Source: https://patents.justia.com/patent/10570916
Timestamp: 2020-03-30 08:38:18
Document Index: 779193876

Matched Legal Cases: ['Application No. 61', 'Application No. 14882896', 'Application No. 14883503', 'Application No. 15752013', 'Application No. 15751617', 'Application No. 15793127', 'Application No. 15792720', 'Application No. 15792720', 'Application No. 15793193', 'Application No. 14883154', 'Application No. 14883117', 'Application No. 15752432', 'Application No. 15793425', 'Application No. 15793193', 'Application No. 15752887', 'Application No. 14883515', 'Application No. 15752124', 'Application No. 15751498', 'Application No. 15751454', 'Application No. 15793323', 'Application No. 15796827', 'Application No. 15792194', 'Application No. 15751738', 'Application No. 15752593', 'Application No. 14883036', 'Application No. 15793112', 'Application No. 15793268', 'Application No. 14883170', 'Application No. 15752593']

US Patent for Gas turbine engine airfoil Patent (Patent # 10,570,916 issued February 25, 2020) - Justia Patents Search
Justia Patents US Patent for Gas turbine engine airfoil Patent (Patent # 10,570,916)
Feb 16, 2015 - UNITED TECHNOLOGIES CORPORATION
This application claims priority to U.S. Provisional Application No. 61/941,685, which was filed on Feb. 19, 2014 and is incorporated herein by reference.
This disclosure relates to gas turbine engine airfoils. More particularly, the disclosure relates to airfoil axial stacking offset in, for example, a gas turbine engine compressor.
A turbine engine such as a gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes at least low and high pressure compressors, and the turbine section includes at least low and high pressure turbines.
Direct drive gas turbine engines include a fan section that is driven directly by one of the turbine shafts. Rotor blades in the fan section and a low pressure compressor of the compressor section of direct drive engines rotate in the same direction.
Gas turbine engines have been proposed in which a geared architecture is arranged between the fan section and at least some turbines in the turbine section. The geared architecture enables the associated compressor of the compressor section to be driven at much higher rotational speeds, improving overall efficiency of the engine. The propulsive efficiency of a gas turbine engine depends on many different factors, such as the design of the engine and the resulting performance debits on the fan that propels the engine and the compressor section downstream from the fan. Physical interaction between the fan and the air causes downstream turbulence and further losses. Although some basic principles behind such losses are understood, identifying and changing appropriate design factors to reduce such losses for a given engine architecture has proven to be a complex and elusive task.
Prior compressor airfoil geometries may not be suitable for the compressor section of gas turbine engines using a geared architecture, since the significantly different speeds of the compressor changes the desired aerodynamics of the airfoils within the compressor section. Counter-rotating fan and compressor blades, which may be used in geared architecture engines, also present design challenges.
In one exemplary embodiment, a compressor airfoil of a turbine engine having a geared architecture includes pressure and suction sides that extend in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between an axial stacking offset and span position that includes a curve with a negative slope from 90% span to 100% span. The negative slope leans forward relative to an engine axis.
In a further embodiment of the above, the curve has a negative slope from 80% span to 100% span.
In a further embodiment of any of the above, the curve has a negative slope from 70% span to 100% span.
In a further embodiment of any of the above, the curve has a positive slope beginning at 0% span. The positive slope leans aftward relative to the engine axis.
In a further embodiment of any of the above, the positive slope extends from 0% span to 40% span.
In a further embodiment of any of the above, the curve transitions from the positive slope to the negative slope in a range of 40% span to 75% span.
In another exemplary embodiment, a gas turbine engine includes a combustor section that is arranged between a compressor section and a turbine section. A fan section has an array of twenty-six or fewer fan blades. The fan section has a fan pressure ratio of less than 1.55. A geared architecture couples the fan section to the turbine section or the compressor section. An airfoil is arranged in the compressor section and includes pressure and suction sides that extend in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between an axial stacking offset and span position that includes a curve with a negative slope from 90% span to 100% span. The negative slope leans forward relative to an engine axis.
In a further embodiment of any of the above, the compressor section includes at least a low pressure compressor and a high pressure compressor. The high pressure compressor is arranged immediately upstream of the combustor section.
In a further embodiment of any of the above, the airfoil is provided in a compressor outside the high pressure compressor.
In a further embodiment of any of the above, the low pressure compressor is counter-rotating relative to the fan blades.
In a further embodiment of any of the above, the gas turbine engine is a two-spool configuration.
In a further embodiment of any of the above, the low pressure compressor is immediately downstream from the fan section.
In a further embodiment of any of the above, the airfoil is rotatable relative to an engine static structure.
In a further embodiment of any of the above, the curve has a negative slope from 80% span to 100% span.
FIG. 1 schematically illustrates a gas turbine engine embodiment with a geared architecture.
FIG. 2 schematically illustrates a low pressure compressor section of the gas turbine engine of FIG. 1.
