial gap that occurs during operation of the axial turbomachine is as low as possible , In order for the axial turbomachine to have a high degree of efficiency, it is desirable for the leakage flow through the radial gap to be as low as possible.

If the axial turbomachine is installed in an aircraft engine, the housing is filigree designed to have the lowest possible weight. In contrast, the rotor is massively designed to withstand the pressure and temperature stresses during operation of the axial turbomachine can. The blades are less massive compared to the rotor and are mounted on the rotor.

In operation of the axial turbomachine, the inside of the housing and the blades are in contact with hot gas, the housing having large-area contact with the hot gas on its inside. The fact that the housing is filigree than the rotor, the rotor heats up more slowly than the housing. This has the consequence that for starting and stopping the axial turbomachine, the rotor and the housing have different Wärmeausdehnungsgeschwndigkeiten, so that when starting and stopping the axial turbomachine, the height of the radial gap formed between the blade tips of the blades and the inside of the housing , changes. It turns when starting the radial gap as large and when driving off as a small. Thus, the blade tips of the blades do not rub against the housing during the shutdown and damaged, the radial gap is provided with such a dimensioned minimum height that during Shutting down the axial turbomachine, the blade tips almost never touch the housing. This has the consequence that a correspondingly dimensioned radial gap is kept at the blade tips. On the other hand, especially when starting the axial turbomachine, the radial gap should be designed so large that a reduction in the power density and the efficiency of the axial turbomachine caused by the leakage flow keeps within acceptable limits.

Modern blades have a very high aerodynamic efficiency, which is achieved by a high pressure load of the blades. Caused by this high pressure load, the leakage flow through the radial gap is particularly strong, so that the total efficiency of the blade is greatly affected by the leakage flow. Especially with blades with small height and large radial gaps about 50% of the total loss of the blades is caused by the leakage flow. A reduction in the leakage flow causes an improvement in the overall efficiency of the blade.

It is conventionally known to reduce the leakage flow, for example, by means of an "active-clearance-controP" device. In the case of the "active-clearance-control" device, the housing is cooled during start-up and heated during retraction, so that the thermal expansion rate of the housing is increased Furthermore, a special shaping of the blade tips is known, as for example from US 4,738,586 the formation of a blade-shaped blade tip, for the reduction of the leakage flow.

Another blade tip contoured in the direction of the span of the blade is known from EP 675 290 A2. The blade tip and the opposite channel wall are contoured to each other corresponding to each other, wherein the channel wall having a circumferential recess and the blade tip having a radial tip extension adapted to the recess. Through this Measure can be achieved a rapid reduction in the gas velocity in the depression, which is the strength of shock waves weakened.

Another blade tip contouring and channel wall contouring is evident from FR 996967.

The object of the invention is to provide an axial turbomachine, which has a high aerodynamic efficiency. Another object of the invention is to provide a blade thereto.

The axial turbomachine according to the invention has a blade grid according to the features of claim 1.

The profiling of the blade of the axial turbomachine according to the invention may be conventional. The radial elevations of the blade extend parallel to the radial recess of the annular space inside in the main flow direction of the axial turbomachine, so that the radial gap has a uniform and multi-wavy course. The annular space inside and blade tip formed in the manner of a double shaft - and accordingly also the radial gap - each comprise at least four curved sections delimited by inflection points, the curves of adjacent curved sections having different signs. As a result, the leakage flow, which occurs during operation of the axial turbomachine through the radial gap, is alternately accelerated and decelerated. By the acceleration and deceleration, the overflow velocity and the direction of the leakage flow are changed so that a gap vortex, which forms when mixing the leakage flow with the main flow, is prevented in the formation phase and growth. This advantageously ensures that the flow through the blade grid is homogeneous and low loss, whereby the efficiency of the blade lattice and thus the axial turbomachine is high.

The gap which is constant in its gap has a uniform, not abruptly changing course along the main flow direction, so that the flow in the region of the blade tip has low losses.

In addition, due to the reduced influence of the leakage flow on the main flow, the working conversion of the blade is high and the flow of a guide blade downstream of the blade is improved. As a result, a Fehlanströmung this vane is reduced and / or the vane may have a simpler form.

The mass flow of the leakage flow and its unfavorable effect on the overall efficiency of the blade lattice are advantageously reduced. This results in having to provide without additional design measures, an improved aerodynamic quality of the blade.

