DE10244199A1 - Device for supplying secondary fluid to transsonic primary flow e.g. for supplying cooling air for film cooling in turbine plant - Google Patents

Abstract

The device (3,12,13) supplies a secondary fluid (5) to the transsonic primary flow (4) via a bore (3) in a surface (15) over which the latter travels, the exit opening of the bore located in a recess (13) provided in this surface. An Independent claim for a method for supplying a secondary fluid to a transsonic primary flow is also included.

Description

TECHNICAL TERRITORY

The present invention relates to an apparatus and a method for introducing a secondary fluid by means of a hole in a transonic primary flow flowing over a surface. there it is particularly preferably a device, respectively a method for introducing cooling air into a transonic, via a surface flowing Hot gas flow in a turbine plant for film cooling of the surface.

STATE OF THE ART

The trend in industrial gas turbines to higher Turbine inlet temperatures, greater performance, smaller Emissions and reduced costs are unbroken. On the one hand on the material side, progress towards higher permissible material temperatures and more permissible Material tensions reached, on the other hand there is a need to the cooling for new ones Generations of gas turbines continue to improve.

Guide and rotor blades immediately downstream the combustion chambers are the greatest thermal stress exposed. The blades are generally on the hot gas side film-cooled. cooling air is suitably blown out through holes so that a protective Cooling air film the Shovel "shields" from the hot gas. These cooling air holes have cylindrical or diffuser shapes and are suitably arranged, to cool the film optimize and prevent, for example, dust entry. By the diffuser becomes the exit pulse current density of the cooling air of the hot gas so that the cooling film remains on the wall and evenly distributed across the current direction.

For example, B. the US 6,129,515 specific forms of openings in cooling air bores. Specifically, bores are proposed which expand conically before exiting into the hot air flow, ie shortly below the surface, so that the cooling air flow triggers as little interference as possible in the hot air flow and a homogeneous cooling air film is formed.

The power output of the turbine can increase by the pressure difference between the pressure side and the suction side a shovel raised will (with unchanged Gradually entering state). A reduction in costs is achieved through a smaller number of stages of the turbine possible. With the same power output have to the blades enlarged accordingly become. For both cases the load of a step is increased.

An increase in the pressure difference leads to higher flow velocities and Mach numbers (the Mach number is defined as the quotient of the speed through the speed of sound and describes the influence of compressibility). A limitation the pressure difference is necessary because the losses increase rapidly, if the local speed is the local speed of sound approaches.

This state is called transonic flow state designated. Usually the transonic flow state is defined as the flow state with a Mach number between 0.8 and 1.2.

Flow losses can occur in Pressure resistance, friction resistance and wave resistance divided (see e.g. Prandtl, Oswatitsch and Wieghardt "Guide through Fluid Dynamics", Vieweg, 9th ed. Braunschweig, 1993). Cause of are the rapidly increasing losses in the case of transonic flow conditions local shock waves, which increase the wave resistance to lead. Shock waves occur when the local Mach number exceeds one.

The transonic flow state is characterized by a relatively large dependency the current of the geometry and small fluctuations in the flow. For civil applications this condition is generally avoided.

If cooling air is blown out into transonic flows, this reduces the flow area for the hot gas flow and the flow rate and thus the mach number is increased and can reach the sound state. For modern gas turbines the local Mach number on the first blades immediately downstream after the combustion chamber nearby from one. Modern gas turbines are therefore operated at the limit. Out For this reason, the narrowest cross section of the first guide vane for example, generally no cooling air is blown out on the suction side, because otherwise such shock waves form and accordingly the flow resistance increases.

But there is a need for higher performance which z. B. by increasing the turbine inlet temperature, larger pressure differences between Pressure and suction side and larger Dimensions of the blades downstream of the narrowest cross section realized can be. This need may require cooling air to cool the film in transonic flow areas blow.

PRESENTATION THE INVENTION

The invention is therefore based on the object to provide a device or a method which involves the introduction of a secondary fluid, for. B. cooling air, by means of a hole, for. B. a cooling air hole, in a transonic, flowing over a surface primary flow, for. B. a hot air flow in a gas turbine ne, allowed, without having the disadvantages mentioned above.

This will solve this problem reached the exit opening the hole in a recess, or trough, arranged in the surface is.

