BACKGROUND ART Gas turbines comprise vanes and blades that are exposed to high temperatures, and for this reason they must be intensively cooled. Cooling is realised by the flow of the cooling medium, mostly air, within turbine blade passages. Because of the shape and dimensions of coolant passages desirable aerodynamical shape and proportions of blades cannot be realised, and the width of the channels between blades is decreased. Cooling intensity could grow by narrowing of the coolant passages, but this would be accompanied by the increase of the flow resistance, and the higher inlet pressure of the cooling air would be required. Increasing of the cooling air compression ratio and/or its consumption decreases the overall turbine efficiency. When the cooling air flows through passages within the blades, the unfavourable distribution of the local heat transfer coefficients is pronounced. The incompatibility of the cooling intensity within cooling air flow passages with the distribution of the thermal load on the gas side causes big nonuniformities of the temperature distribution within the blade material. The biggest thermal loads result near the leading edge and near the trailing edge. As the coolant passages near the trailing edge are the narrowest, it is very difficult to realise the cooling of this area. Near the leading edge the cooling passages are wider, but the cooling of this area is also a problem because of high thermal load concentration. The adverse shape of the coolant passages is the main difficulty in realising of the desirable local heat transfer coefficients. The nonuniformity of the temperature distribution within the blade material causes thermal stresses, shortening the life of the blades. Very unfavourable conditions arise during the turbine acceleration period, when the leading and trailing edges are heated faster than the middle part of the airfoil, and during deceleration period, when! the leading and trailing edges are cooled faster. Power load changes are typical for aircraft turbines.

Turbine blade designers try to minimise nonuniformity of the blade temperature. The aerodynamic losses of the gas flow between the blades should be as low as possible, and the blade cooling improvement should not decrease the overall efficiency of the turbine.

It is difficult to describe in detail all the advantages and disadvantages of a huge number of various turbine blade cooling methods. Generally, the aims are as follows: low temperature of the blade material low temperature gradients within the blade material 'low consumption of the cooling air low compression ratio of the cooling air * high total efficiency of the turbine small weight of the blade simple and inexpensive blade manufacturing The above mentioned criteria have to be satisfied by the blade design. The intention is to increase cooled surface area within passages and to increase the heat transfer coefficient to the air, with small cooling air consumption and flow resistance. There are many ways to enhance the heat transfer to the air, and the maximum enhancement has to be realised where the thermal load is high. On the locations where thermal load reaches its highest value, e. g. on the leading edge of the blade, the cooling air is released through an array of holes into the boundary layer of the gas flowing near the airfoil surface. By such method the local cooled area is increased and the injected air forms a cooling film near the leading edge. But, in spite of the high drilling technology, the stress concentration near the holes reduces the life of the blade. At the trailing edge region the cooling air flows circumferentially out of the blade through the narrow slots, because it is not possible to place wide radially oriented passages within the thin trailing edge, that would insure the necessary air flow rate.

Within cooling passages, the heat transfer can be enhanced by the means of flow turbulators, and/or deflectors that locally direct air flow towards the surfaces with higher thermal loads. When compared to the radial air flow direction, the circumferential air flow direction has the advantage in reduced heating of the cooling air due to the compression in rotating blades at high angular velocities. By artificial roughening of the surfaces within cooling passages which are exposed to high thermal loads, both the heat transfer coefficient and the heat transferring area are increased.

But there are cases of rotating blades where the transversal ribs cause so high stresses due to their inertia reaction to the centripetal forces so that this negative effect is higher than the positive effect of the enhanced cooling.

The technology of blade manufacturing is a very important criterion in choosing the proper blade cooling method. When the blades are produced from two parts which are bonded by welding or soldering, it is possible to realise much higher precision of the dimensions of cooling passages, the blade wall thicknesses and the final dimensions of the blade. In that case, the artificial roughening of the surfaces in coolant passages can be done much easier as well. The blades are usually produced from nickel based alloys. The thermal conductivity of those alloys is relatively low, app.

20 W/(m K). Por this reason, it is desirable to make the blade walls as thin as possible in order to avoid to high temperature gradients in transversal direction. Sometimes, the cooling of the leading and trailing edges is enhanced by the use of material with higher thermal conductivity, e. g. nickel-aluminide. Material with higher thermal conductivity distributes the heat flux also in longitudinal direction, thus reducing the heat flux concentration at a small surface. In this case, the blade should be made of three parts. The leading and the trailing parts are made from nickel- aluminide, and are cemented to the central part which is made from nickel based superalloy. All three parts have separate coolant passages. In that case the overall weight of the blade is also reduced because nickel-aluminide has lower density than the superalloy. Unsatisfactory is the necessity to make a durable and reliable bonding of three blade parts.

