Patents by Inventor Bruce L. Morin
Bruce L. Morin has filed for patents to protect the following inventions. This listing includes patent applications that are pending as well as patents that have already been granted by the United States Patent and Trademark Office (USPTO).
-
Patent number: 9957840Abstract: A gas turbine engine has a propulsor including a fan and a liner positioned upstream of the fan. The liner has a backing plate, a cellular structure with cells extending from the backing plate, and a perforated sheet with a depth defined as a distance between the perforated sheet and the backing sheet. The depth is selected to achieve a desired ratio of the depth relative to a gap?. A depth to gap ratio is substantially in a range of 0.035 to 0.08. A method is also disclosed.Type: GrantFiled: February 19, 2014Date of Patent: May 1, 2018Assignee: United Technologies CorporationInventors: David A. Topol, Bruce L. Morin, Dilip Prasad, Thomas J. Ouellette
-
Patent number: 9879611Abstract: Bleed valve assemblies in a gas turbine engine are disclosed herein. A bleed valve assembly in a low pressure compressor may include a bleed valve and a resonator chamber. A manifold may allow passage of air from a bleed duct into the resonator chamber. The resonator chamber may alter resonation properties of the bleed duct in order to prevent damage to components in the low pressure compressor.Type: GrantFiled: June 29, 2015Date of Patent: January 30, 2018Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Bruce L. Morin, David A. Topol, Mark Boyer, Brian Desfosses, W. Marshall Quin, John Holchin, Jonathan D. Little, Bryan Roseberry
-
Publication number: 20170343574Abstract: A gas turbine engine has a fan, a turbine section having a first turbine including a first turbine rotor, a compressor rotor, and a gear reduction that effects a reduction in a speed of the fan relative to an input speed from the first turbine rotor. Each of the compressor rotor and the first turbine rotor includes a number of blades in each of a plurality of blade rows, the number of blades configured to operate at least some of the time at a rotational speed, and the number of blades and the rotational speed being such that the following formula holds true for at least a plurality of the blade rows of the first turbine rotor: (number of blades×rotational speed)/60?5500, and the rotational speed being an approach speed in revolutions per minute, and the following formula holds true for at least a plurality of the blade rows of the compressor rotor: (number of blades×rotational speed)/60?10000, the rotational speed being an approach speed in revolutions per minute.Type: ApplicationFiled: July 28, 2017Publication date: November 30, 2017Inventors: David A. Topol, Bruce L. Morin, Detlef Korte
-
Publication number: 20170321611Abstract: A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor, and having a gear reduction ratio of greater than 2.5:1. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor: (number of blades×rotational speed)/60 sec?5500 Hz, and the rotational speed is in revolutions per minute. A method of designing a gas turbine engine also disclosed.Type: ApplicationFiled: February 20, 2017Publication date: November 9, 2017Applicant: MTU AERO ENGINES AGInventors: David A. Topol, Bruce L. Morin
-
Patent number: 9733266Abstract: A gas turbine engine has a fan, a turbine section having a first turbine including a first turbine rotor, a compressor rotor, and a gear reduction configured to effect a reduction in a speed of the fan relative to an input speed from the first turbine rotor. Each of the compressor rotor and the first turbine rotor includes a number of blades in each of a plurality of blade rows, the number of blades configured to operate at least some of the time at a rotational speed, and the number of blades and the rotational speed being such that the following formula holds true for at least a majority of the blade rows of the first turbine rotor, but does not hold true for any of the blade rows of the compressor rotor: (number of blades×rotational speed)/60?5500, and the rotational speed being an approach speed in revolutions per minute.Type: GrantFiled: September 20, 2016Date of Patent: August 15, 2017Assignees: UNITED TECHNOLOGIES CORPORATION, MTU AERO ENGINES AGInventors: David A. Topol, Bruce L. Morin, Detlef Korte
-
Patent number: 9726019Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan, a turbine section that has a fan drive turbine rotor, and a compressor rotor. A gear reduction effects a reduction in a speed of the fan relative to an input speed from the fan drive turbine rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of blade rows of the compressor rotor, and the blades are configured to operate at least some of the time at a rotational speed. The number of compressor blades in at least one of the blade rows and the rotational speed are such that the following formula holds true for the at least one of the plurality of blade rows of the compressor rotor: (the number of blades×the rotational speed)/60 sec?about 5500 Hz. A method of designing a gas turbine engine is also disclosed.Type: GrantFiled: September 8, 2016Date of Patent: August 8, 2017Assignees: UNITED TECHNOLOGIES CORPORATION, MTU AERO ENGINES AGInventors: David A. Topol, Bruce L. Morin
-