Here, the radial distance of the radial recess to the envelope of the blade tips along the main flow direction of the axial turbomachine is constant.

Furthermore, it is preferred that the first curvature section is located in the region of the annular space inner side, which is opposite to the region of a front half of the chordail of the airfoil tip viewed from the front edge. In addition, it is preferred that the maximum of the first radial recess is located in the region or in that point, which is at 10% to 30%, preferably at 20%, the profile chord viewed from the front edge of view. Thus, advantageously, the radial elevation and the radial recess in the region of the highest pressure load of the blade tip of the conventionally profiled blade is settled, so that the leakage flow is reduced by the radial gap.

Furthermore, it is preferred that the curved sections are shaped in such a way that in the main flow direction of the axial turbomachine the course of the radial gap runs substantially edge-free and without steps. In this case, more than four curved sections, both in the annular space inside and at the blade tip, may be provided in order to reduce the leakage flow through the radial gap.

It is preferred that before the first bend section, after the fourth bend section and / or after an adjoining bend section, a further section is provided on the annular space inside or on the blade tip, whose course is straight in the main flow direction of the axial turbomachine.

Alternatively, it is preferable that the further portion (s) in the main flow direction of the axial turbomachine is conical.

As a result, a smooth transition from the radial recess or the radial elevations is achieved downstream of the trailing edge of the blade, so that the flow in the region of the blade tip is loss.

It is preferable that in the main flow direction of the axial turbomachine and from the leading edge, the first inflection point at 5% to 15%, preferably at 10%, the chord length of the blade and / or the bottom of the radial recess at 15% to 25%, preferably at 20%, which is located chord length of the blade. In principle, the contour of the blade tip and the opposing annular space inside are always corresponding, so that both contours are identical to describe. Thus, the advantages occurring for an axial turbomachine apply mutatis mutandis to a blade.

The axial turbomachine is preferably a compressor in a stationary gas turbine, in an aircraft engine, in a process compressor, in a fan, in a fan, in a steam pressure turbine, in a hydraulic turbine and / or in a pump.

In the following the invention will be explained with reference to a preferred embodiment of an axial compressor according to the invention with reference to the accompanying schematic drawings.

Show it:

Fig. 1 is a side view of a housing portion of a first

Embodiment of the axial compressor according to the invention,

Fig. 2 is a perspective view of a blade tip of

Embodiment of Fig. 1,

Fig. 3 is a side view of the first and a second embodiment of a housing of the axial compressor according to the invention, and

Fig. 4 shows another embodiment of a Ringraumwandung with a correspondingly adapted contour of the blade tip.

As is apparent from FIGS. 1 to 3, an axial compressor 1 has a rotor 2 having a blade lattice formed by a plurality of Blades 3 is formed. Seen in Figures 1 and 3, the axial compressor

I flows through from left to right.

Furthermore, the axial compressor 1 has a first stator 4 upstream of the rotor blade 3 and a second stator 5 downstream of the rotor blade 3. The first stator 4 is formed by a plurality of first vanes 6 and the second stator 5 is formed by a plurality of second vanes 7.

The blade 3 has at its upstream end a leading edge 8 and at its downstream end facing a trailing edge 9, wherein the one side between the front edge 8 and the rear edge 9, the pressure side 10 and the other side between the front edge 8 and the trailing edge. 9 the Saugseϊte

II of the blade 3 is. From the leading edge 8 to the trailing edge 9 extends a straight chord having a normal chordal length of 100%, with the starting point equal to 0% of the chord length of the chord at the leading edge and 100% of the chord length of the chord at the trailing edge.

The rotor blade 3 is surrounded radially on the outside by an annular space wall 13, wherein the annular space wall 13 has an annular space inner side 14, which faces the rotor blade 3. The blade 3 is attached to its radially inner longitudinal end and free-standing with its radially outer longitudinal end, wherein at the freestanding end of a blade tip 15 is formed. Between the annular space inside 14 and the blade tip 15, a gap 16 is provided.

The blade tip 15 is provided on its side facing the annular space 14 side with a radial elevation 18, the maximum radial extent is located at 20% of the chord length of the chord 3 bucket. Following the course of the blade tip 15, in a section of the annular space inner side (14) which lies opposite the blade tip 15, a radial depression 17 is provided in the annular space wall 13 on the annular space inside 14, which is shown in FIG and 3 as seen from left to right runs parallel to the radial elevation 18. The radial recess 17 has a base 12, which is arranged radially outside at the height of the maximum radial extent of the radial elevation 18.

