Gas turbine engine airfoil

Information

  • Patent Grant
  • 9777580
  • Patent Number
    9,777,580
  • Date Filed
    Friday, January 27, 2017
    7 years ago
  • Date Issued
    Tuesday, October 3, 2017
    7 years ago
Abstract
An airfoil for a turbine engine includes an airfoil that has pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a leading edge dihedral and a span position. The leading edge dihedral is negative from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning. The airfoil has a relationship between a trailing edge dihedral and a span position. The trailing edge dihedral is positive from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning and a negative dihedral corresponds to pressure side-leaning. The airfoil includes at least one of a least negative dihedral in a range of 5-15% span position, a least negative dihedral in a range of 0-10% span position and a least positive trailing edge dihedral in a 40%-55% span position.
Description
BACKGROUND

This disclosure relates generally to an airfoil for gas turbine engines, and more particularly to leading and trailing edge aerodynamic dihedral relative to span for gas turbine engine blades.


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 low and high pressure compressors, and the turbine section includes low and high pressure turbines.


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. As an example, the fan may rotate at a high rate of speed such that air passes over the fan airfoils at transonic or supersonic speeds. The fast-moving air creates flow discontinuities or shocks that result in irreversible propulsive losses. Additionally, 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.


SUMMARY

In one exemplary embodiment, an airfoil for a turbine engine includes an airfoil that has pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a leading edge dihedral and a span position. The leading edge dihedral is negative from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning. The airfoil has a relationship between a trailing edge dihedral and a span position. The trailing edge dihedral is positive from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning and a negative dihedral corresponds to pressure side-leaning. The airfoil includes at least one of a least negative dihedral in a range of 5-15% span position, a least negative dihedral in a range of 0-10% span position and a least positive trailing edge dihedral in a 40%-55% span position.


In a further embodiment of the above airfoil, the leading edge dihedral at the 0% span position is in the range of −3° to −12°.


In a further embodiment of any of the above airfoils, the leading edge dihedral at the 0% span position is about −4°.


In a further embodiment of any of the above airfoils, the leading edge dihedral at the 0% span position is about −10°.


In a further embodiment of any of the above airfoils, the leading edge dihedral extends from the 0% span position to a 20% span position and has a leading edge dihedral in a range of −2° to −6°.


In a further embodiment of any of the above airfoils, the leading edge dihedral includes a first point at a 75% span position and extends generally linearly from the first point to a second point at the 85% span position.


In a further embodiment of any of the above airfoils, the first point is in a range of −8° to −10° dihedral.


In a further embodiment of any of the above airfoils, the second point is in a range of −3° to −6° dihedral.


In a further embodiment of any of the above airfoils, a maximum negative dihedral is in a range of 95-100% span position.


In a further embodiment of any of the above airfoils, a maximum negative dihedral is in a range of 65-75% span position.


In a further embodiment of any of the above airfoils, a maximum negative dihedral is in a range of 50-60% span position.


In a further embodiment of any of the above airfoils, the relationship provides a generally C-shaped curve from the 0% span position to a 50% span position and then a 90% span position.


In a further embodiment of any of the above airfoils, a trailing edge dihedral at the 0% span position is in a range of 20° to 25°.


In a further embodiment of any of the above airfoils, a trailing edge dihedral at about the 50% span position is in a range of 2° to 6°.


In a further embodiment of any of the above airfoils, a trailing edge dihedral at the 90% span position is in a range of 16° to 22°.


In a further embodiment of any of the above airfoils, from a 65% span position to a 75% span position the trailing edge dihedral increases about 5°.


In a further embodiment of any of the above airfoils, a positive-most trailing edge dihedral in a 80%-100% span position is within 5° of the trailing edge dihedral in the 0% span position.


In a further embodiment of any of the above airfoils, the airfoil is a fan blade for a gas turbine engine.





BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:



FIG. 1 schematically illustrates a gas turbine engine embodiment.



FIG. 2A is a perspective view of a portion of a fan section.



FIG. 2B is a schematic cross-sectional view of the fan section.



FIG. 2C is a cross-sectional view a fan blade taken along line 2C-2C in FIG. 2B.



FIG. 3A is a schematic view of fan blade span positions.



FIG. 3B is a schematic representation of a dihedral angle for an airfoil.



FIG. 4A illustrates a relationship between a leading edge aerodynamic dihedral angle and a span position for a set of first example airfoils and a prior art curve.



FIG. 4B illustrates a relationship between a trailing edge aerodynamic dihedral angle and a span position for a set of first example airfoils and a prior art curve.



FIG. 5A illustrates a relationship between a leading edge aerodynamic dihedral angle and a span position for a set of second example airfoils and a prior art curve.



FIG. 5B illustrates a relationship between a trailing edge aerodynamic dihedral angle and a span position for set of second example airfoils and a prior art curve.





The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible.


DETAILED DESCRIPTION


FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. That is, the disclosed airfoils may be used for engine configurations such as, for example, direct fan drives, or two- or three-spool engines with a speed change mechanism coupling the fan with a compressor or a turbine sections.


The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis X relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.


The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis X which is collinear with their longitudinal axes.


The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.


The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1). Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.


The example gas turbine engine includes the fan 42 that comprises in one non-limiting embodiment less than about twenty-six (26) fan blades. In another non-limiting embodiment, the fan section 22 includes less than about twenty (20) fan blades. Moreover, in one disclosed embodiment the low pressure turbine 46 includes no more than about six (6) turbine rotors schematically indicated at 34. In another non-limiting example embodiment the low pressure turbine 46 includes about three (3) turbine rotors. A ratio between the number of fan blades 42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine 46 provides the driving power to rotate the fan section 22 and therefore the relationship between the number of turbine rotors 34 in the low pressure turbine 46 and the number of blades 42 in the fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.


