Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means
Reexamination Certificate
1999-02-22
2001-02-06
Lopez, F. Daniel (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
With heating, cooling or thermal insulation means
C416S228000, C416S235000, C415S178000, C415S914000
Reexamination Certificate
active
06183197
ABSTRACT:
BACKGROUND OF THE INVENTION
This invention relates to reducing external heat load n an airfoil, either stationary, rotating, or their associated endwalls through use of heat reducing dimples located on the airfoil or its associated endwall.
Turbine airfoils are subject to extreme heat loads in high performance machines such as aircraft engines and power turbines. The temperature of hot gases entering the turbine can be well above the melting point temperatures of the alloys from which the airfoils are fabricated to such a degree that highly air-cooled airfoils are now common in the industry. Features that have the ability to reduce this heat load can directly benefit the performance of the turbines.
It is known in the industry to cool airfoils by flowing a cooling fluid through the hollow interior of the airfoil. The most common method is to bleed air from the compressor that is at a relatively lower temperature into the interior of the airfoil. Generally, the cooling is accomplished by external film cooling, and internal air impingement and convection cooling, or a combination of both. Air impingement cooling contemplates compressor bleed air channeled to the inside of the airfoil and directed onto the inside wall of the airfoil. The air then exits through a set of film cooling holes on the surface of the airfoil.
Film cooling has been shown to be very effective but requires a great deal of fluid flow that typically requires the use of power and is therefore looked upon as penalizing fuel efficiency and power. Also, film cooling is sometimes actively controlled, which is complex and expensive. Another disadvantage of film cooling is the degree of complexity in fabricating and machining the airfoils. The added complexity of film cooling results in more features that can break down while operating. Thus, film cooling of airfoils greatly increases the cost of operating.
As such, features and techniques that reduce the heat load on the airfoil surface and are relatively simple in nature, not requiring the complexity or power penalties of film cooling, are quite desirable. It was this desire which led to the so-called riblets on the leading edge described in U.S. Pat. No. 5,337,568 issued to Lee et. al. and titled Micro-Grooved Heat Transfer Wall. Riblets are a series of continuous grooves in a surface that are aligned in the direction of the flow, that serve to reduce the surface drag and the heat transfer when formed with the correct height, width, shape and spacing. Such riblets have been demonstrated to reduce drag on the fuselages of aircraft. However, for application to turbine airfoils, riblets are not as straightforward to use due to the extremely small size required for precision machining in the surfaces, and due to the fact that the riblet geometry would need to be altered for each region of the airfoil surface or each operating condition.
Accordingly, there is a need in the art for an airfoil where heat load is reduced passively over the entire airfoil surface and is also easy to fabricate or machine.
SUMMARY OF THE INVENTION
According to the present invention, an airfoil with a reduced heat load for use in either a turbine or a compressor through which a hot gas stream flow passes comprises a body having a leading edge and a trailing edge, the body having an exterior surface shaped from the leading edge to the trailing so as to have a suction area which has a convex shape. The suction area will experience a relatively low gas pressure as the hot gas stream flow passes thereover from the leading edge to the trailing edge. The body will be further shaped so as to have a pressure area from the leading edge to the trailing edge which will be concave in shape. The pressure area will experience a relatively high gas pressure as the hot gas stream flow passes thereover from the leading edge to the trailing edge.
The heat load is reduced on the airfoil by having at least one heat reducing dimple on the body of the airfoil. Each dimple has a length that is preferably aligned with the expected direction of hot gas stream flow. The width of a dimple is accordingly preferably aligned in a direction transverse to the expected direction of hot gas stream flow. To achieve the benefits of this invention, the length must be at least equal to or greater than the width.
The heat reducing dimple reduces the heat transfer coefficient for the airfoil downstream of its location. This serves to reduce the total heat load on the airfoil. Thus, the present invention is advantageous because the heat reducing dimple can be placed anywhere on the airfoil body. Furthermore, heat load can be reduced without the need for the expensive, power consuming film cooling mechanism of the prior art. Since the heat reducing dimples contemplated for use with this invention are relatively large, they are easy to machine into the surface of an airfoil.
The present invention also contemplates a method of reducing heat load on a turbine or compressor airfoil by forming at least one heat reducing dimple in the body of the airfoil. This method can be employed in existing airfoils to further reduce heat load or to save power for film cooled airfoils since less film cooling will be needed to accomplish the same reduction in heat load.
REFERENCES:
patent: 4522360 (1985-06-01), Barnwell et al.
patent: 4720239 (1988-01-01), Owczarek
patent: 4859150 (1989-08-01), Takigawa
patent: 4872484 (1989-10-01), Hickey
patent: 4974633 (1990-12-01), Hickey
patent: 5337568 (1994-08-01), Lee et al.
Abuaf Nesim
Bunker Ronald Scott
General Electric Company
Lopez F. Daniel
Nguyen Ninh
Patnode Patrick K.
Snyder Marvin
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