The radial recess 17 is formed circumferentially in the annular space wall 13. As a result, when the rotor 2 rotates about the rotation axis 28, each blade 3 with its radial elevation 18 can rotate into the radial recess 17 in an engaging manner.

Seen in the main flow direction of the axial compressor 1, the radial recess 17 and, analogously, the radial elevation 18 of a first embodiment of four curved sections 19, 21, 23, 25 are formed, wherein the curved sections 19, 21, 23, 25 each have a curvature, wherein the sign the curvatures of curvature section to curvature section changes. The curved sections 19, 21, 23, 25 are arranged lying one behind the other, wherein the first curved section 19 is delimited from the second curved section 21 by a first turning point 20. Furthermore, the second curvature portion 22 is separated from the third curvature portion 23 by a second inflection point 22. The third curvature section 23 is delimited from the fourth curvature section 25 by a third inflection point 24. Due to the successive arrangement of the curved portions 19, 21, 23, 25 and the intermediate inflection points 20, 22, 24, the gap 16 between the blade tip 15 and the annular space inside 14 is formed like a wave.

In Fig. 2, the radial boundary of a conventional blade tip is shown with the line 27, so that the radial projection caused by the provision of the radial projection 18 is clear.

In Fig. 3, the course of the annulus inner side 14 is shown in relation to a rotation axis 28 of the axial compressor 1 according to the first embodiment. Upstream of the first curvature section 19 and / or downstream of the fourth curvature section 25, the course of the annular space inner side 14, at least in that region which is opposite the airfoil tip 15, is conical and thus rectilinear in order to obtain a channel contraction.

Of course, it is possible that the course of the annular space inside 14 and accordingly the blade tip 15 not only has four curved sections, but even more curved sections.

FIG. 4 shows an exemplary embodiment of an annular space wall 13 with a correspondingly adapted contour of the blade tip 15, in which numerous further curved sections 21, 23, 25, 27, 29, 31, 33, 35 are provided behind an upstream first curved section 19. Adjacent curvature sections are ϊnvers curved to each other. This means that when one curvature section is concavely curved, the next following curvature section is convexly curved, and vice versa. Concave and convex curvature sections alternate. Between them there is a turning point W1, W2, W3, W4, W5, W6, W7, W8.

In Figure 4, the chord of the Ringraumwandung 13 is designated 40. The chord falls from the inflow end to the discharge end of the flow channel. The central axis of the annular space bounded by the annular space 13 is designated A. The terms maximum and minimum are based on this axis A. The largest maximum, which is referred to as the main maximum, is located in the curved section 21. It is the first maximum within the axial length of the blade. This is followed by smaller maxima in the curvature sections 25, 29 and 33.

The second curvature portion 21 thus forms, with respect to the axis A of the annular space wall 13 and the blade, the main maximum. Every second of the following curvature sections, namely the curvature sections 25, 29 and 33, forms another maximum of smaller amplitude than the main maximum.

The amplitudes of all maxima fall monotonically, starting with the main maximum.

The amplitudes of the maxima preferably fall off according to a second-order function, according to the formula

Y = cX ² + bX + C

Here, Y forms the amplitude (perpendicular to axis A), X the position along axis A. c and b are constant coefficients, and C is a constant. The coefficients c and b are not equal to zero.

Between two adjacent maxima are exactly two turning points.

The moving blade 3 is not yet located in the first curved section 19. This starts only at the first turning point W 1. Consequently, the first curvature portion 19 refers exclusively to the annular space wall 13 which is convexly curved in this curvature portion. In the region of the first inflection point W1 there is an increase and then the main maximum follows with a subsequent decrease to the inflection point W2.

The embodiment of Figure 4 is particularly, but not exclusively, suitable for operating in the transonic region front stage of a multi-stage gas turbine. By disturbing the resulting during overflow of the blade tip vortex of the efficiency of the axial turbomachine is increased.