DETAILED DESCRIPTION OF THE INVENTION The cooling of turbine blades is improved by present invention with decreasing temperature gradients in the blade walls by means of a thin, graphite foil with anisotropic thermal conductivity, which is placed on surfaces of coolant passages and/or on the gas side surfaces. The essential properties of graphite foil are also its low density, softness and flexibility, that enable a simple realisation of a tight contact between the heat transferring surfaces and the foil. According to the data of a producer of such foils, the thermal conductivity in the longitudinal direction is for temperatures up to 800 ° C higher than 70 Wl (m K), while for lower temperatures its value increases and reaches values that are higher than 100 W/(m K). In the direction perpendicular to the foil surface the thermal conductivity is substantially lower, only app. 4 Wl (m K). The graphite foil must be covered by the metal foil, that serves to protect the graphite foil from oxidation and to give strength to it. Thanks to the softness and the flexibility of the graphite foil, its tight contact to the blade wall on one side and to the protecting metal foil on the other side can be realised by evacuating the air and by the gas tight welding or soldering the circumference of the metal foil with the blade alloy substrate. This enables an undisturbed heat flow between the blade wall and graphite foil, as well as between the graphite foil and the protecting metal foil. If necessary, in rotating blades, the bonding of the graphite foil with the blade wall and with the metal foil could be enhanced by means of a very thin layer of thermally conducting cement, in order to avoid bending of the graphite foil due to inertia forces. For the case of very high loads it is possible to apply the graphite foil which is reinforced by metallic fibers or mesh, which could be produced specially for this purpose.

The graphite foil could be applied in many existing designs, but also the new coolant passages could be designed for the optimal exploiting of the advantages of this method. Because of the decreasing of the cross-sectional area of coolant passages, due to the thickness of the graphite and metal foils, the cooling air flow rate should also be reduced in order to avoid the pressure drop increasing. The resulting change of the heat transfer coefficient can be neglected. The increase of the mean temperature of the blade walls due to the additional resistance to the heat flux caused by the graphite foil will be compensated by the higher uniformity of temperature distribution, and thus also with lower thermal stresses. In the overall thermal resistance to the heat flux through the blade wall, the resistance caused by the boundary layer of the cooling air is several times higher than the corresponding resistance to the transversal conduction within the blade wall. For this reason, the additional resistance to the heat conduction, caused by the graphite foil, is relatively small. If the graphite foil and the protecting metal foil were placed on the gas side on places where the thermal load reaches its peak values, the mean temperature of the blade wall would decrease, and the temperature distribution uniformity would increase. Estimating of the optimal thickness of the graphite foil is a very complex problem, and for this purpose it is important to know the local heat transfer coefficients within coolant passages of different geometries. When the high thermal load is exposed to the part of the surface of the coolant passage with unfavourable geometry the graphite foil directs the part of the heat flux to the area with a higher heat transfer coefficient. Thus the graphite foil could compensate the nonuniformities of the heat transfer on both the gas side and the cooling air side.

The increase in the blade mass due to the application of the graphite foil could be neglected.

Graphite foil density ranges from 700 to 1000 kglm', and this is about ten times less than the density of the nickel based superalloys. But in the rotating blades the inertial forces would cause high stresses within the graphite foil. This problem could be solved by the use of artificially roughened surfaces, and the graphite foil with protecting metal foil could follow the shape of the artificial roughness. The roughness could be two-dimensional (ribs transversal to the flow direction) or three-dimensional. Three-dimensional surface roughness results with higher heat transfer enhancements. In that case the surface protrudings could pass through the graphite foil and be bonded with protecting metal foil, preventing foil defonnations due to inertial forces.

There are many different possibilities for the fastening of the graphite foil. For example, it is possible to make a thin metallic mesh (or a grid) and the graphite foil, which should be thinner than the mesh, could be placed within the mesh openings, so that the mesh could also serve as an artificial roughness for heat transfer enhancement. The mesh could be welded or soldered to the surface coolant passages. The artificially roughened wall surface could also include the parts of the graphite foil between the protrudings. There are also many technological possibilities of producing thin coatings on the graphite foil surface which could give the possibility to apply graphite foil without protecting metal foil. Such methods would be applicable for the stationary vanes, but also for the rotating blades when combining the metallic mesh with parts of the graphite foil embeded within the mesh. Optimal solutions, which would define the appropriate geometry, the height, and the arrangement of the surface roughness, as well as foil thicknesses, could be reached only by the use of complex engineering methods, which should include numerical calculations and experimental research.