FIG. 3 is a schematic view of airfoil span positions.
FIG. 4 is a schematic view of a cross-section of an airfoil sectioned at a particular span position and depicting directional indicators.
FIG. 5 graphically depicts curves for several example airfoil axial stacking offset relative to span, including two prior art curves and several inventive curves according to this disclosure.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.55. In another non-limiting embodiment the low fan pressure ratio is less than about 1.45. In another non-limiting embodiment the low fan pressure ratio is from 1.1 to 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7°R)]°0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1200 ft/second (365.7 meters/second).
Referring to FIG. 2, which schematically illustrates an example low pressure compressor (LPC) 44, a variable inlet guide vane (IGV) is arranged downstream from a fan exit stator (FES). The figure is highly schematic, and the geometry and orientation of various features may be other than shown. An actuator driven by a controller actuates the IGV about their respective axes. Multiple airfoils are arranged downstream from the IGV. The airfoils include alternating stages of rotors (ROTOR1, ROTOR2, ROTOR3, ROTOR4) and stators (STATOR1, STATOR2, STATOR3, STATOR4). In the example shown in FIG. 2, the LPC includes four rotors alternating with four stators. However, in another example, a different number of rotors and a different number of stators may be used. Moreover, the IGV and stator stages may all be variable, fixed or a combination thereof.
The disclosed airfoils may be used in a low pressure compressor of a two spool engine or in portions of other compressor configurations, such as low, intermediate and/or high pressure areas of a three spool engine. However, it should be understood that any of the disclosed airfoils may be used for blades or vanes, and in any of the compressor section, turbine section and fan section.
Referring to FIG. 3, span positions on an airfoil 64 are schematically illustrated from 0% to 100% in 10% increments (i.e., 0, 10, 20, 30, 40, 50, 60, 70, 80, 90, 100). Each section at a given span position is provided by a conical cut that corresponds to the shape of the core flow path, as shown by the large dashed lines. In the case of an airfoil with an integral platform, the 0% span position corresponds to the radially innermost location where the airfoil meets the fillet joining the airfoil to the inner platform. In the case of an airfoil without an integral platform, the 0% span position corresponds to the radially innermost location where the discrete platform meets the exterior surface of the airfoil. For airfoils having no outer platform, such as blades, the 100% span position corresponds to the tip 66. For airfoils having no platform at the inner diameter, such as cantilevered stators, the 0% span position corresponds to the inner diameter location of the airfoil. For stators, the 100% span position corresponds to the outermost location where the airfoil meets the fillet joining the airfoil to the outer platform.
Airfoils in each stage of the LPC are specifically designed radially from an inner airfoil location (0% span) to an outer airfoil location (100% span) and along circumferentially opposite pressure and suction sides 72, 74 extending in chord between a leading and trailing edges 68, 70 (see FIG. 4). Each airfoil is specifically twisted with a corresponding stagger angle and bent with specific sweep and/or dihedral angles along the airfoil. Airfoil geometric shapes, stacking offsets, chord profiles, stagger angles, sweep and dihedral angles, among other associated features, are incorporated individually or collectively to improve characteristics such as aerodynamic efficiency, structural integrity, and vibration mitigation, for example, in a gas turbine engine with a geared architecture in view of the higher LPC rotational speeds.
The airfoil 64 has an exterior surface 76 providing a contour that extends from a leading edge 68 generally aftward in a chord-wise direction H to a trailing edge 70, as shown in FIG. 4. Pressure and suction sides 72, 74 join one another at the leading and trailing edges 68, 70 and are spaced apart from one another in an airfoil thickness direction T. An array of airfoils 64 are positioned about the axis X (corresponding to an X direction) in a circumferential or tangential direction Y. Any suitable number of airfoils may be used for a particular stage in a given engine application.
An axial stacking offset Xd corresponds to the location of the center of gravity XCG for a particular section at a given span location relative to a reference point 80 in the X direction. The reference point 80 is a location such as the axial center of the root or the axial center of a rotor bore, for example. The value Xd corresponds to the axial distance from the reference point 80 to the center of gravity. A positive X is on the aft side of reference point 80, and a negative X is on the forward side of the reference point 80. A positive slope to the axial stacking offset profile is where the span section leans aftward, and a negative slope to the axial stacking offset profile is where the span section leans forward relative to the engine's axis X.
The exterior surface 76 of the airfoil 64 generates lift based upon its geometry and directs flow along the core flow path C. The airfoil 64 may be constructed from a composite material, or an aluminum alloy or titanium alloy, or a combination of one or more of these. Abrasion-resistant coatings or other protective coatings may be applied to the airfoil. The rotor stages may constructed as an integrally bladed rotor, if desired, or discrete blades having roots secured within corresponding rotor slots of a disc. The stators may be provided by individual vanes, clusters of vanes, or a full ring of vanes.
