Swept turbomachinery blade
A swept turbomachinery blade for use in a cascade of such blades is disclosed. The blade (12) has an airfoil (22) uniquely swept so that an endwall shock (64) of limited radial extent and a passage shock (66) are coincident and a working medium (48) flowing through interblade passages (50) is subjected to a single coincident shock rather than the individual shocks. In one embodiment of the invention the forwardmost extremity of the airfoil defines an inner transition point (40) located at an inner transition radius rt-inner. The sweep angle of the airfoil is nondecreasing with increasing radius from the inner transition radius to an outer transition radius rt-outer, radially inward of the airfoil tip (26), and is nonincreasing with increasing radius between the outer transition radius and the airfoil tip.
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This is a continuation of application Ser. No. 09/343,736, filed Jun. 30, 1999 now U.S. Pat. No. Re. 38,040, seeking reissue of U.S. Pat. No. 5,642,985, issued Jul. 1, 1997.
STATEMENT REGARDING GOVERNMENT RIGHTSThe government has certain rights to this invention under Department of Defense Contract No. N00140-91-C-2793.
TECHNICAL FIELDThis invention relates to turbomachinery blades, and particularly to blades whose airfoils are swept to minimize the adverse effects of supersonic flow of a working medium over the airfoil surfaces.
BACKGROUND OF THE INVENTIONGas turbine engines employ cascades of blades to exchange energy with a compressible working medium gas that flows axially through the engine. Each blade in the cascade has an attachment which engages a slot in a rotatable hub so that the blades extend radially outward from the hub. Each blade has a radially extending airfoil, and each airfoil cooperates with the airfoils of the neighboring blades to define a series of interblade flow passages through the cascade. The radially outer boundary of the flow passages is formed by a case which circumscribes the airfoil tips. The radially inner boundary of the passages is formed by abutting platforms which extend circumferentially from each blade.
During engine operation the hub, and therefore the blades attached thereto, rotate about a longitudinally extending rotational axis. The velocity of the working medium relative to the blades increases with increasing radius. Accordingly, it is not uncommon for the airfoil leading edges to be swept forward or swept back to mitigate the adverse aerodynamic effects associated with the compressibility of the working medium at high velocities.
One disadvantage of a swept blade results from pressure waves which extend along the span of each airfoil suction surface and reflect off the surrounding case. Because the airfoil is swept, both the incident waves and the reflected waves are oblique to the case. The reflected waves interact with the incident waves and coalesce into a planar aerodynamic shock which extends across the interblade flow channel between neighboring airfoils. These “endwall shocks” extend radially inward a limited distance from the case. In addition, the compressibility of the working medium causes a passage shock, which is unrelated to the above described endwall shock, to extend across the passage from the leading edge of each blade to the suction surface of the adjacent blade. As a result, the working medium gas flowing into the channels encounters multiple shocks and experiences unrecoverable losses in velocity and total pressure, both of which degrade the engine's efficiency. What is needed is a turbomachinery blade whose airfoil is swept to mitigate the effects of working medium compressibility while also avoiding the adverse influences of multiple shocks.
DISCLOSURE OF THE INVENTIONIt is therefore an object of the invention to minimize the aerodynamic losses and efficiency degradation associated with endwall shocks by limiting the number of shocks in each interblade passage.
According to the invention, a blade for a blade cascade has an airfoil which is swept over at least a portion of its span, and the section of the airfoil radially coextensive with the endwall shock intercepts the endwall shock extending from the neighboring airfoil so that the endwall shock and the passage shock are coincident.
In one embodiment the axially forwardmost extremity of the airfoil's leading edge defines an inner transition point located at an inner transition radius radially inward of the airfoil tip. An outer transition point is located at an outer transition radius radially intermediate the inner transition radius and the airfoil tip. The outer transition radius and the tip bound a blade tip region while the inner and outer transition radii bound an intermediate region. The leading edge is swept at a first sweep angle in the intermediate region and is swept at a second sweep angle over at least a portion of the tip region. The first sweep angle is generally nondecreasing with increasing radius and the second sweep angle is generally non-increasing with increasing radius.
