Hollow gas-turbine blade, and a method for external-film cooling thereof

(42) Date of commencement of the patent: 15. 4.1997 (45) Date of issue: 25. 9.1998 (56) Documentation:

(73) Patent owner:

EP 326508A1 GB U81481A US 4384452A US 4705455A US 4726735A US 4859147A

INSTITUTE OF THERMAL TURBQUE MACHINES AND MACHINE SYNAMICS TECHNICAL UNIVERSITY GRAZ A-8010 GRAZ, STYRIA (AT).

(72) Inventor:

3ERICHA HERBERT DR.

VIENNA (AT).

DISCOUNT JACOB DR. GRAZ, STEIERMARK (AT). SANZ WOLFGANG DR.

GRAZ, STYRIA (AT).

(54) HOLLOWING GAS TURBINE BLADES AND OUTSIDE FILM COOLING METHODS OF THE SAME (57) External film cooling of hollow gas turbine blades by means of under-expanded, large-area cooling films emerging from slots with sound velocity, which are related by a sufficiently high pressure ratio 2 to 3 of the refrigerant in the feed be generated at the pressure of the gas main flow at the exit point. A safe and high cooling effect is achieved by the use of the property of such a jet, the bottom expands from a cooling slot exits, by the further expansion also to the possibly highly curved profile surface (profile nose) aniegt and thereby forms large-scale cooling films, by the turbulence of the Mainstream are influenced only slightly.

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The present invention relates to a method for external film cooling of hollow gas turbine blades with the aid of under-expanded, emerging from slots with sound velocity, large-area cooling films according to the preamble of patent claim 1, and the formation of a gas turbine blade for use of this cooling method according to claim 2.

Modern gas turbines operate at inlet temperatures in the range of 1300 degrees Celsius for industrial and power plants, for aircraft engines, these temperatures have already risen to 1500 degrees Celsius. This was only possible due to the development of modern materials and cooling methods. Even the most modern materials only allow metal temperatures in the range of 800 to 900 degrees Celsius, so that a range of several hundred degrees Celsius must be produced by the blade cooling as a temperature difference between the gas temperature and the metal temperature.

Particularly successful were the so-called film cooling process in which cooling air is fed into the gas stream through a series of holes from the inside of the hollow blade. Obviously, this cooling air may be blown out of the bores only at speeds which are a little greater than the gas velocity, otherwise the single cooling air jet will not touch the surface of the blade and produce the intended cooling effect. Moreover, so far only a series of relatively thin holes have been made in the range of one half to one millimeter diameter, so that the emerging from these holes cooling air jet is laterally enclosed by the gas flow and it is a difficult flow problem, these speeds and the exit angle of the cooling air jet from the Select surface so that the desired formation of a cooling film is effected. This is particularly difficult at the leading edge of a turbine profile and also in the vicinity of the narrowest cross section on the suction side, as in the first case, although a laminar but very thin boundary layer in this zone prevails and in the latter case, a highly turbulent boundary layer rapid mixing of the cooling air jet with the Mainstream causes.

In one embodiment of the present invention closest to EP 0 326 508 A1, a cooling film is passed through slots to the surface of the blade, but the amount of coolant which is introduced into the slot, is limited by a Zumeßbohrung, so that in Exit of the slot to the tread surface only subsonic speed may exist, similar in size as the velocity of the boundary layer on the surface of the blade. Thus, the intention of this design solution is clear, just to introduce the cooling film in the lower zone of the boundary layer directly to the surface of the blade at low speed. As shown in Figure 2 of this patent, such cooling slots are provided on both sides of the rounded profile nose in the direction of flow of the main flow ausblasend, so that also here the above-mentioned low speed of the exiting film is apparent. This patent therefore brings progress over the solutions that blow out directly from individual rows of holes and where there is a risk of the boundary layer penetrating the beam. However, the cited patent does not disclose a solution to effectively cool the rounded profile nose itself in which, due to the proximity of the stagnation point and the laminar boundary layers of small thickness developing from both sides thereof, a particularly high heat transfer exists.

