Power plants – Reaction motor – Interrelated reaction motors
Reexamination Certificate
1999-02-12
2001-03-06
Thorpe, Timothy S. (Department: 3746)
Power plants
Reaction motor
Interrelated reaction motors
C416S22300B, C415S208100
Reexamination Certificate
active
06195983
ABSTRACT:
BACKGROUND OF THE INVENTION
The Government has rights to this invention pursuant to a contract by the United States Air Force.
FIELD OF THE INVENTION
The present invention relates generally to an aircraft gas turbine engine having fan bypass ducts, and more particularly, to exit guide vanes and frame struts and splitters for dividing the fan flow between core and bypass ducts.
DISCUSSION OF THE BACKGROUND ART
A gas turbine engine includes a core engine having, in serial axial flow relationship, a high pressure compressor to compress the airflow entering the core engine, a combustor in which a mixture of fuel and the compressed air is burned to generate a propulsive gas flow, and a high pressure turbine which is rotated by the propulsive gas flow and which is connected by a large diameter shaft to drive the high pressure compressor. A typical bypass turbofan engine adds a low pressure turbine aft of the high pressure turbine which drives a fan forward of the high pressure compressor. A splitter aft of the fan divides fan flow exiting the fan into core engine flow and bypass flow around the core engine. The splitter is mounted to struts of a forward engine frame, referred to as the fan frame, and separates a bypass duct inlet from a core engine duct inlet.
The fan frame struts and the splitter are placed at a sufficient distance from fan outlet guide vanes directly aft of fan blades of the fan such that there is enough length in an empty gap extending from an exit guide vane trailing edge to a strut leading edge such that the static pressure field from the struts do not locally back-pressure the exit guide vanes and cause non-uniform flows which create high pressure losses, reduced stall margin, or high airflow distortion transfer. It is highly desirable to minimize the length of gap without producing this undesired aerodynamic back-pressure which causes the loss of engine performance, aerodynamic instability of the fan and compressor, or adversely impacting the fan airflow distortion transfer to the compressor.
Prior art studies have suggested the use of non-axisymmetric stator or fan exit guide vane configurations ahead of pylons and/or struts to achieve a uniform flow field upstream of the stators. Two and three-dimensional potential flow analyses have been used to conceptually develop stators with varying degree of camber angles in the vicinity of the strut to protect the rotor from the pressure disturbance induced forward by the service strut. Experiments of O'Brien, Reimers and Richardson (1983) and Woodard and Balombin (1984) confirmed the importance of the potential flow interaction between rotor, stator and struts in the production of rotor blade pressure fluctuations. In 1996, Parry using a simpler calculation scheme arrived at results similar to those obtained by Rubbert showing the relative benefits of increasing the number of vane types to alleviate the exit guide vane and frame flow interaction. Shrinivas and Giles (1995) performed detailed sensitivity studies to arrive at a cyclically varying exit guide vane trailing edge camber configuration, while maintaining the leading edge airfoil shape, thus retaining uniform exit guide vane incidence angle. These, and other prior art studies, are disclosed in the prior art publications cited herein. These types of stator camber modifications described above, while technically very effective, are usually unattractive from a manufacturing cost and maintenance perspective.
As a practical alternative to building stator cascades with different camber angles, a slightly more attractive option that lowers the cost of implementation uses circumferentially re-staggered stator vanes in front of the struts to channel the exit flow from the stator trailing edge smoothly around the strut leading edge. Yokoi, Nagano and Kakehi (1981) tested several such stator re-stagger options showing large reductions in upstream pressure disturbance. Cerri and O'Brien (1989) used the classical Douglas-Neumann singularity superposition method to solve the cascade-strut system to arrive at an optimal staggering of the stator configuration that predicted the lowest upstream pressure disturbance. Jones, Barton and O'Brien (1996) used the Douglas-Neumann formulation of Cerri and O'Brien to design and test exit guide vanes with circumferentially non-uniform stator stagger angle distribution and showed reduced static pressure disturbance on the rotor. Studies have also been directed at reducing the flow interactions to improve fan acoustic levels are described by McArdle, Jones, Heidelberg and Homyak (1980), Ho (1981), Nakamura, Isomura and Kodama (1986) and Preisser, Schoenster, Golub and Home (1981). These engine configurations are still relatively unattractive from a manufacturing cost and maintenance perspective. It is highly desirable that all the exit guide vanes have the same shape, size and stagger angle.
SUMMARY OF THE INVENTION
A gas turbine engine fan assembly has a plurality of axially swept fan exit guide vanes circumferentially disposed around an axially extending centerline, a plurality of fan frame struts having axially swept strut leading edges and circumferentially disposed around the centerline directly aft of the exit guide vanes, and an axially extending annular gap between the trailing edges of the fan exit guide vanes and the strut leading edges. Each of the fan exit guide vanes has pressure and suction sides, vane trailing edges, and is circumferentially leaned such that the pressure side facing radially inward. Preferably, the axially swept vane trailing edges and the axially swept strut leading edges generally conform to each other in shape. An annular splitter is radially disposed at a spanwise position to split airflow from the fan into fan bypass airflow and core engine airflow and includes a splitter leading edge preferably positioned aft of the strut leading edges.
Various embodiments include the swept strut leading and swept vane trailing edge having sweeps chosen from a plurality of sweeps in a radially outer direction, the plurality consisting of forward, aft, aft inner and forward outer, and forward inner and aft outer sweeps.
In one particular embodiment of the invention, the vanes are circumferentially curved having stacking lines that curve in a first circumferential in a first radial portion of the vane and curve in a second circumferential direction in a second radial portion of the vane. In one particular embodiment of the invention, a circumferential lean of the vane is concentrated around a pitch line of the vanes between the first and second radial portions of the vane. A yet more particular embodiment of the invention provides that the stacking lines are substantially linear around the pitch lines between the first and second radial portions of the vanes.
ADVANTAGES OF THE INVENTION
The present invention allows leading edges of fan struts to be positioned closer to fan exit guide vanes with a smaller empty gap therebetween. It allows such a configuration without having tailored flow near the struts by using different tailored airfoils or stagger angles for the fan exit guide vanes. The invention, therefore, offers aircraft gas turbine engine designers the advantage of designing shorter and less complex engines that have reduced cost and maintenance fees than would otherwise be possible. The present invention provides an aerodynamically closely coupled swept and leaned exit guide vane and frame strut and splitter system. The invention allows aircraft gas turbine engine designers to minimize the length of gap without producing an undesired aerodynamic back-pressure, loss of engine performance, aerodynamic instability of the fan and compressor, or adversely impacting the fan airflow distortion transfer to the compressor.
One benefit of the circumferential lean in the exit guide vanes is lowering the exit guide vane hub exit Mach number, thus, reducing the required duct velocity diffusion ratio. The lean is such to custom tailor the flow radially inwards in to the hub endwall. Combining the lean w
Hemmelgarn Bruce A.
Szucs Peter N.
Wadia Aspi R.
General Electric Company
Hess Andrew C.
Thorpe Timothy S.
Torrente David J.
Young Rodney M.
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