The thickness of the graphite foil could be minimised if the protecting metal foil is made of the material of high thermal conductivity, e. g. copper or, at lower temperatures, also aluminum. This metal foil would additionally direct the heat flux parallel to the walls of coolant passages. As the thermal conductivity of copper is app. four to five times higher than the longitudinal thermal conductivity of the graphite foil, and app. ten times higher than the thermal conductivity of the nickel based superalloy, good results could be obtained with a very thin graphite foil. This solution could be favourable in the trailing edge area, where the coolant passages are very narrow.

As the graphite foil has a very low transversal thermal conductivity, it would stimulate the longitudinal heat flux redistribution within the conductive metal foil. The roll of the graphite foil is also to conduct the heat in longitudinal direction and to serve for the perfect thermal contact between the blade wall and the conductive metal foil. The advantage of the tight thermal contact that can be realised by the use of graphite foil is also in its simple realisation. The pressing of the parts together, evacuating the air, and the subsequent gas tight welding of the metal foil on its circumference with the blade material, do not necessitate very high technologies.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a cross section of a turbine blade, with the graphite foil 2 placed on the inside surface of the turbine blades walls 1, in coolant passages, and with the protecting metal foil 3, that is placed over the graphite foil 2.

FIG. 2 is a detail of a turbine blade wall 1, with the metallic layer 5, characterised by the high thermal conductivity, placed on the surface of a coolant passage, between the graphite foil 2 and the turbine blade wall 1, and with the protecting metal foil 3, that is placed over the graphite foil 2.

FIG. 3 is a detail of a turbine blade wall 1, with the protrudings 4 of the three-dimensionally roughened surface in coolant passages, that are covered by the graphite foil 2 and protecting metal foil 3.

FIG. 4 is a detail of a turbine blade wall 1, with the three-dimensionally roughened surface in coolant passages, covered by the graphite foil 2 and protecting metal foil 3, where the protrudings 4 of the artificial roughness are passing through the holes in the graphite foil 2.

FIG. 5 is a detail of a turbine blade wall 1, with the two-dimensionally roughened surface of coolant passages covered by the graphite foil 2 and the protecting metal foil 3, where the protrudings 4 of the artificial roughness are passing through the holes in the graphite foil 2.

FIG. 6 is a detail of a turbine blade near the leading edge, where the graphite foil 2 and the protecting metal foil 6 are placed on the gas side surface of the turbine blade wall 1, where the thermal load reaches its highest value. The graphite foil 2 and the protecting metal foil 3 are placed on the surface of the coolant passage in the middle part of the turbine blade, where thermal load is lower.

DESCRIPTION OF THE REALISATION OF THE INVENTION Figure 1 shows the typical shape of coolant passages in a turbine blade, with the graphite foil 2 and the protecting metal foil 3 on the surface. Technology details that have to be used for applying the graphite foil and the protecting metal foil on the walls of coolant passages are not the subject of this invention. Generally, for this purpose traditional techniques in the manufacturing of turbine blades could be applied, such as casting, forging, welding, soldering, cementing etc.

Satisfactory contact heat transfer could be realised by evacuating the air between the blade wall, the graphite foil and the protecting metal foil, with additional pressing if necessary, and finally with gas tight welding or soldering the circumference of the metal foil with the blade wall. There is also the possibility to use cement with high thermal conductivity. Protection of the graphite foil from oxidation at temperatures above 500 ° C could also be realised as an integral part of its surface. In that case, the protecting metal foil would not be necessary.

In stationary vanes the satisfactory contact could be realised by evacuating the air between the graphite foil and metal foil. The necessary pressure on the outside surface of the metal foil could be obtained by the compressed cooling air. In the rotating blades, with high centripetal accelerations, the relatively low tensile strength of the graphite foil, app. 4 N/mm, should be taken into account because the foil could not withstand bending. For this reason long radial graphite foil parts should be avoided, if the foil is not cemented. The length of the graphite foil can be decreased by the artificial roughness of coolant passages, which could be two-or three- dimensional. The literature shows that the greatest heat transfer enhancements can be obtained with the three-dimensional surface roughness. Also, for the same heat transfer enhancement, the flow resistance is lower in channels with artificial roughness having rounded protrudings. Figure 3 shows an example of the three-dimensional surface roughness with rounded protrudings. The blade wall 1 is in tight contact with the graphite foil 2, covered by protecting metal foil 3.