It is known that under-expanded rays, which emerge at a speed of sound in a space of lower pressure, in the subsequent expansion to invest in a curved surface, even if it deviates greatly from the straight direction of the beam. The object of the present invention is to make use of this effect for cooling blade profiles of gas turbines. This object is achieved with the characterizing features of claims 1 and 2 in the way, that slits are incorporated in the surface, the inside of the hollow shell, which forms the gas turbine blade, be supplied with coolant, this coolant experiences a steady acceleration up to the narrowest cross section of the slot directly at the surface, in which then the speed of sound prevails, if the coolant is supplied with sufficient pressure increase, So with a pressure ratio of two to three based on the ratio of the feed pressure to the static pressure of the gas main flow at the surface.

It occurs in the narrowest cross section, just in the slot, speed of sound. The influx to the narrowest cross-section is expediently designed by two slightly mutually inclined planes that produce an acceleration to the slot, on the other hand, the lateral boundary is designed in a more or less circular or elliptical contour.

The transition of the slot in the profile surface of the blade is now designed according to the invention with a radius, the tangential both to the inner wall of the inflow slot, as well as outwards to tangentially merges into the profile contour of the blade, causing an acceleration to supersonic of the film, the large surface both in the longitudinal direction of the blade, as well as the direction of the expanding flow along the profile bears against this blade surface, through the Relativge2

AT 404 160 Β speed causes a good cooling effect and just a good heat transfer coefficient. Due to the properties of the under-expanded jet, fed with the high pressure ratio according to the invention in the main flow, it exerts the desired large-scale cooling effect, without, as in a single hole, the coolant breaks through the boundary layer of the main flow and thus without exerting its cooling effect in hot Gas flow is lost.

The invention will now be explained with reference to an exemplary embodiment, wherein the figures of the figures show the following: 1a and b show elevation and plan view of a gas turbine blade wherein SP represents the blade profile, KK the cooling channels, KS the cooling slots according to the invention and KSEK the cooling slot on the profile nose (leading edge) arranged on the pressure side of the profile, directed against the main flow. The attachment of the blade in the rotor disk, for example, by a Christmas tree foot TF. FIG. 1c shows a section through the cast gas turbine blade of the gas turbine blade with the gas flow GS, the coolant flow KM to the feed cylinder SZ, with the cooling slot KS, formed from the outer plane AE and the inner plane IE, in conjunction, the inner plane IE with the Rounding radius RA tangentially merges into the blade profile SP on the outside. FIG. 1d shows a view perpendicular to the profile surface in which the opening of the slot is visible in a corresponding projection and the transition of the fillet radius RA in the blade profile SP is indicated by a thin line. The side boundary SB is visible in this zone as an intersection, and is indicated by dashed lines in the area where it runs inside the bucket, as well as the, in this view, concealed feed cylinder. Here, the lateral boundary SB is semi-circular or elliptical, while Fig. Figure 1e shows a solution with even page boundaries. In Fig.2a, b, the fluidic function of the proposal according to the invention is demonstrated, which could already be proved by experiments in the manner shown as valid. 2a shows the outflow of the coolant KM from the cooling slot KS, as at the suction or pressure side of the profile substantially in the direction of the main flow (gas flow GS), takes place, wherein in the narrowest cross section the sound line SL is indicated by dashed lines, and the forming boundary line GL between the further expanding cooling film and the main flow GS is registered in dash-dotted lines, thin lines indicate the expansion fans and the dash lines, determine the outflow of the blade profile curvature SP following cooling film KF. In this way, it comes to the formation of a large surface supersonic cooling film KF. In Fig.2b the special situation of the flow around the profile nose PN (also referred to as the leading edge in turbine construction) with the cooling according to the invention by the cooling slot KSEK (cooling slot leading edge) is shown, where KM means the coolant again. The gas flow GS is here indicated by streamlines in the vicinity of the front stagnation point, and the cooling film overflowing the entire profile nose is denoted by KF as above.

The fluidic function of the invention will now be described with reference to these drawings. 1 a shows the outline of a gas turbine blade, FIG. 1b the elevation of such a blade with the Tannenbaumfuß TF, and the blade profile SP in which the mentioned cooling slots KS in several rows behind one another, but are arranged correspondingly offset and allow in a simple way by the large-scale formation of the desired cooling films to cover the entire surface of the blade , It is particularly pointed to the cooling slot on the profile nose KSEK, which causes a blow-out against the incoming gas flow, but by the property of the under-expanded jet this is able to wrap around the strongly curved leading edge and there to effect the most necessary cooling.