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. The flight condition of 0.8 Mach and 35,000 ft, 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 another non-limiting embodiment is less than about 1200 ft/second.


Referring to FIG. 2A-2C, the fan blade 42 is supported by a fan hub 60 that is rotatable about the axis X, which corresponds to the X direction. Each fan blade 42 includes an airfoil 64 extending in a radial span direction R from a root 62 to a tip 66. A 0% span position corresponds to a section of the airfoil 64 at the inner flow path (e.g., a platform), and a 100% span position corresponds to a section of the airfoil 64 at the tip 66.


The root 62 is received in a correspondingly shaped slot in the fan hub 60. The airfoil 64 extends radially outward of the platform, which provides the inner flow path. The platform may be integral with the fan blade or separately secured to the fan hub, for example. A spinner 66 is supported relative to the fan hub 60 to provide an aerodynamic inner flow path into the fan section 22.


The airfoil 64 has an exterior surface 76 providing a contour that extends from a leading edge 68 aftward in a chord-wise direction H to a trailing edge 70, as shown in FIG. 2C. 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 the fan blades 42 are positioned about the axis X in a circumferential or tangential direction Y. Any suitable number of fan blades may be used in a given application.


The exterior surface 76 of the airfoil 64 generates lift based upon its geometry and directs flow along the core flow path C. The fan blade 42 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 fan blade 42. The curves and associated values assume a fan in a hot, running condition (typically cruise).


One characteristic of fan blade performance relates to the fan blade's leading and trailing edge aerodynamic dihedral angle relative to a particular span position (R direction). Referring to FIG. 3A, span positions a schematically illustrated from 0% to 100% in 10% increments. 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 a fan blade 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 platform. In the case of a fan blade 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. In addition to varying with span, leading and trailing edge sweep varies between a hot, running condition and a cold, static (“on the bench”) condition.


An aerodynamic dihedral angle D (simply referred to as “dihedral”) is schematically illustrated in FIG. 3B for a simple airfoil. An axisymmetric stream surface S passes through the airfoil 64 at a location that corresponds to a span location (FIG. 3A). For the sake of simplicity, the dihedral D relates to the angle at which a line L along the leading or trailing edge tilts with respect to the stream surface S. A plane P is normal to the line L and forms an angle with the tangential direction Y, providing the dihedral D. A positive dihedral D corresponds to the line tilting toward the suction side (suction side-leaning), and a negative dihedral D corresponds to the line tilting toward the pressure side (pressure side-leaning). The method of determining and calculating the dihedral for more complex airfoil geometries is disclosed in Smith Jr., Leroy H., and Yeh, Hsuan “Sweep and Dihedral Effects in Axial-Flow Turbomachinery.” J. Basic Eng. Vol. 85 Iss. 3, pp. 401-414 (Sep. 1, 1963), which is incorporated by reference in its entirety.


Example relationships between the leading edge dihedral (LE DIHEDRAL) and the span position (LE SPAN %) are shown in FIGS. 4A and 5A for several example fan blades, each represented by a curve. Only one curve in each graph is discussed for simplicity. In the examples, the leading edge dihedral is negative from the 0% span position to the 100% span position.


The leading edge dihedral at the 0% span position (92 in FIG. 4A; 104 in FIG. 5A) is in the range of −3° to −12°. In the examples shown in FIGS. 4A and 5A, the leading edge dihedral at the 0% span position is about −4°.


The leading edge dihedral extends from the 0% span position to a 20% span position (96 in FIG. 4A; 108 in FIG. 5A) having a leading edge dihedral in a range of −2° to −6°.


In the examples shown in FIGS. 4A and 5A, the leading edge dihedral includes a first point (100 in FIG. 4A; 110 in FIG. 5A) at a 75% span position and extends generally linearly from the first point to a second point (102 in FIG. 4A; 114 in FIG. 5A) at the 85% span position. The first point is in a range of −8° to −10° dihedral, and the second point is in a range of −3° to −6° dihedral.


Referring to FIG. 4A, a maximum negative dihedral 98 is in a range of 95-100% span position. A least negative dihedral 94 is in a range of 5-15% span position. Referring to FIG. 5A, a maximum negative dihedral 110 is in a range of 65-75% span position. A least negative dihedral 106 is in a range of 0-10% span position.


Example relationships between the trailing edge dihedral (TE DIHEDRAL) and the span position (TE SPAN %) are shown in FIGS. 4B and 5B for several example fan blades, each represented by a curve. Only one curve in each graph is discussed for simplicity. In the examples, the trailing edge dihedral is positive from the 0% span position to the 100% span position. The relationship provides a generally C-shaped curve from the 0% span position to a 50% span position and then a 90% span position.


A trailing edge dihedral (130 in FIG. 4B; 140 in FIG. 5B) at the 0% span position is in a range of 20° to 25°. A trailing edge dihedral (132 in FIG. 4B; 142 in FIG. 5B) at about the 50% span position is in a range of 2° to 6°. A trailing edge dihedral (137 in FIG. 4B; 147 in FIG. 5B) at the 90% span position is in a range of 16° to 22°. From a 65% span position (134 in FIG. 4B; 144 in FIG. 5B) to a 75% span position (136 in FIG. 4B; 146 in FIG. 5B) the trailing edge dihedral increases about 5°.


A positive-most trailing edge dihedral (138 in FIG. 4B; 148 in FIG. 5B) in a 80%-100% span position is within 5° of the trailing edge dihedral in the 0% span position (130 in FIG. 4B; 140 in FIG. 5B). A least positive trailing edge dihedral (132 in FIG. 4B; 142 in FIG. 5B) is in a 40%-55% span position.