Publication number: 20170191415Abstract: A gas turbine engine has a fan, a turbine section having a first turbine including a first turbine rotor, a compressor rotor, and a gear reduction configured to effect a reduction in a speed of the fan relative to an input speed from the first turbine rotor. Each of the compressor rotor and the first turbine rotor includes a number of blades in each of a plurality of blade rows, the number of blades configured to operate at least some of the time at a rotational speed, and the number of blades and the rotational speed being such that the following formula holds true for at least a majority of the blade rows of the first turbine rotor, but does not hold true for any of the blade rows of the compressor rotor: (number of blades×rotational speed)/60?5500, and the rotational speed being an approach speed in revolutions per minute.Type: ApplicationFiled: September 20, 2016Publication date: July 6, 2017Inventors: David A. Topol, Bruce L. Morin, Detlef Korte
-
Publication number: 20170191424Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan, a turbine section that has a fan drive turbine rotor, and a compressor rotor. A gear reduction effects a reduction in a speed of the fan relative to an input speed from the fan drive turbine rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of blade rows of the compressor rotor, and the blades are configured to operate at least some of the time at a rotational speed. The number of compressor blades in at least one of the blade rows and the rotational speed are such that the following formula holds true for the at least one of the plurality of blade rows of the compressor rotor: (the number of blades×the rotational speed)/60 sec?about 5500 Hz. A method of designing a gas turbine engine is also disclosed.Type: ApplicationFiled: September 8, 2016Publication date: July 6, 2017Inventors: David A. Topol, Bruce L. Morin
-
Publication number: 20170184128Abstract: A gas turbine engine has a fan section including a fan. A turbine section has a first turbine and a second turbine. A gear reduction between the fan and the first turbine includes an epicycle gear train. The gear reduction is configured to receive an input from the first turbine and to turn the fan at a lower speed than the first turbine in operation. The first turbine further includes a number of turbine blades in each of a plurality of rows of the first turbine. The first turbine blades operate at least some of the time at a rotational speed. The number of blades and the rotational speed is such that the following formula holds true for at least one of the blade rows of the first turbine: (number of blades×speed)/60?5500. A turbine section is also disclosed.Type: ApplicationFiled: August 24, 2016Publication date: June 29, 2017Inventors: Bruce L. Morin, Detlef Korte
-
Patent number: 9650965Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan, a turbine section having a fan drive turbine rotor, and a compressor rotor. A gear reduction is configured to effect a reduction in a speed of the fan relative to an input speed from the fan drive turbine rotor. Each of the compressor rotor and the fan drive turbine rotor includes a number of blades in each of a plurality of blade rows. The number of blades are configured to operate at least some of the time at a rotational speed.Type: GrantFiled: February 3, 2016Date of Patent: May 16, 2017Assignees: United Technologies Corporation, MTU Aero Engines AGInventors: David A. Topol, Bruce L. Morin, Detlef Korte
-
Publication number: 20170122217Abstract: A gas turbine engine has a fan including a plurality of fan blades, a turbine section having a first turbine, a compressor, and a gear reduction positioned between the fan and the first turbine. Each of the compressor and the first turbine includes a number of blades in each of a plurality of blade rows, the number of blades rotatable at least some of the time at a rotational speed in operation, and the number of blades and the rotational speed being such that the following formula holds true for at least a majority of the blade rows of the first turbine, but does not hold true for any of the blade rows of the compressor: (number of blades×rotational speed)/60?5500, and the rotational speed being an approach speed in revolutions per minute.Type: ApplicationFiled: January 12, 2017Publication date: May 4, 2017Inventors: David A. Topol, Bruce L. Morin, Detlef Korte
-
Publication number: 20170122218Abstract: A gas turbine engine has a fan, a turbine section having a first turbine including a first turbine, a compressor, and a gear reduction positioned between the fan and the first turbine. Each of the compressor and the first turbine includes a number of blades in each of a plurality of blade rows, the number of blades rotatable at least some of the time at a rotational speed in operation, and the number of blades and the rotational speed being such that the following formula holds true for the plurality of the blade rows of the first turbine, but does not hold true for any of the blade rows of the compressor rotor: (number of blades×rotational speed)/60?5500, and the rotational speed being an approach speed in revolutions per minute.Type: ApplicationFiled: January 12, 2017Publication date: May 4, 2017Inventors: David A. Topol, Bruce L. Morin, Detlef Korte
-