FIG. 1c now shows a section through the blade shell in a sectional plane approximately perpendicular to the radial, corresponding also to the section shown in the plan view of Fig.1a. In Fig.lc the mentioned details of the coolant flow KM are shown to the narrowest cross-section. From the feed cylinder SZ follows a limited by two levels cooling channel, which represents the actual cooling slot KS, this opens at its narrowest point to the surface, the lower contour of this cooling slot merges tangentially with the mentioned fillet radius RA in the blade profile SP.

This geometric arrangement of the cooling slot is shown in plan in Fig. 1d. We see from the outside on the gas turbine blade, therefore conceals the feed cylinder and the intersection of the two planes of the cooling slot with the feed cylinder, as well as the contour of their lateral boundary. An alternative of this arrangement is repeated again for another contour in FIG. 1 ©.

The advantages of the construction according to the invention are the following:

The pressure losses of the feed are significantly reduced because the highest speed occurs only in the outlet cross-section of the slot. There is thus a well-defined well-computable pressure in this outlet cross section and the outlet cross-section itself can be brought by the chosen manufacturing process and by any plastic deformation exactly to the desired shape.

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The coolant enters the main stream in a wide fan so that it can be circulated therefrom. The main flow therefore forces the coolant to abut against the surface, to flow along the intended rounding as a boundary layer over the surface and thus to form a cooling film with a strong cooling effect.

There are also advantages in terms of strength. For a cooled blade, the highest stresses in the cold zone occur at the inner edge of the bucket's bucket. In the conventional method with the numerous holes small diameter occur here very high notch effects at the mouth of the holes on the inside of the shell. In the construction according to the invention a cylinder with large radii of curvature is given here, whereby the notch effects are significantly reduced.

The cooling method according to the invention can be used both in systems with air-sucking gas turbines, as well as in plants that are operated in the so-called combined process, ie work with a gas turbine with downstream steam system. As a result of the usual pressure conditions in the gas turbine process, or the usual boiler pressure in such downstream systems here is a convenient way to make a feed into the gas turbine blades for cooling the same with steam of high pressure but low temperature, so that no special measures for producing the pressure ratio of 2 to 3 according to the invention are required. On the other hand, in air-sucking gas turbines, air can very well also be used for the cooling, since the first blade already has a sufficiently high pressure ratio if the feed is from the compressor and only for the first guide blade, if it is cooled in accordance with the invention, a relatively small additional compressor is required. Thus, the method according to the invention for air-sucking gas turbines can generally be used.

The advantages of the invention with respect to the achievable by the geometric arrangement and the high pressure ratio flow characteristics are shown in Fig.2. Thus, FIG. 2a shows the inflow of the coolant KM through the cooling slot KS in the section perpendicular to the blade axis as well as the boundary planes and the narrowest cross section in which the speed of sound (sound line SL) occurs, as well as the course of the under-expanded beam during application to the surface (cooling film KF) as a result of the then continuing expansion to supersonic speed. It is drawn in this picture, the boundary line GL between the free gas flow and the cooling film and it is noted that due to the different temperature of the coolant and the main flow and due to the high velocity of the cooling jet, that differs only slightly from that of the mainstream only a slight turbulent exchange takes place along that boundary line, so that over a wide range of cooling jet with a low temperature rests against the surface of the profile.

Fig.2b shows in a section the particularly difficult situation at the profile nose PN of such a gas turbine blade. At this occurs a stagnation point of the incoming gas flow (GS), which means that the boundary layer thickness is zero at this point, and on both sides forms a laminar, but very thin boundary layer. This thin boundary layer represents only a small thermal insulation compared to the high temperature of the main flow, so that normally takes place here the peak value of the heat transfer and the heating of the blade. The low-speed coolant streams produced here in general by roaring of cooling bores are blown off here particularly easily, so that a particularly serious problem of gas turbine cooling is to be solved here. The proposal according to the invention, however, is to arrange here a cooling slot, which lies on the pressure side of the profile, a supersonic cooling jet (cooling film KF), directed in part against the main flow, sends out. Due to its properties caused by the sub-expansion (sound velocity in the slot cross-section), this cooling film is also subject to the extreme curvature of the leading edge. It thus occurs here on a supersonic cooling jet, which flows around the entire profile nose (leading edge), which is also only slightly influenced by turbulence in the main flow.