The essence of the invention is with in other words, the flow cross section at the point where the secondary fluid is introduced into the primary flow is going to raise locally so that in the area of introducing the primary flow an essentially unchanged Flow cross section for disposal stands and so there are no shock waves in the area of insertion can. Downstream of the depression, when it again adjusts to the surface, has the secondary fluid already so widely distributed and homogeneously adapted to the primary current that shock waves can also be essentially prevented in this area. Corresponding allows the arrangement of such depressions in the area of Incorporation of a reduction in the otherwise occurring flow resistance. The present invention reduces the characteristic impedance of the z. B. Cooling air blowing into transonic flow areas arises. The invention is in no way based on the cooling air blowing limited to gas turbines, but it refers to all types of introduction of a secondary fluid into a transonic primary flow. The The invention described here minimizes the loss of air blowing in regions with a high external flow number arises from an increase in the wave resistance. This is achieved due to the geometric shape of the cooling air outlet geometry. The Invention can be achieved by optimizing the diffuser of the cooling air hole be combined appropriately.

The invention transfers Whitcombs "transonic area calls" to the blowing a secondary fluid into a primary fluid. By contouring the cooling air geometry and blade outer skin in the case of gas turbine blade application the resistance and thereby the losses are significantly reduced. The invention enables Blowout into transonic flow areas.

According to a first preferred embodiment According to the invention, the primary flow is a flow with a Mach number M in the range from 0.8 to 1.2, particularly preferred from 0.8 to less than 1. Especially in the transonic range below M = 1, the arrangement of depressions can Shock waves are prevented and the flow resistance be kept.

Another preferred embodiment of the Invention in which the introduction of the secondary fluid, d. H. z. B. the cooling air, particularly gentle in the primary current happens, is characterized in that the bore essentially is arranged in a plane which is perpendicular to the surface in the direction the primary flow. The secondary flow points the right moment in the direction of flow of the primary flow. there preferably closes the bore with the flow direction of the primary flow an angle between 0 and 90 degrees, particularly preferably between 30 and 60 degrees.

Another preferred embodiment draws is characterized by the deepening at its widest point in the direction of flow an area which is in the range of one tenth to ten times, preferably from one tenth to 1 times the flow cross section at the considered Position (without deepening). In order to influence the primary current as little as possible, it proves it is advantageous to deepen the longitudinal and / or transversely concave direction substantially semicircular. This should generally be taken into account when designing the specialization be sure not to form any sharp edges as these become stress zones to lead, in which slight cracks can occur under thermal stress. Furthermore to lead sharp edges often to stall and generally poor flow behavior.

The bore through which the secondary fluid is introduced into the primary flow can preferably have a widening in the area immediately before its opening to the surface of the depression. It can be widenings such as z. B. in the US 6,129,515 are described, but other diffuser shapes are also possible. When designing the bores or their widening and when designing the depression, it proves to be advantageous to dimension them as follows: the diameter of the bore relates to the diameter of the widening on the surface of the depression in a ratio between 1.2 and 1, 8 and / or the diameter of the widening on the surface of the depression relates to the diameter of the depression on the surface between 1.2 and 1.8. In particular in the area of the introduction of cooling air into a hot air stream in a gas turbine z. B. immediately downstream of the combustion chamber in connection with the film cooling of moving blades or guide vanes, it turns out that the bores are designed with a diameter in the range from 0.5 to 1.5 mm, particularly preferably in the range from 0.8 mm is suitable. Previous documents on cooling air hole optimization only include the aspect of optimizing the thermal effect of the cooling film, dust entry and manufacture of the complicated diffuser holes.

According to a further preferred embodiment of the invention, the primary flow is a hot gas flow in a turbine system, particularly preferably in a gas turbine system, the secondary fluid being cooling air which is introduced into the primary flow for film cooling. Acting it is preferred for the surface to be the surface of a component of the turbine system arranged essentially immediately downstream of the combustion chamber, such as, for. B. in particular around the surface of a moving blade, guide blade, and / or a housing component.

Further preferred embodiments the inventive Device are in the dependent claims described.

Furthermore, the present concerns Invention a method for introducing a secondary fluid by means of a bore in a transonic primary flow flowing over a surface, in particular for the introduction of cooling air into a transonic, about a surface flowing hot gas flow in a turbine system for film cooling. The secondary fluid flows through a hole introduced the outlet opening is arranged in a recess in the surface. Prefers is such a method using a device as described above. The procedure can be done with great advantage if the Mach number of the transonic primary flow in Range moved from 0.8 to 1.

SHORT EXPLANATION THE FIGURES

The invention is based on the following of embodiments are explained in more detail in connection with the drawings. Show it:

1 a section in the flow direction through a surface flowed over by a transonic hot gas with a cooling air bore;

2 a) a transsonically approached aircraft without rejuvenation; b) a transsonically approached aircraft with a taper; c) resistance as a function of Mach number; and

3 a) a section in the flow direction according to 1 with a cooling air hole in a concave recess; b) a cut according to 3a ) but perpendicular to the flow direction of the hot gas; c) a top view of a hole with a recess.