Claims

claims

An axial turbomachine (1) with a blade grid formed by blades (2) each having a leading edge (8) and a radially outer freestanding blade tip (15), and an annular space wall (13) enveloping the blade grid (8) having an annular space inside (Fig. 14), with which the annular space wall (13) is arranged immediately adjacent to the blade tips (15) forming a radial gap (16) between the envelope of the blade tips (15) and the annular space inside (14), wherein within one of the blade tip (15) the annular space wall (13) on the annular space inside (14) at least one circumferential first radial recess (17) having a first and a second curved portion (19, 21) in along the Hauptdurchströmungsrichtung of the axial turbomachine (1 ) constant radial distance to the envelope of the annular space inside (14) corresponding Scha ufelspitzen (15) is arranged,

characterized in that

in the main flow direction of the axial turbomachine (1), the course of the annular space inside (14) at least

• a third curvature section (23) adjoining the second curvature section (21) and

Having a fourth curvature section (25) adjoining the third curvature section (23), in which

The first curvature section (19) of the second curvature section (21) having a first inflection point (20),

The second curvature section (21) of the third curvature section (23) with a second inflection point (22) and

The third curvature section (23) is delimited from the fourth curvature section (25) by a third inflection point (24),

wherein the curvatures of adjacent curved sections (19, 21, 23, 24) have different signs, and

in that the second curvature section (21) forms a main maximum with respect to the axis (A) of the annular space wall (13) and the blade, and that every second (25, 29, 33) of the subsequent curved sections (23-35) constitutes a further maximum of smaller amplitude than the main maximum.

An axial turbomachine (1) according to claim 1, wherein the first curved portion (19) is located in that portion of the annular space inner side (14) opposite to the front half portion of the chord blade tip (15) from the front edge (8) ,

3. axial turbomachine (1) according to one of claims 1 or 2, wherein the maximum of the first radial recess (17) is located in the region or in that point, which at 10% to 30%, in particular at 20%, the chord of the front edge (8) seen from opposite.

4. axial turbomachine (1) according to one of claims 1 to 3, wherein the curved portions (19, 21, 23, 25) are formed such that along the main flow direction of the axial turbomachine (1), the course of the radial gap (16) substantially edge-free and runs steplessly.

5. axial turbomachine (1) according to one of claims 1 to 4, wherein in the main flow direction of the axial turbomachine (1) the course on the annular space inside (14) further to the fourth curved portion (23) adjoining Krümmungsabschnϊtte.

6. axial turbomachine (1) according to one of claims 1 to 5, wherein before the first curved portion (19), at the fourth curved portion (23) and / or at the fourth curved portion (23) subsequent further curved portion another portion of Annular space inside (14) is provided, whose course in the main flow direction of the axial turbomachine (1) is straight.

An axial turbomachine (1) according to claim 6, wherein said further portion (s) is (are) cone-shaped.

8. axial turbomachine (1) according to one of claims 4 to 7, wherein seen in the main flow direction of the axial turbomachine (1) and from one of the leading edge (8) of the blade opposite point of the first inflection point (20) at 5% to 15%, in particular at 10%, the chord length of the blade (3) and / or the bottom of the radial recess at 15% to 25%, in particular at 20%, of the chord length of the blade (3) is located.

9. axial turbomachine (1) according to claim 1, characterized in that the amplitudes of all maxima, starting from the main maximum, monotonically decrease.

10. axial turbomachine (1) according to claim 1, characterized in that the amplitudes of the maxima decrease according to a second order function, according to the formula

Y = cX ² + bX + C,

where Y is the amplitude, X is the position along the axis (A) and c, b and C are constants.

11. Blade (2) for an axial turbomachine, with a front edge (8) and a radially outer freestanding blade tip (15) at its blade tip (15) at least a first radial elevation (18) having a first curvature portion (19) and a second curved portion (21), characterized in that along a chord of the blade (2) the blade tip (15) at least

Having a fourth curvature section (25) adjoining the third curvature section (23),

in which

the first curvature portion (19) of the second curvature portion (21) having a first inflection point (20),

the second curvature portion (21) of the third curvature portion (23) having a second inflection point (22) and The third curvature section (23) is delimited from the fourth curvature section (25) by a third inflection point (24),

12. Blade (2) according to claim 11, whose first radial elevation (18) is located in the region of the front half of the chord length.

13. The moving blade (2) according to claim 11, wherein the maximum of the first radial elevation (18) is located at 10% to 30%, in particular at 20%, of the chord length from the leading edge (8).