The invention has the advantage of limiting the number of shocks in each interblade passage so that engine efficiency is maximized.
Referring to
The hub 16 is attached to a shaft 52. During engine operation, a turbine (not shown) rotates the shaft, and therefore the hub and the blades, about the axis 18 in direction R. Each blade, therefore, has a leading neighbor which precedes it and a trailing neighbor which follows it during rotation of the blades about the rotational axis.
The axial velocity Vx (
Sweeping the blade leading edge, while useful for minimizing the adverse effects of supersonic working medium velocity, has the undesirable side effect of creating an endwall reflection shock. The flow of the working medium over the blade suction surface generates pressure waves 60 (shown only in
The endwall shock can be eliminated by making the case wall perpendicular to the incident expansion waves so that the incident waves coincide with their reflections. However other design considerations, such as constraints on the flowpath area and limitations on the case construction, may make this option unattractive or unavailable. In circumstances where the endwall shock cannot be eliminated, it is desirable for the endwall shock to coincide with the passage shock since the aerodynamic penalty of coincident shocks is less than that of multiple individual shocks.
According to the present invention, coincidence of the endwall shock and the passage shock is achieved by uniquely shaping the airfoil so that the airfoil intercepts the endwall shock extending from the airfoil's leading neighbor and results in coincidence between the endwall shock and the passage shock.
A swept back airfoil according to the present invention has a leading edge 28, a trailing edge 30, a root 24 and a tip 26 located at a tip radius rtip. An inner transition point 40 located at an inner transition radius rt-inner is the axially forwardmost point on the leading edge. The leading edge of the airfoil is swept back by a radially varying first sweep angle σ1 in an intermediate region 70 of the airfoil (in
The leading edge 28 of the airfoil is also swept back by a radially varying second sweep angle σ2 in a tip region 74 of the airfoil. The tip region is radially bounded by the outer transition radius rt-outer and a tip radius rtip. The second sweep angle is nonincreasing (decreases, or at least does not increase) with increasing radius. This is in sharp contrast to the prior art airfoil 22′ whose sweep angle increases with increasing radius radially outward of the inner transition radius.
The beneficial effect of the invention is appreciated primarily by reference to
The embodiment of
The invention's beneficial effects also apply to a blade having a forward swept airfoil. Referring to
The leading edge 128 of the airfoil is also swept forward by a radially varying second sweep angle σ2 in a tip region 74 of the airfoil. The tip region is radially bounded by the outer transition radius rt-outer and the tip radius rtip. The second sweep angle is nonincreasing (decreases, or at least does not increase) with increasing radius. This is in sharp contrast to the prior art airfoil 122′ whose sweep angle increases with increasing radius radially outward of the inner transition radius.
In the forward swept embodiment of the invention, as in the swept back embodiment, the nonincreasing sweep angle σ2 in the tip region 74 causes the endwall shock 64 to be coincident with the passage shock 66 for reducing the aerodynamic losses as discussed previously. This is in contrast to the prior art blade, shown in phantom where the endwall shock and the passage shock are distinct and therefore impose multiple aerodynamic losses on the working medium.
In the swept back embodiment of
The invention has been presented in the context of a fan blade for a gas turbine engine, however, the invention's applicability extends to any turbomachinery airfoil wherein flow passages between neighboring airfoils are subjected to multiple shocks.
Claims
1. A turbomachinery blade for a turbine engine having a cascade of blades rotatable about a rotational axis so that each blade in the cascade has a leading neighbor and a trailing neighbor, and each blade cooperates with its neighbors to define flow passages for a working medium gas, the blade cascade being circumscribed by a case and under some operational conditions an endwall shock extends a limited distance radially inward from the case and also extends axially and circumferentially across the flow passages, and a passage shock also extends across the flow passages, the turbomachinery blade including an airfoil having a leading edge, a trailing edge, a root, a tip and an inner transition point located at an inner transition radius radially inward of the tip, the blade characterized in that at least a portion of the leading edge radially outward of the inner transition point is swept and a section of the airfoil radially coextensive with the endwall shock extending from the leading neighbor intercepts the endwall shock so that the endwall shock and the passage shock are coincident.