WAYS TO EXECUTE THE INVENTION

1 shows a section through a turbine blade 1 perpendicular to the plane or surface 15 of the turbine blade 1 and parallel to the direction of flow of a hot air flow 4 (Primary flow) according to the prior art. The cut leads through a cooling air hole 3 through which cooling air 5 in the hot air flow 4 is blown in. The cooling air 5 serves the downstream of the cooling air hole 3 arranged area with a cooling air film 7 to act upon.

Is it the hot air flow 4 a transonic flow, i.e. a Mach number M of between 0.8 and 1.2, is formed in the cooling air bore 3 a range of shock waves 6 out. The reduction of the hot air flow 4 available cross section through the introduction of cooling air 5 leads to this formation of shock waves 6 ,

It serves to illustrate the principle according to the invention 2 , Airplanes that reach transonic flight speeds have a continuous taper in the area of the wing attachment, as a result of which a smooth transition in the flow direction of the cross-sectional area is achieved and the maximum cross-section that the flow sees is reduced. This taper provides additional space for the flow at the wing attachment, which is accelerated due to the finite thickness of the wings. This measure reduces the aircraft's resistance at speeds close to sound by up to 75%. The idea goes back to R. Whitcomb (see for example JD Anderson, "Modern compressible flow", Mc Graw Hill, 2nd Edition, 19990, USA). 2a schematically shows an airplane 9 with wings 10 , which flies in the area close to the sound. The inflow 8th , z. B. already slightly above Mach 0.8 accelerated to the speed of sound by the aircraft as a result of displacement by the fuselage at the fuselage surface. In the wing area, the speed of the flow is increased even further, whereby the resistance also increases steeply. This behavior is shown in the graphic 2c ), where the curve W shows the resistance W of an aircraft without tapering in the wing area. Above a critical value of the Mach number (M _crit , typically 0.8), the resistance suddenly increases massively. 2 B ) now shows an airplane, where tapered 11 in the area of the wing 10 available. If one considers the resistance as a function of the Mach number for such a construction, the behavior according to curve b in 2c ). The rejuvenation 11 leads to a strong reduction R of the resistance above the critical value of the Mach number M _crit . This behavior is known as the area rule.

This principle, which has not yet been used in connection with the introduction of a secondary fluid in a primary flow, in particular in connection with the cooling of turbine components, is to be used according to the invention as follows. 3a ) shows a section perpendicular to the surface 15 of a turbine blade 1 in a gas turbine. The surface 15 is thereby from a hot air flow 4 swept over, the hot air flow flowing in the transonic area. The cooling air 5 for the formation of the cooling air film 7 is no longer directly on the surface here 15 blown in like in 1 shown. The cooling air duct 3 does not open directly on the surface 15 in the hot air flow 4 , but ends up in the surface 15 into a depression 16 , The hole 3 expands already shortly before the mouth into the depression 16 in an expansion 12 (Diffuser). This expansion improves the introduction of the cooling air 5 in the hot air flow 4 , The cooling air duct 3 lies essentially in a plane which is perpendicular to the surface 15 through the flow direction of the hot air flow 4 is spanned. The cooling air duct 3 is at an angle α to the plane of the surface 15 bent. This angle can be in the range of close to 0 to 90 degrees, typically angles of between 30 to 60 degrees are sensible, the inflow of cooling air should be shallower the higher the speed of the hot air flow. In this way, unnecessary asymmetrical lateral and orthogonal components of the cooling air become 5 avoided. 3b ) shows a section perpendicular to the flow direction of the transonic hot gas 4 , The circles with a point mean a direction of flow from the paper plane towards the viewer. 3c ) shows a top view of such a device for introducing cooling air 5 into a hot air stream 4 ,

In this case, the depression is both in the longitudinal direction of the hot gas flow 4 as a semicircular, conical depression 13 trained, as well as in the transverse direction perpendicular to the hot gas flow 4 as indicated by the reference symbol 14 is indicated. The depression points, as in 3c ) has an elliptical shape that is elongated in the direction of flow, but it can also take on a different shape, for example a circular shape. The cooling air hole 3 can be in the center and / or at the deepest point of the depression 16 be arranged, but it can also, for. B. be slightly offset upstream. The depression does not have to be semicircular concave, but can take special concave contours, depending on the flow requirements.

In the area of film cooling of rotor blades, the cooling air holes have 3 typically a diameter of approximately 0.8 millimeters. The exit diameter of the expansion 12 is typically 2 to 8 times larger than this diameter of the bore 3 , The outside diameter of the recess 16 is again 2 to 8 times larger than the outlet-side diameter of the expansion 12 ,

Typically, the depression has a depth T, so that the additional flow cross-section available as a result 0 . 1 to 10 times, preferably 0.1 to 1 times the flow cross section, which is present without a recess.