Protrudings 4 are in a"chess"arrangement. In this case, for the rotating blades the contact of the graphite foil with the blade wall and metal foil should be intensified by cementing, at least at protrudings. In the example shown on Figure 4 protrudings pass through the holes in the graphite foil and metal foil and are exceeding the surface of the metal foil. In that case, the gas tight connection must be ensured between all protrudings and the metal foil. Similarly, Figure 5 shows the case with two-dimensional artificial roughness with ribs exceeding the surface of the metal foil. Figure 2 shows the detail with the wall 1 of the turbine blade in the case when the thin ply 5 of the metal with high thermal conductivity, such as copper or aluminum, is placed between the graphite foil 2 and the surface of the blade wall 1. The metal ply 5 could be applied directly to the surfaces of coolant passages, e. g. by galvanic means. The tight contact between the surfaces of all the plies is an imperative independently to the technology applied. If the graphite foil should be cemented, the cement should have high thermal conductivity and it should be as thin as possible.

Figure 6 shows the detail of the blade near the leading edge. The graphite foil 2, with the protecting metal foil 6 is applied on the gas side of the blade, but the coolant passage below it can be without the graphite foil. The metal foil 6 must have extraordinary resistance to high temperatures, and therefore differs from the metal foil 3 which is applied within coolant passages and is exposed to substantially lower temperatures. Coolant passages in the middle of the same blade can have the graphite foil applied on their surface.

THE USE OF THE INVENTION This invention is primarily aimed to improve gas turbine blade cooling. The blades of the aircraft gas turbines with improved temperature uniformity would be less sensitive to sudden rise or decrease of turbine power, and thus also longer life.

CLAIMS What is claimed is: 1. A gas turbine blade, characterised by, that the surface of coolant passages within the said blade or only one part of the surface of the said coolant passages is covered by a thin ply of graphite foil with anisotropic thermal conductivity, the thermal conductivity of the said graphite foil being bigger in the direction parallel to the surface of the said coolant passages than in the direction perpendicular to the surface of the said coolant passages, the said graphite foil being bonded with the surface of the said coolant passages in a way, that the bonding is realised by means of covering the said graphite foil by a metal foil, by the application of pressure and/or by evacuating the air between the said metal foil, the said graphite foil and the surface of the said coolant passages, and with the subsequent gas tight bonding of the said metal foil with the wall of a turbine blade, so that the said metal foil protects the said graphite foil from oxidation at high temperatures.

2. A gas turbine blade in accordance with claim 1, characterised by, that the said graphite foil is reinforced by the fibers, or by the mesh, that is made of strong and thermally conductive material.

3. A gas turbine blade in accordance with claim 1, or claim 2, characterised by, that the said graphite foil and the said metal foil are applied on the surface or only a part of the surface on the gas side of the turbine blade.

4. A gas turbine blade in accordance with claim 1, or claim 2, or claim 3, characterised by, that the said bonding of the said graphite foil with the wall surface of the said turbine blade is realised by means of a cement with high thermal conductivity, resistant to high temperatures.

5. A gas turbine blade in accordance with claim 1, claim 2, or claim 4, characterised by, that the said graphite foil is not covered by the said metal foil, but the surface of the said graphite foil containes a thin overlayer that protects the said graphite foil from oxidation.

6. A gas turbine blade in accordance with claim 1, or claim 2, or claim 4, or claim 5, characterised by, that the surface of the said coolant passages is artificially roughened, and the said graphite foil, thanks to its softness and flexibility, follows the shape of the surface of the said coolant passages, and therefore the said artificial roughness has two benefits, firstly, the enhancement of the heat transfer in both rotating and stationary blades, and secondly, preventing the said graphite foil from bending due to the inertial forces in rotating blades.

7. A gas turbine blade in accordance with claim 6, characterised by, that the elements of the said artificial roughness of the coolant passages wall surface are shaped as protrudings passing through the holes in the said graphite foil, the geometry and dimensions of the said holes being adjusted to the said protrudings, wherein the protection of the said graphite foil from oxidation can be realised in accordance with the method described in claim 5, or by the use of the metal foil in accordance with claim 1.