The leading and trailing edge aerodynamic dihedral in a hot, running condition along the span of the airfoils 64 relate to the contour of the airfoil and provide necessary fan operation in cruise at the lower, preferential speeds enabled by the geared architecture 48 in order to 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 leading and trailing edge aerodynamic dihedral in the hot, running condition can be determined in a known manner using numerical analysis, such as finite element analysis.


It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention.


Although the different examples have specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.


Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.

Claims
  • 1. An airfoil for a turbine engine comprising: an airfoil having pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip, wherein the airfoil has a relationship between a leading edge dihedral and a span position, the leading edge dihedral negative from the 0% span position to the 100% span position, wherein a positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning;wherein the airfoil has a relationship between a trailing edge dihedral and a span position, the trailing edge dihedral positive from the 0% span position to the 100% span position, wherein a positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning; andwherein the airfoil includes one of:a least negative dihedral is in a range of 5-15% span position;a least negative dihedral is in a range of 0-10% span position; anda least positive trailing edge dihedral is in a 40%-55% span position.
  • 2. The airfoil according to claim 1, wherein the leading edge dihedral at the 0% span position is in the range of −3° to −12°.
  • 3. The airfoil according to claim 2, wherein the leading edge dihedral at the 0% span position is about −4°.
  • 4. The airfoil according to claim 2, wherein the leading edge dihedral at the 0% span position is about −10°.
  • 5. The airfoil according to claim 2, wherein the leading edge dihedral extends from the 0% span position to a 20% span position having a leading edge dihedral in a range of −2° to −6°.
  • 6. The airfoil according to claim 5, wherein the leading edge dihedral includes a first point at a 75% span position and extends generally linearly from the first point to a second point at the 85% span position.
  • 7. The airfoil according to claim 6, wherein the first point is in a range of −8° to −10° dihedral.
  • 8. The airfoil according to claim 6, wherein the second point is in a range of −3° to −6° dihedral.
  • 9. The airfoil according to claim 2, wherein a maximum negative dihedral is in a range of 95-100% span position.
  • 10. The airfoil according to claim 2, wherein a maximum negative dihedral is in a range of 65-75% span position.
  • 11. The airfoil according to claim 2, wherein a maximum negative dihedral is in a range of 50-60% span position.
  • 12. The airfoil according to claim 11, wherein the relationship provides a generally C-shaped curve from the 0% span position to a 50% span position and then a 90% span position.
  • 13. The airfoil according to claim 12, wherein a trailing edge dihedral at the 0% span position is in a range of 20° to 25°.
  • 14. The airfoil according to claim 13, wherein a trailing edge dihedral at about the 50% span position is in a range of 2° to 6°.
  • 15. The airfoil according to claim 13, wherein a trailing edge dihedral at the 90% span position is in a range of 16° to 22°.
  • 16. The airfoil according to claim 13, wherein from a 65% span position to a 75% span position the trailing edge dihedral increases about 5°.
  • 17. The airfoil according to claim 14, wherein a positive-most trailing edge dihedral in a 80%-100% span position is within 5° of the trailing edge dihedral in the 0% span position.
  • 18. The airfoil according to claim 12, wherein the airfoil is a fan blade for a gas turbine engine.
CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. application Ser. No. 15/191,860 filed on Jun. 24, 2016 which in a continuation of U.S. application Ser. No. 14/876,995 filed on Oct. 7, 2015 which is now U.S. Pat. No. 9,399,917 granted Jul. 26, 2016 which is a continuation of U.S. application Ser. No. 14/624,025 filed on Feb. 17, 2015, which is now U.S. Pat. No. 9,353,628 issued May 31, 2016, which claims priority to U.S. Provisional Application No. 61/942,025 filed on Feb. 19, 2014.