For the application of the invention is the flow cross-section that the cooling air 5 to determine swept out of the mass balance. The flow cross-section is the decisive parameter for the geometry changes to be made. Subsequently, several measures are used to reduce the flow cross-section for the hot gas flow 4 achieved continuously in the direction of flow. Abrupt changes should be avoided. The measures include:

• 1. Longitudinal trough in the outer skin of the blade ( 3a )
• 2. Transversal trough in the outer skin of the flow ( 3b )
• 3. Starting the cooling air hole in the radial direction ( 3a ), Angle α)

These measures are local measures that take place in the vicinity of the cooling air hole 3 can be carried out. This can be done, for example, by modifying the mold. Far away, downstream of the hole 3 , the resistance can be reduced by slightly withdrawing the blade contour. The local flow cross-section, which is the cooling air, is decisive 5 occupies. The geometry modifications 1 , -3 , can with diffuser shapes 12 be combined appropriately.

1

turbine blade

2

cavity

3

Cooling air hole

4

Hot air flow (primary flow)

5

Cooling air flow (secondary flow)

6

Area the shock waves

7

Cooling air film

8th

inflow

9

plane

10

Hydrofoil

11

Rejuvenation in Area of the wing

12

widening the cooling air hole

13

conical Cutout in the cooling air hole in the longitudinal direction

14

conical Cutout in the cooling air hole in the transverse direction

15

Surface of 1

16

deepening

W

resistance

M

Mach number

M _crit ,

critical Mach number

a

plane without rejuvenation in the wing area

b

plane with rejuvenation in the wing area

R

reduction resistance due to rejuvenation

α

tilt angle the cooling air hole in the direction of flow the hot air flow

T

depth the recess 16

Claims (11)

Contraption ( 3 . 12 . 13 ) for the introduction of a secondary fluid ( 5 ) by means of a hole ( 3 . 12 ) in a transonic, over a surface ( 15 ) stream primary flow ( 4 ) characterized in that the outlet opening of the bore ( 3 . 12 ) in a recess ( 13 ) in the surface ( 15 ) is arranged. Device according to claim 1, characterized in that the primary flow ( 4 ) has a Mach number (M) in the range from 0.8 to 1.2, in particular from 0.8 to less than 1. Device according to one of claims 1 or 2, characterized in that the bore ( 3 . 12 ) is essentially arranged in a plane that is perpendicular to the surface ( 15 ) in the direction of the primary flow ( 4 ), the bore ( 3 . 12 ) with the flow direction of the primary flow ( 4 ) includes an angle between 0 and 90 degrees, particularly preferably between 30 and 60 degrees. Device according to one of the preceding claims, characterized in that the recess ( 13 ) has a depth (T) so that the through the recess ( 13 ) additionally available cross-sectional area 0.1 to 10 times, preferably 0.1 to 1 times that without recess ( 13 ) corresponds to the present cross-sectional area, the depression ( 13 ) is preferably concave essentially semicircular in the longitudinal and / or transverse direction. Device according to one of the preceding claims, characterized in that the bore ( 3 ) in the area immediately before its opening to the surface of the depression ( 13 ) an expansion ( 12 ) having. Device according to claim 5, characterized in that the diameter of the bore ( 3 ) the diameter of the expansion ( 12 ) on the surface of the depression ( 13 ) between 1.2 and 1.8, and / or that the diameter of the expansion ( 12 ) on the surface of the depression ( 13 ) the diameter of the depression on the surface ( 15 ) between 1: 2 and 1: 8. Device according to one of claims 5 or 6, characterized in that the bore ( 3 ) has a diameter in the range from 0.5 to 1.5 mm, in particular in the range from 0.8 mm, Device according to one of the preceding claims, characterized in that the primary flow ( 4 ) is a hot gas flow in a turbine system, particularly preferably in a gas turbine system, and that the secondary fluid ( 5 ) is cooling air, which is used for film cooling ( 7 ) into the primary flow ( 4 ) is introduced. Device according to claim 8, characterized in that the surface ( 15 ) is the surface of a component of the turbine system arranged substantially immediately downstream of the combustion chamber, such as, for. B. especially around the surface ( 15 ) a moving blade, guide blade, and / or a housing component. Process for introducing a secondary fluid ( 5 ) by means of a hole ( 3 ) in a transonic, over a surface ( 15 ) flowing primary flow ( 4 ), especially for the introduction of cooling air ( 5 ) in a transonic, over a surface ( 15 ) flowing hot gas flow ( 4 ) in a turbine plant for film cooling ( 7 ), characterized in that the secondary fluid ( 5 ) through a hole ( 5 ) is introduced, the outlet opening of which is in a recess ( 13 ) in the surface ( 15 ) is arranged. Procedure according to Claim 10, characterized in that a device according to one of the Expectations 1 to 9 is used.