Fluid reaction surfaces (i.e. – impellers) – With heating – cooling or thermal insulation means – Changing state mass within or fluid flow through working...
Reexamination Certificate
2002-05-06
2004-09-07
Look, Edward K. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
With heating, cooling or thermal insulation means
Changing state mass within or fluid flow through working...
C416S19300A, C416S248000
Reexamination Certificate
active
06786696
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates generally to gas turbine engines, and, more specifically, to turbine blades therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. High and low pressure turbines include corresponding rows of turbine rotor blades extending radially outwardly from supporting rotor disks, with energy being extracted from the hot combustion gases by the rotor blades for rotating the disk which in turn is used for powering the compressor and an upstream fan in a typical turbofan aircraft engine application.
A primary design objective for aircraft engines is maximizing performance thereof while minimizing weight. Accordingly, the various components of the gas turbine engine are designed for removing as much weight as possible therefrom without exceeding acceptable stress limits therein.
Weight reduction is particularly more difficult in engine rotor components since they rotate during operation and must carry substantial centrifugal loads which generate corresponding stress therein. In addition to centrifugal loads, rotor blades are subject to aerodynamic or pressure loads due to the air being compressed or the combustion gases being expanded through the compressor and turbines.
Since a turbine rotor blade is subject to hot combustion gases during operation, it is typically cooled using air bled from the compressor. A typical turbine blade includes a hollow airfoil having various cooling circuits therein which are fed with cooling air obtained from the compressor which enters the turbine blade through inlet apertures extending radially through supporting dovetails of the blades.
The dovetail of a turbine blade typically includes corresponding pairs of upper and lower dovetail lobes or tangs in a fir tree configuration. The perimeter of the rotor disk includes a row of axial dovetail slots defined between corresponding disk posts having complementary upper and lower supporting lobes or tangs.
The disk slots are typically manufactured using a conventional broaching process in which a series of increasingly larger cutting tools are carried axially through the rotor perimeter until the final fir tree configuration of the disk slots is achieved.
The disk tangs are therefore axially straight between the forward and aft endfaces of the disk. And, the corresponding dovetail tangs are also axially straight for mating with the complementary disk tangs.
In this way, the two pairs of dovetail tangs provide four axially straight pressure surfaces which engage the corresponding pressure surfaces of the disk tangs for carrying centrifugal and other loads from each blade during operation into the perimeter of the disk which supports the blades.
In other designs, the dovetail slots may be skewed or angled through the disk rim relative to the engine centerline or axial axis, and the blade dovetails are correspondingly skewed or angled.
Since the dovetail tangs extend circumferentially oppositely from each other they define corresponding necks of locally minimum area directly above each of the lower and upper pairs of tangs. These dovetail necks must be sufficiently sized in area to spread the centrifugal loads thereacross for minimizing the maximum or peak stresses in the dovetail. The peak stress in the dovetail must be limited to ensure a suitable useful life of the blade in operation. Accordingly, the blade dovetails have minimum sizes controlled by a maximum acceptable peak stress therein.
In order to enhance the strength of the radially innermost or lower dovetail tang pair, the root end thereof may include a rectangular block of additional material extending the full axial length of the dovetail between the forward and aft endfaces thereof, and extending over the lateral or circumferential width of the dovetail between the base ends of the corresponding opposite tangs. The root block typically joins the lower tangs at corresponding fillets of suitable radius for reducing stress concentrations thereat.
The resulting blade dovetail is relatively complex in view of the fir tree configuration required therefor for transferring all operational loads from each blade into the rotor disk.
Reduction in the mass or weight of the individual blade dovetails is presently limited by the maximum acceptable peak stress therein. Since a typical gas turbine engine includes a substantial number of turbine blades in each stage row, it would be desirable to further reduce weight of the engine by correspondingly reducing weight of the dovetails, provided the acceptable peak dovetail stress is not exceeded.
Accordingly, it is desired to provide an improved dovetail having further weight reduction without exceeding maximum permissible peak stress therein.
BRIEF DESCRIPTION OF THE INVENTION
A turbine blade includes an airfoil and dovetail. The dovetail includes a pair of supporting tangs. A rectangular root block bridges the tangs over a majority of the root end thereof. The block terminates short of one endface of the dovetail to form a root notch thereat for preferentially reducing weight.
REFERENCES:
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Herman William Charles
Keith Brian David
Norton Brian Alan
Andes William S.
Conte Francis L.
McAleenan J. M.
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