US Referenced Citations (156)
Number Name Date Kind
2714499 Warner Aug 1955 A
2746672 Doll, Jr. May 1956 A
2934259 Hausmann Apr 1960 A
3194487 Tyler et al. Jul 1965 A
3287906 McCormick Nov 1966 A
3747343 Rosen Jul 1973 A
3754484 Roberts Aug 1973 A
3867062 Troller Feb 1975 A
3892358 Gisslen Jul 1975 A
3905191 Matto Sep 1975 A
3932058 Harner et al. Jan 1976 A
3935558 Miller et al. Jan 1976 A
4012172 Schwaar et al. Mar 1977 A
4130872 Harloff Dec 1978 A
4431376 Lubenstein et al. Feb 1984 A
4682935 Martin Jul 1987 A
4826400 Gregory May 1989 A
4900230 Patel Feb 1990 A
5088892 Weingold et al. Feb 1992 A
5141400 Murphy et al. Aug 1992 A
5167489 Wadia et al. Dec 1992 A
5192190 Ferleger Mar 1993 A
5211703 Ferleger May 1993 A
5221181 Ferleger Jun 1993 A
5277549 Chen et al. Jan 1994 A
5433674 Sheridan et al. Jul 1995 A
5443367 Samit et al. Aug 1995 A
5447411 Curley et al. Sep 1995 A
5524341 Ferleger Jun 1996 A
5524847 Brodell et al. Jun 1996 A
5525038 Sharma et al. Jun 1996 A
5624234 Neely et al. Apr 1997 A
5642985 Spear et al. Jul 1997 A
5725354 Wadia et al. Mar 1998 A
5778659 Duesler et al. Jul 1998 A
5857836 Stickler et al. Jan 1999 A
5915917 Eveker et al. Jun 1999 A
5975841 Lindemuth et al. Nov 1999 A
6059532 Chen et al. May 2000 A
6071077 Rowlands Jun 2000 A
6079948 Sasaki et al. Jun 2000 A
6195983 Wadia Mar 2001 B1
6223616 Sheridan May 2001 B1
6299412 Wood et al. Oct 2001 B1
6312219 Wood et al. Nov 2001 B1
6318070 Rey et al. Nov 2001 B1
6328533 Decker et al. Dec 2001 B1
6331100 Liu et al. Dec 2001 B1
6341942 Chou et al. Jan 2002 B1
6565334 Bradbury May 2003 B1
6607165 Manteiga et al. Aug 2003 B1
6814541 Evans et al. Nov 2004 B2
6899526 Doloresco et al. May 2005 B2
6994524 Owen et al. Feb 2006 B2
7021042 Law Apr 2006 B2
7204676 Dutton et al. Apr 2007 B2
7328580 Lee et al. Feb 2008 B2
7374403 Decker et al. May 2008 B2
7396205 Dube et al. Jul 2008 B2
7476086 Wadia et al. Jan 2009 B2
7497664 Walter et al. Mar 2009 B2
7547186 Schuster et al. Jun 2009 B2
7591754 Duong et al. Sep 2009 B2
7785075 Botrel et al. Aug 2010 B2
7806651 Kennepohl et al. Oct 2010 B2
7806653 Burton et al. Oct 2010 B2
7824305 Duong et al. Nov 2010 B2
7926260 Sheridan et al. Apr 2011 B2
7967571 Wood et al. Jun 2011 B2
7997868 Liang et al. Aug 2011 B1
7997872 Wilson Aug 2011 B2
7997882 Shulver Aug 2011 B2
8087885 Suciu et al. Jan 2012 B2
8147207 Orosa et al. Apr 2012 B2
8167548 Greim May 2012 B2
8167567 Kirchner et al. May 2012 B2
8205432 Sheridan Jun 2012 B2
8246292 Morin et al. Aug 2012 B1
RE43710 Spear et al. Oct 2012 E
8382438 Guemmer Feb 2013 B2
8393870 Nash et al. Mar 2013 B2
8464426 Kirchner et al. Jun 2013 B2
9574574 Gallagher Feb 2017 B2
20020141863 Liu Oct 2002 A1
20030086788 Chandraker May 2003 A1
20030163983 Seda et al. Sep 2003 A1
20040091353 Shahpar et al. May 2004 A1
20050031454 Doloresco et al. Feb 2005 A1
20050169761 Dube et al. Aug 2005 A1
20050254956 Dutton et al. Nov 2005 A1
20060210395 Schuster et al. Sep 2006 A1
20060222488 Fessou Oct 2006 A1
20060228206 Decker et al. Oct 2006 A1
20070041841 Walter et al. Feb 2007 A1
20070160478 Jarrah et al. Jul 2007 A1
20070201983 Arinci et al. Aug 2007 A1
20070243068 Wadia et al. Oct 2007 A1
20080003096 Kohli et al. Jan 2008 A1
20080101959 McRae et al. May 2008 A1
20080120839 Schilling May 2008 A1
20080148564 Burton et al. Jun 2008 A1
20080226454 Decker et al. Sep 2008 A1
20090056343 Suciu et al. Mar 2009 A1
20090226322 Clemen Sep 2009 A1
20090257866 Greim Oct 2009 A1
20090274554 Guemmer Nov 2009 A1
20090297355 Herr et al. Dec 2009 A1
20090304518 Kodama et al. Dec 2009 A1
20090317227 Grover et al. Dec 2009 A1
20100054946 Orosa et al. Mar 2010 A1
20100148396 Xie et al. Jun 2010 A1
20100215503 Myoren et al. Aug 2010 A1
20100232970 Murooka et al. Sep 2010 A1
20100254797 Grover et al. Oct 2010 A1
20100260609 Wood et al. Oct 2010 A1
20100331139 McCune Dec 2010 A1
20110081252 Li Apr 2011 A1
20110116917 Wang et al. May 2011 A1
20110135482 Nash et al. Jun 2011 A1
20110150660 Micheli Jun 2011 A1
20110159797 Beltman et al. Jun 2011 A1
20110206527 Harvey et al. Aug 2011 A1
20110225979 Hoeger et al. Sep 2011 A1
20110268578 Praisner et al. Nov 2011 A1
20110286850 Micheli et al. Nov 2011 A1
20110286856 Micheli Nov 2011 A1
20110293423 Bunker et al. Dec 2011 A1
20120057982 O'Hearn et al. Mar 2012 A1
20120124964 Hasel et al. May 2012 A1
20120195767 Gervais et al. Aug 2012 A1
20120237344 Wood et al. Sep 2012 A1
20120243975 Breeze-Stringfellow et al. Sep 2012 A1
20120244005 Breeze-Stringfellow et al. Sep 2012 A1
20130008170 Gallagher et al. Jan 2013 A1
20130022473 Tran Jan 2013 A1
20130089415 Brown et al. Apr 2013 A1
20130149108 Webster Jun 2013 A1
20130164488 Wood et al. Jun 2013 A1
20130189117 Baltas et al. Jul 2013 A1
20130192199 Merry et al. Aug 2013 A1
20130192261 Mayer et al. Aug 2013 A1
20130192266 Houston Aug 2013 A1
20130202403 Morin et al. Aug 2013 A1