Patent number: 9624834Abstract: A gas turbine engine has a fan and a turbine having a fan drive turbine rotor. The fan drive turbine rotor drives a compressor rotor. A gear reduction effects a reduction in the speed of the fan relative to an input speed from the fan drive turbine rotor that drives the compressor rotor. The compressor rotor has a number of compressor blades in at least one of a plurality of rows of the compressor rotor. The blades operate at least some of the time at a rotational speed. The number of compressor blades in at least one row and the rotational speed are such that the following formula holds true for at least one row of the compressor rotor turbine: (number of blades×rotational speed)/60 s?5500 Hz, and the rotational speed is in revolutions per minute. A method of designing a gas turbine engine and a compressor module are also disclosed.Type: GrantFiled: January 8, 2015Date of Patent: April 18, 2017Assignee: United Technologies CorporationInventors: David A. Topol, Bruce L. Morin
-
Publication number: 20160362983Abstract: A method of designing a gas turbine engine comprises the steps of including a fan section with a fan. A turbine section is included having a first turbine and a second turbine. A gear reduction is included between the fan and the first turbine, the gear reduction being configured to receive an input from the first turbine and to turn the fan at a lower speed than the first turbine in operation. The first turbine is designed to include a number of turbine blades in each of a plurality of rows of the first turbine, the first turbine blades operating at least some of the time at a rotational speed, and the number of blades and the rotational speed being such that the following formula holds true for at least one of the blade rows of the first turbine: (number of blades×speed)/60?5500.Type: ApplicationFiled: August 24, 2016Publication date: December 15, 2016Inventors: Bruce L. Morin, Detlef Korte
-
Publication number: 20160195021Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan, a turbine section having a fan drive turbine rotor, and a compressor rotor. A gear reduction is configured to effect a reduction in a speed of the fan relative to an input speed from the fan drive turbine rotor. Each of the compressor rotor and the fan drive turbine rotor includes a number of blades in each of a plurality of blade rows. The number of blades are configured to operate at least some of the time at a rotational speed.Type: ApplicationFiled: February 3, 2016Publication date: July 7, 2016Inventors: David A. Topol, Bruce L. Morin
-
Publication number: 20160177774Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.Type: ApplicationFiled: January 27, 2015Publication date: June 23, 2016Applicant: MTU AERO ENGINES GMBHInventors: Bruce L. Morin, David A. Topol, Detlef Korte
-
Publication number: 20160153361Abstract: Bleed valve assemblies in a gas turbine engine are disclosed herein. A bleed valve assembly in a low pressure compressor may include a bleed valve and a resonator chamber. A manifold may allow passage of air from a bleed duct into the resonator chamber. The resonator chamber may alter resonation properties of the bleed duct in order to prevent damage to components in the low pressure compressor.Type: ApplicationFiled: June 29, 2015Publication date: June 2, 2016Applicant: UNITED TECHNOLOGIES CORPORATIONInventors: BRUCE L. MORIN, DAVID A. TOPOL, MARK BOYER, BRIAN DESFOSSES, W. MARSHALL QUIN, JOHN HOLCHIN, JONATHAN D. LITTLE, BRYAN ROSEBERRY
-
Publication number: 20160153279Abstract: A turbine module according to an example of the present disclosure includes, among other things, a fan drive rotor having a plurality of blade rows each including a number of blades. A majority of the blade rows are capable of rotating at a rotational speed, so that when measuring the rotational speed in revolutions per minute: (number of blades×the rotational speed)/60?about 5500 Hz.Type: ApplicationFiled: January 27, 2016Publication date: June 2, 2016Inventors: Bruce L. Morin, Detlef Korte
-
Publication number: 20160153310Abstract: A turbine section including a high pressure turbine, an intermediate pressure turbine and a fan drive turbine, the fan drive turbine driving a gear reduction to in turn drive a fan, and effecting a reduction in the speed of the fan relative to an input speed from the fan drive turbine and said high pressure turbine driving a high pressure compressor, and the intermediate pressure turbine driving a low pressure compressor, with the intermediate pressure turbine having a number of turbine blades in at least one row, and the turbine blades operating at least some of the time at a rotational speed, and the number of turbine blades in the at least one row, and the rotational speed being such that the following formula holds true for the at least one row of the intermediate pressure turbine: (number of blades×speed)/60?5500 Hz.Type: ApplicationFiled: July 2, 2014Publication date: June 2, 2016Applicant: UNITED TECHNOLOGIES CORPORATIONInventors: Frederick M. Schwarz, Bruce L. Morin
-
Publication number: 20160138474Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan and a turbine section having a fan drive turbine rotor, and a compressor rotor. A gear reduction effects a reduction in a speed of the fan relative to an input speed from the fan drive turbine rotor. The compressor rotor has a number of compressor blades in at least half of a plurality of blade rows of the compressor rotor. The blades are configured to operate at least some of the time at a rotational speed. The number of compressor blades in the at least half of the blade rows and the rotational speed is such that the following formula holds true for each row of the at least half of the blade rows of the compressor rotor: (the number of blades×the rotational speed)/60 s?about 5500 Hz.Type: ApplicationFiled: December 14, 2015Publication date: May 19, 2016Inventors: David A. Topol, Bruce L. Morin