2. A turbomachinery blade for a turbine engine having a cascade of blades rotatable about a rotational axis so that each blade in the cascade has a leading neighbor and a trailing neighbor, and each blade cooperates with its neighbors to define flow passages for a working medium gas, the blade cascade being circumscribed by a case and under some operational conditions an endwall shock extends a limited distance radially inward from the case and also extends axially and circumferentially across the flow passages and a passage shock also extends across the flow passages, the turbomachinery blade including an airfoil having a leading edge, a trailing edge, a root, a tip located at a tip radius, an inner transition point located at an inner transition radius radially inward of the tip, and an outer transition point at an outer transition radius radially intermediate the inner transition radius and the tip radius, the blade having a tip region bounded by the outer transition radius and the tip radius, and an intermediate region bounded by the inner transition radius and the outer transition radius, the blade characterized in that the leading edge is swept in the intermediate region at a first sweep angle which is generally nondecreasing with increasing radius, and the leading edge is swept over at least a portion of the tip region at a second sweep angle which is generally nonincreasing with increasing radius so that the section of the airfoil radially coextensive with the endwall shock extending from the leading neighbor intercepts the endwall shock so that the endwall shock and the passage shock are coincident.
3. The turbomachinery blade of claim 1 or 2 characterized in that the inner transition radius is coincident with the root at the leading edge of the blade.
4. A fan stage of a ducted fan gas turbine engine that is rotatable about an axis of rotation and defines a downstream direction along the axis of rotation, comprising:
- a fan casing that defines an inner duct wall having a fan rotor region;
- a hub disposed concentrically relative to the fan casing;
- a fan rotor that includes multiple swept fan blades, the swept fan blades being spaced apart around the hub, each of the multiple swept fan blades having:
- a tip profile that corresponds to the inner duct wall of the fan casing;
- a leading edge that defines a variable sweep angle in a direction perpendicular to the axis of rotation, the leading edge including: an inner region adjacent the hub, the inner region defining a forward sweep angle; an intermediate region between the inner region and the fan casing, the intermediate region defining a rearward sweep angle; and an outer region between the intermediate region and the fan casing, the outer region being translated forward relative to a leading edge with the same sweep angle as an outward boundary of the intermediate region.
5. The fan stage according to claim 4, wherein the leading edge at a boundary between the intermediate region and the inner region extends further upstream along the axis of rotation than the leading edge of the inner region.
6. The fan stage according to claim 4, wherein each of the multiple swept fan blades includes a hub contacting surface that extends further than the tip profile along the axis of rotation.
7. A fan stage of a ducted fan gas turbine engine that is rotatable about an axis of rotation and defines a downstream direction along the axis of rotation, comprising:
- a fan casing that defines an inner duct wall having a fan rotor region;
- a hub disposed concentrically relative to the fan casing;
- a fan rotor that includes multiple swept fan blades, the swept fan blades being spaced apart around the hub and being capable of rotating at speeds providing supersonic working medium gas velocities over the blades to cause a shock in the gas adjacent the inner duct wall, each of the multiple swept fan blades having:
- a tip profile that corresponds to the inner duct wall of the fan casing;
- a leading edge that defines a variable sweep angle in a direction perpendicular to the axis of rotation, the leading edge including: an inner region adjacent the hub, the inner region defining a forward sweep angle; an intermediate region between the inner region and the fan casing, the intermediate region defining a rearward sweep angle; and an outer region between the intermediate region and the fan casing, the outer region being translated forward relative to a leading edge with the same sweep angle as an outward boundary of the intermediate region to provide a sweep angle that causes the blade to intercept the shock.
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Type: Grant
Filed: Jun 5, 2001
Date of Patent: Oct 2, 2012
Assignee: United Technologies Corp. (Hartford, CT)
Inventors: David A. Spear (Manchester, CT), Dennis N. Kantor, legal representative (East Hartford, CT), Bruce P. Biederman (West Hartford, CT), John A. Orosa (Palm Beach Gardens, FL)
Primary Examiner: Christopher Verdier
Application Number: 09/874,931
International Classification: F01D 5/14 (20060101);