20130219859 Suciu Aug 2013 A1
20130219922 Gilson et al. Aug 2013 A1
20130224040 Straccia Aug 2013 A1
20130259668 Myoren et al. Oct 2013 A1
20130266451 Pesteil et al. Oct 2013 A1
20130315739 Cellier Nov 2013 A1
20130340406 Gallagher et al. Dec 2013 A1
20140030060 Magowan Jan 2014 A1
20140248155 Merville et al. Sep 2014 A1
20150017012 Pouzadoux et al. Jan 2015 A1
20150118059 Perrot Apr 2015 A1
20150354367 Gallagher et al. Dec 2015 A1
20160195104 Cellier Jul 2016 A1
Foreign Referenced Citations (34)
Number Date Country
1903642 Aug 1970 DE
3251978 Jan 1988 EP
3661413 Jul 1995 EP
3745755 Dec 1996 EP
1074700 Feb 2001 EP
1098092 May 2001 EP
1106835 Jun 2001 EP
1106836 Jun 2001 EP
1111188 Jun 2001 EP
1505302 Feb 2005 EP
1508669 Feb 2005 EP
1524405 Apr 2005 EP
1582695 Oct 2005 EP
1939399 Jul 2008 EP
0801230 May 2009 EP
2075408 Jul 2009 EP
2133573 Dec 2009 EP
2226468 Sep 2010 EP
2226468 Sep 2010 EP
1930598 Aug 2012 EP
2535527 Dec 2012 EP
2543818 Jan 2013 EP
2543818 Jan 2013 EP
2631491 Aug 2013 EP
1516041 Jun 1978 GB
2041090 Sep 1980 GB
2170868 Aug 1986 GB
2007001389 Jan 2007 WO
2007038674 Apr 2007 WO
2008109036 Sep 2008 WO
2009103528 Aug 2009 WO
WO 2013141935 Sep 2013 WO
2014066503 May 2014 WO
2015126449 Aug 2015 WO
Non-Patent Literature Citations (146)
Entry
The International Search Report and Written Opinion for PCT Application No. PCT/US2015/016187, mailed May 20, 2015.
The International Search Report and Written Opinion for PCT Application No. PCT/US2015/016011, mailed May 21, 2015.
The International Search Report and Written Opinion for PCT Application No. PCT/US2015/016078, mailed May 29, 2015.
The International Search Report and Written Opinion for PCT Application No. PCT/US2015/016154, mailed May 22, 2015.
The International Search Report and Written Opinion for PCT Application No. PCT/US2015/016086, mailed May 26, 2015.
The International Search Report and Written Opinion for PCT Application No. PCT/US2015/016554, mailed May 26, 2015.
The International Search Report and Written Opinion for PCT Application No. PCT/US2015/015554, mailed May 21, 2015.
The International Search Report and Written Opinion for PCT Application No. PCT/US2014/052325, mailed May 29, 2015.
The International Search Report and Written Opinion for PCT Application No. PCT/US2015/016378, mailed May 29, 2015.
The International Search Report and Written Opinion for PCT Application No. PCT/US2014/052293, mailed May 29, 2015.
The International Search Report and Written Opinion for PCT Application No. PCT/US2014/052516, mailed Jun. 10, 2015.
InternationalSearch Report and Written Opinion for PCT Application PCT/US2014/052447 dated Dec. 8, 2014.
International Search Report and Written Opinion for PCT Application PCT/US2014/052441 dated Dec. 12, 2014.
International Search Report and Written Opinion for PCT Application PCT/US2014/052282 dated Dec. 5, 2014.
International Search Report and Written Opinion for PCT Application PCT/US2014/052474 dated Dec. 5, 2014.
International Search Report and Written Opinion for PCT Application PCT/US2014/052434 dated Nov. 27, 2014.
International Search Report and Written Opinion for PCT Application PCT/US2014/052468 dated Dec. 12, 2014.
InternationalSearch Report and Written Opinion for PCT Application PCT/US2015/016083 dated Jul. 21, 2015.
International Search Report and Written Opinion for PCT Application PCT/US2014/052440 dated Nov. 27, 2014.
International Search Report and Written Opinion for PCT Application PCT/US2014/052437 dated Dec. 26, 2014.
International Search Report and Written Opinion for PCT Application PCT/US2014/052238 dated Dec. 11, 2014.
InternationalSearch Report and Written Opinion for PCT Application PCT/US2014/052080 dated Aug. 21, 2014.
InternationalSearch Report and Written Opinion for PCT Application PCT/US2014/052096 dated Nov. 28, 2014.
McMillian, A. (2008) Material development for fan blade containment casing. Abstract. p. 1. Conference on Engineering and Physics: Synergy for Success 2006. Journal of Physics: Conference Series vol. 105. London, UK. Oct. 5, 2006.
Kurzke, J. (2009). Fundamental differences between conventional and geared turbofans. Proceedings of ASME Turbo Expo: Power for Land, Sea, and Air. 2009, Orlando, Florida. pp. 145-153.
Agarwal, B.D and Broutman, L.J. (1990). Analysis and performance of fiber composites, 2nd Edition. John Wiley & Sons, Inc. New York: New York. pp. 1-30, 50-1, 56-8, 60-1, 64-71, 87-9, 324-9, 436-7.
Carney, K., Pereira, M. Revilock, and Matheny, P. (2003). Jet engine fan blade containment using two alternate geometries. 4th European LS-DYNA Users Conference. pp. 1-10.
Brines, G.L. (1990). The turbofan of tomorrow. Mechanical Engineering: The Journal of the American Society of Mechanical Engineers,108(8), 65-67.
Faghri, A. (1995). Heat pipe and science technology. Washington, D.C.: Taylor & Francis. pp. 1-60.