Airfoil geometries can be described with respect to various parameters provided. The disclosed graph(s) illustrate the relationships between the referenced parameters within 10% of the desired values, which correspond to a hot aerodynamic design point for the airfoil. In another example, the referenced parameters are within 5% of the desired values, and in another example, the reference parameters are within 2% of the desired values. It should be understood that the airfoils may be oriented differently than depicted, depending on the rotational direction of the blades. The signs (positive or negative) used, if any, in the graphs of this disclosure are controlling and the drawings should then be understood as a schematic representation of one example airfoil if inconsistent with the graphs. The signs in this disclosure, including any graphs, comply with the “right hand rule.” The axial stacking offset varies with position along the span, and varies between a hot, running condition and a cold, static (“on the bench”) condition.
The geared architecture 48 of the disclosed example permits the fan 42 to be driven by the low pressure turbine 46 through the low speed spool 30 at a lower angular speed than the low pressure turbine 46, which enables the LPC 44 to rotate at higher, more useful speeds. The axial stacking offset in a hot, running condition along the span of the airfoils 64 provides necessary compressor operation in cruise at the higher speeds enabled by the geared architecture 48, to thereby enhance aerodynamic functionality and thermal efficiency. As used herein, the hot, running condition is the condition during cruise of the gas turbine engine 20. For example, the axial stacking offset in the hot, running condition can be determined in a known manner using numerical analysis, such as finite element analysis.
FIG. 5 illustrates the relationship between the axial stacking offset and the average span (AVERAGE SPAN %), which is the average of the radial position at the leading and trailing edges 68, 70. The axial stacking offset assumes a center of gravity based upon a homogeneous material throughout the airfoil cross-section. In one example, the airfoils are LPC rotor blades. Two prior art curves (“PRIOR ART”) are illustrated as well as several example inventive curves 88, 90, 92, 94, 96. The airfoil 64 has a relationship between an axial stacking offset and span position that includes a curve with a negative slope from 90% span to 100% span. The prior art airfoil curves are substantially linear across the entire span of the airfoil with substantially no offset. The curves 88, 90, 92, 94, 96 have a negative slope from 80% span to 100% span in the example. The curves may have a negative slope from at least 70% span to 100% span, for example.
The inventive curves 90, 92, 94, 96 have a positive slope beginning at 0% span and may extend from 0% span to at least 40% span, for example. The curves transition from the positive slope to the negative slope in a range of 40% span to 75% span, providing a maximum aftward leaning axial stacking offset in this range.
The prior art has generally used straight, axially non-stacked LPC blades. The disclosed airfoils include significant forward and aftward stacking to improve the aerodynamic efficiency of the high speed LPC blades downstream from a counter-rotating fan.
a combustor section arranged between a compressor section and a turbine section, wherein the compressor section includes at least a low pressure compressor and a high pressure compressor, the high pressure compressor arranged upstream of the combustor section;
a fan section having an array of twenty-six or fewer fan blades, wherein the fan section has a fan pressure ratio of less than 1.55, wherein the low pressure compressor is counter-rotating relative to the fan blades, wherein the low pressure compressor is downstream from the fan section;
a geared architecture coupling the fan section to the turbine section or the compressor section; and
an airfoil arranged in the low pressure compressor and including pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position, wherein the airfoil has a relationship between an axial stacking offset and span position that includes a curve with a negative slope from 90% span to 100% span, where the negative slope leans forward relative to an engine axis.
2. The gas turbine engine according to claim 1, wherein the gas turbine engine is a two-spool configuration.
3. The gas turbine engine according to claim 1, wherein the airfoil is rotatable relative to an engine static structure.
4. The gas turbine engine according to claim 1, wherein the negative slope is from 80% span to 100% span.
5. The gas turbine engine according to claim 4, wherein the negative slope is from 70% span to 100% span.
6. The gas turbine engine according to claim 1, wherein the curve has a positive slope beginning at 0% span, where the positive slope leans aftward relative to the engine axis.
7. The gas turbine engine according to claim 6, wherein the positive slope extends from 0% span to 40% span.
8. The gas turbine engine according to claim 7, wherein the curve transitions from the positive slope to the negative slope in a range of 40% span to 75% span.
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Patent number: 10570916
Patent Publication Number: 20170159670
Inventors: Edward J. Gallagher (West Hartford, CT), Lisa I. Brilliant (Middletown, CT), Joseph C. Straccia (Middletown, CT), Stanley J. Balamucki (The Villages, FL), Mark A. Stephens (Wethersfield, CT), Kate Hudon (Superior, CO)
Application Number: 15/115,360
International Classification: F04D 29/32 (20060101); F04D 29/54 (20060101); F01D 5/14 (20060101); F01D 9/02 (20060101); F02C 3/04 (20060101); F02C 7/36 (20060101); F02K 3/06 (20060101); F04D 29/38 (20060101);