Hess, C. (1998). Pratt & Whitney develops geared turbofan. Flug Revue 43(7). Oct. 1998.
Grady, J.E., Weir, D.S., Lamoureux, M.C., and Martinez, M.M. (2007). Engine noise research in NASA's quiet aircraft technology project. Papers from the International Symposium on Air Breathing Engines (ISABE). 2007.
Griffiths, B. (2005). Composite fan blade containment case. Modem Machine Shop. Retrieved from: http://www.mmsonline.com/articles/composite-fan-blade-containment-case pp. 1-4.
Hall, C.A. and Crichton, D. (2007). Engine design studies for a silent aircraft. Journal of Turbomachinery, 129, 479-487.
Haque, A. and Shamsuzzoha, M., Hussain, F., and Dean, D. (2003). S20-glass/epoxy polymer nanocomposites: Manufacturing, structures, thermal and mechanical properties. Journal of Composite Materials, 37(20), 1821-1837.
Brennan, P.J. and Kroliczek, E.J. (1979). Heat pipe design handbook. Prepared for National Aeronautics and Space Administration by B & K Engineering, Inc. Jun. 1979. pp. 1-348.
Horikoshi, S. and Serpone, N. (2013). Introduction to nanoparticles. Microwaves in nanoparticle synthesis. Wiley-VCH Verlag GmbH & Co. KGaA. pp. 1-24.
Kerrebrock, J.L. (1977). Aircraft engines and gas turbines. Cambridge, MA: The MIT Press. p. 11.
Xie, M. (2008). Intelligent engine systems: Smart case system. NASA/CR-2008-215233. pp. 1-31.
Knip, Jr., G. (1987). Analysis of an advanced technology subsonic turbofan incorporating revolutionary materials. NASA Technical Memorandum. May 1987. pp. 1-23.
Willis, W.S. (1979). Quiet clean short-haul experimental engine (QCSEE) final report. NASA/CR-159473 pp. 1-289.
Kojima, Y., Usuki, A. Kawasumi, M., Okada, A., Fukushim, Y., Kurauchi, T., and Kamigaito, O. (1992). Mechanical properties of nylon 6-clay hybrid. Journal of Materials Research, 8(5), 1185-1189.
Kollar, L.P. and Springer, G.S. (2003). Mechanics of composite structures. Cambridge, UK: Cambridge University Press. p. 465.
Ramsden, J.M. (Ed). (1978). The new European airliner. Flight International, 113(3590). Jan. 7, 1978. pp. 39-43.
Langston, L. and Faghri, A. Heat pipe turbine vane cooling. Prepared for Advanced Turbine Systems Annual Program Review. Morgantown, West Virginia. Oct. 17-19, 1995. pp. 3-9.
Oates, G.C. (Ed). (1989). Aircraft propulsion systems and technology and design. Washington, D.C.: American Institute of Aeronautics, Inc. pp. 341-344.
Lau, K., Gu, C., and Hui, D. (2005). A critical review on nanotube and nanotube/nanoclay related polymer composite materials. Composites: Part B 37(2006) 425-436.
Shorter Oxford English dictionary, 6th Edition. (2007). vol. 2, N-Z. p. 1888.
Lynwander, P. (1983). Gear drive systems: Design and application. New York, New York: Marcel Dekker, Inc. pp. 145, 355-358.
Sweetman, B. and Sutton, O. (1998). Pratt & Whitney's surprise leap. Interavia Business & Technology, 53.621, p. 25.
Mattingly, J.D. (1996). Elements of gas turbine propulsion. New York, New York: McGraw-Hill, Inc. pp. 8-15.
Pyrograf-III Carbon Nanofiber. Product guide. Retrieved Dec. 1, 2015 from: http://pyrografproducts.com/Merchant5/merchant.mvc?Screen=cp—nanofiber.
Nanocor Technical Data for Epoxy Nanocomposites using Nanomer 1.30E Nanoclay. Nnacor, Inc. Oct. 2004.
Ratna, D. (2009). Handbook of thermoset resins. Shawbury, UK: iSmithers. pp. 187-216.
Wendus, B.E., Stark, D.F., Holler, R.P., and Funkhouser, M.E. (2003). Follow-on technology requirement study for advanced subsonic transport. NASA/CR-2003-212467. pp. 1-37.
Silverstein, C.C., Gottschlich, J.M., and Meininger, M. The feasibility of heat pipe turbine vane cooling. Presented at the International Gas Turbine and Aeroengine Congress and Exposition, The Hague, Netherlands. Jun. 13-16, 1994.pp. 1-7.
Merriam-Webster's collegiate dictionary, 11th Ed. (2009). p. 824.
Merriam-Webster's collegiate dictionary, 10th Ed. (2001). p. 1125-1126.
Whitaker, R. (1982). ALF 502: plugging the turbofan gap. Flight International, p. 237-241, Jan. 30, 1982.
Hughes, C. (2010). Geared turbofan technology. NASA Environmentally Responsible Aviation Project. Green Aviation Summit. NASA Ames Research Center. Sep. 8-9, 2010. pp. 1-8.
Gliebe, P.R. and Janardan, B.A. (2003). Ultra-high bypass engine aeroacoustic study. NASA/CR-2003-21252. GE Aircraft Engines, Cincinnati, Ohio. Oct. 2003. pp. 1-103.
Moxon, J. How to save fuel in tomorrow's engines. Flight International. Jul. 30, 1983. 3873(124). pp. 272-273.
File History for U.S. Appl. No. 12/131,876.
Cusick, M. (1981). Avco Lycoming's ALF 502 high bypass fan engine. Society of Automotive Engineers, inc. Business Aircraft Meeting & Exposition. Wichita, Kansas. Apr. 7-10, 1981. pp. 1-9.
Fledderjohn, K.R. (1983). The TFE731-5: Evolution of a decade of business jet service. SAE Technical Paper Series. Business Aircraft Meeting & Exposition. Wichita, Kansas. Apr. 12-15, 1983. pp. 1-12.
Dickey, T.A. and Dobak, E.R. (1972). The evolution and development status of ALF 502 turbofan engine. National Aerospace Engineering and Manufacturing Meeting. San Diego, California. Oct. 2-5, 1972. pp. 1-12.
Gunston, B. (Ed.) (2000). Jane's aero-engines, Issue seven. Coulsdon, Surrey, UK: Jane's Information Group Limited. pp. 510-512.
Ivchenko-Progress D-436. Jane's Aero-engines, Aero-engines- Turbofan. Feb. 8, 2012.
Ivchenko-Progress AI-727M. Jane's Aero-engines, Aero-engines- Turbofan. Nov. 27, 2011.
Ivchenko-Progress D-727. Jane's Aero-engines, Aero-engines- Turbofan. Feb. 7, 2007.
Turbomeca Aubisque. Jane's Aero-engines, Aero-engines- Turbofan. Nov. 2, 2009.
Aviadvigatel D-110. Jane's Aero-engines, Aero-engines- Turbofan. Jun. 1, 2010.
Rolls-Royce M45H. Jane's Aero-engines, Aero-engines- Turbofan. Feb. 24, 2010.
Honeywell LF502. Jane's Aero-engines, Aero-engines- Turbofan. Feb. 9, 2012.
Honeywell LF507. Jane's Aero-engines, Aero-engines- Turbofan. Feb. 9, 2012.
Honeywell TFE731. Jane's Aero-engines, Aero-engines- Turbofan. Jul. 18, 2012.
NASA Conference Publication. Quiet, powered-lift propulsion. Cleveland, Ohio. Nov. 14-15, 1978. pp. 1-420.
“Civil Turbojet/Turbofan Specifications”, Jet Engine Specification Database (Apr. 3, 2005).
Kandebo, S.W. (1993). Geared-turbofan engine design targets cost, complexity. Aviation Week & Space Technology, 148(8). Start p. 32.
Hendricks, E.S. and Tong, M.T. (2012). Performance and weight estimates for an advanced open rotor engine. NASA/TM-2012-217710. pp. 1-13.
Guynn, M. D., Berton, J.J., Fisher, K. L., Haller, W.J., Tong, M. T., and Thurman, D.R. (2011). Refined exploration of turbofan design options for an advanced single-aisle transport. NASA/TM-2011-216883. pp. 1-27.
Zalud, T. (1998). Gears put a new spin on turbofan performance. Machine Design, 70(20), p. 104.
Kurzke, J. (2008). Preliminary Design, Aero-engine design: From state of the art turbofans towards innovative architectures. pp. 1-72.
Zamboni, G. and Xu, L. (2009). Fan root aerodynamics for large bypass gas turbine engines: Influence on the engine performance and 3D design. Proceedings of ASME Turbo Expo 2009: Power for Land, Sea and Air. Jun. 8-12, 2009, Orlando, Florida, USA. pp. 1-12.
Han, J., Dutta, S., and Ekkad, S.V. (2000). Gas turbine heat transfer and cooling technology. New York, NY: Taylor & Francis. pp. 1-25, 129-157, and 160-249.
Mattingly, J.D. (1996). Elements of gas turbine propulsion. New York, New York: McGraw-Hill, Inc. pp. 1-18, 50-62, 85-87, 95-104, 121-123, 223-234, 242-245, 278-280, 303-309, 323-326, 462-479, 517-520, 563-565, 673-675, 582-685, 697-699, 703-705, 802-805, 862-864, and 923-925.
Declaration of Reza Abhari, Ph.D. In re U.S. Pat. No. 8,844,265. Executed Jun. 28, 2016. pp. 1-91.
Declaration of John Eaton, Ph.D. In re U.S. Pat. No. 8,869,568. Executed Mar. 28, 2016. pp. 1-87.
Declaration of Reza Abhari. In re U.S. Pat. No. 8,695,920. Executed Nov. 30. pp. 1-67.
Declaration of Reza Abhari. In re U.S. Pat. No. 8,448,895. Executed Nov. 28. pp. 1-81.
Declaration of Reza Abhari. In re U.S. Pat. No. 8,695,920, claims 1-4, 7-14, 17 and 19. Executed Nov. 29. pp. 1-102.
Declaration of Dr. Magdy Attia. In re U.S. Pat. No. 8,313,280. Executed Oct. 21, 2016. pp. 1-88.
Lord, W.K., MacMartin, D.G., and Tillman, T.G. (2000). Flow control opportunities in gas turbine engines. American Institute of Aeronautics and Astronautics. pp. 1-15.
Daly, M. Ed. (2010). Jane's Aero-Engine. Issue Twenty-seven. Mar. 2010. p. 633-636.
Roux, E. (2007). Turbofan and turbojet engines database handbook. Editions Elodie Roux. Blagnac: France. pp. 1-595.
Wilfert, G. (2008). Geared fan. Aero-Engine Design: From State of the Art Turbofans Towards Innovative Architectures, von Karman Institute for Fluid Dynamics, Belgium, Mar. 3-7, 2008. pp. 1-26.
Petition for Inter Partes Review of U.S. Pat. No. 8,517,668. General Electric Company, Petitioner, v. United Technologies Corporation, Patent Owner. Filed Dec. 14, 2016.
Declaration of Dr. Magdy Attia. In re U.S. Pat. No. 8,517,668. Executed Dec. 8, 2016. pp. 1-81.
Cramoisi, G. Ed. (2012). Death in the Potomac: The crash of Air Florida Flight 90. Air Crash Investigations. Accident Report NTSB/AAR-82-8. p. 45-47.
Norton, M. and Karczub, D. (2003). Fundamentals of noise and vibration analysis for engineers. Press Syndicate of the University of Cambridge. New York: New York. p. 524.
U.S. Department of Transportation: Federal Aviation Administration Advisory Circular. Runway overrun prevention Dated: Nov. 6, 2007. p. 1-8 and Appendix 1 p. 1-15, Appendix 2 p. 1-6, Appendix 3 p. 1-3, and Appendix 4 p. 1-5.
U.S. Department of Transportation: Federal Aviation Administration Advisory Circular. Standard operating procedures for flight deck crewmembers. Dated: Feb. 27, 2003.. p. 1-6 and Appendices.
Smith, L.,Yeh,H.,(1963).Sweep and Dihedral Effects in Axial-Flow Turbomachinery;Journal of Basic Engineering; Sep. 1963.pp. 401-416.
Engine Specifications. Engine Alliance GP7200—The Engine for the A380. Retrieved Feb. 19, 2015 from http://www.enginealliance.com/engine—specifications.html.
International Search Report and Written Opinion for PCT Application No. PCT/US2015/016018, mailed Nov. 24, 2015.
International Search Report and Written Opinion for PCT Application No. PCT/US2015016091, mailed Nov. 24, 2015.
International Search Report and Written Opinion for PCT Application No. PCT/US2015/016032, mailed Nov. 24, 2015.
International Search Report and Written Opinion for PCT Application No. PCT/US2015/016135, mailed Nov. 24, 2015.
International Search Report and Written Opinion for PCT Application No. PCT/US2015/016584, mailed Nov. 24, 2015.
International Search Report and Written Opinion for PCT Application No. PCT/US2015/015561, mailed Nov. 24, 2015.
International Search Report and Written Opinion for PCT Application No. PCT/US2015/015575, mailed Nov. 24, 2015.
International Search Report and Written Opinion for PCT Application No. PCT/US2015/015579, mailed Nov. 24, 2015.
International Search Report and Written Opinion for PCT Application No. PCT/US2015/015586, mailed Nov. 24, 2015.
Aerodynamic Design technique for Optimizing Fan Blade Spacing, Rogalsky et all., Proceedings of the 7th Annual Conference of the Computational Fluid Dynamics Society of Canada, 1999.
Turbine Design and Application. vol. 2. NASA, 1973.
Analytical Parametric Investigation of Low Pressure Ration Fan, NASA, 1973 Metzger et al.
Oyama et al., Multiobjective Optimization of a Multi-Stage Compressor Using Evolutionary Algorithm, NASA, 2002, AIAA 2002-3535 pp. 1-11.
International Preliminary Report on Patentability for PCT Application No. PCT/US2014/052282, mailed Sep. 1, 2016.
International Preliminary Report on Patentability for PCT Application No. PCT/US2015/016554, mailed Sep. 1, 2016.
International Preliminary Report on Patentability for PCT Application No. PCT/US2014/052434, mailed Sep. 1, 2016.
International Preliminary Report on Patentability for PCT Application No. PCT/US2014/052516, mailed Sep. 1, 2016.
International Preliminary Report on Patentability for PCT Application No. PCT/US2014/052447, mailed Sep. 1, 2016.
International Preliminary Report on Patentability for PCT Application No. PCT/US2015/015579, mailed Sep. 1, 2016.
International Preliminary Report on Patentability for PCT Application No. PCT/US2015/015586, mailed Sep. 1, 2016.
International Preliminary Report on Patentability for PCT Application No. PCT/US2014/052080, mailed Sep. 1, 2016.
International Preliminary Report on Patentability for PCT Application No. PCT/US2015/016135, mailed Sep. 1, 2016.
International Preliminary Report on Patentability for PCT Application No. PCT/US2015/016032 mailed Sep. 1, 2016.
International Preliminary Report on Patentability for PCT Application No. PCT/US2015/015561 mailed Sep. 1, 2016.
EP Search Report dated Jan. 24, 2017 for European Application No. 14883154.8.
EP Search Report dated Jan. 23, 2017 for European Application No. 14883117.5.
EP Search Report dated Jan. 24, 2017 for European Application No. 15752432.3.
EP Search Report dated Jan. 26, 2017 for European Application No. 15793425.8.
Partial EP Search Report dated Feb. 8, 2017 for European Application No. 15793193.2.
EP Search Report dated Feb. 9, 2017 for European Application No. 15752887.8.
EP Search Report dated Feb. 9, 2017 for European Application No. 14883515.0.
EP Search Report dated Jan. 30, 2017 for European Application No. 15752124.6.
EP Search Report dated Jan. 30, 2017 for European Application No. 15751498.5.
EP Search Report dated Feb. 3, 2017 for European Application No. 15751454.8.
EP Search Report dated Feb. 3, 2017 for European Application No. 15793323.5.
EP Search Report dated Feb. 3, 2017 for European Application No. 15796827.2.
EP Search Report dated Feb. 13, 2017 for European Application No. 15792194.1.
EP Search Report dated Feb. 13, 2017 for European Application No. 15751738.4.
EP Search Report dated Feb. 13, 2017 for European Application No. 15752593.2.
EP Search Report dated Feb. 13, 2017 for European Application No. 14883036.7.
EP Search Report dated Feb. 22, 2017 for European Application No. 15793112.2.
EP Search Report dated Feb. 20, 2017 for European Application No. 15793268.2.
Extended EP Search report for EP Application No. 15793193.2 dated May 12, 2017.
Related Publications (1)
Number Date Country
20170130586 A1 May 2017 US
Provisional Applications (1)
Number Date Country
61942025 Feb 2014 US
Continuations (3)
Number Date Country
Parent 15191860 Jun 2016 US
Child 15418115 US
Parent 14876995 Oct 2015 US
Child 15191860 US
Parent 14624025 Feb 2015 US
Child 14876995 US