US10385774B2 - Split compressor turbine engine - Google Patents

Split compressor turbine engine Download PDF

Info

Publication number
US10385774B2
US10385774B2 US15/268,713 US201615268713A US10385774B2 US 10385774 B2 US10385774 B2 US 10385774B2 US 201615268713 A US201615268713 A US 201615268713A US 10385774 B2 US10385774 B2 US 10385774B2
Authority
US
United States
Prior art keywords
compressor
turbine
turbine engine
engine
combustor
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active, expires
Application number
US15/268,713
Other versions
US20180080373A1 (en
Inventor
Daniel Bernard Kupratis
Tania Bhatia Kashyap
Mark F. Zelesky
Arthur W. Utay
Gabriel L. Suciu
Thomas N. Slavens
Kevin L. Rugg
Brian Merry
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
United Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corp filed Critical United Technologies Corp
Priority to US15/268,713 priority Critical patent/US10385774B2/en
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SLAVENS, Thomas N., Kashyap, Tania Bhatia, KUPRATIS, Daniel Bernard, MERRY, BRIAN, RUGG, KEVIN L., SUCIU, GABRIEL L., Utay, Arthur W., ZELESKY, MARK F.
Priority to EP17191922.8A priority patent/EP3296542B1/en
Publication of US20180080373A1 publication Critical patent/US20180080373A1/en
Application granted granted Critical
Publication of US10385774B2 publication Critical patent/US10385774B2/en
Assigned to RAYTHEON TECHNOLOGIES CORPORATION reassignment RAYTHEON TECHNOLOGIES CORPORATION CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: UNITED TECHNOLOGIES CORPORATION
Assigned to RAYTHEON TECHNOLOGIES CORPORATION reassignment RAYTHEON TECHNOLOGIES CORPORATION CORRECTIVE ASSIGNMENT TO CORRECT THE AND REMOVE PATENT APPLICATION NUMBER 11886281 AND ADD PATENT APPLICATION NUMBER 14846874. TO CORRECT THE RECEIVING PARTY ADDRESS PREVIOUSLY RECORDED AT REEL: 054062 FRAME: 0001. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF ADDRESS. Assignors: UNITED TECHNOLOGIES CORPORATION
Assigned to RTX CORPORATION reassignment RTX CORPORATION CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: RAYTHEON TECHNOLOGIES CORPORATION
Active legal-status Critical Current
Adjusted expiration legal-status Critical

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/107—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
    • F02C3/145—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant the combustion chamber being in the reverse flow-type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/40—Use of a multiplicity of similar components
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/30—Arrangement of components

Definitions

  • the present disclosure relates generally to turbine engines, and more specifically to a split compressor turbine engine.
  • Turbine engines generally compress air in a compressor section, and provide the compressed air to a combustor.
  • the compressed air is mixed with a fuel, and ignited within the combustor.
  • the resultant combustion products are passed to a turbine section, and are expanded across the turbine section.
  • the expansion of the combustion products drives rotation of the turbine section.
  • the turbine section is connected to the compressor section via one or more shafts, and the rotation of the turbine section, in turn, drives rotation of the compressor section.
  • the compressor section and the turbine section each include multiple compressors and turbines, respectively.
  • the first compressor referred to as a low pressure compressor, compresses ambient air, and provides the compressed air to the second compressor, referred to as the high pressure compressor.
  • This arrangement is referred to as the compressors being in series, and provides compressed air to the combustor section from a single output source in the compressor section.
  • a turbine engine in one exemplary embodiment includes a first compressor and a second compressor fluidly parallel to the first compressor, a reverse flow combustor fluidly connected to the first compressor and the second compressor, and a first turbine and a second turbine fluidly in series, and fluidly connected to an output of the reverse flow combustor.
  • a fluid inlet of the first compressor and a fluid inlet of the second compressor are approximately equal sized, such that fluid flow into each of the first compressor and the second compressor is approximately equal.
  • At least one of the first turbine and the second turbine is a single stage turbine.
  • each of the first turbine and the second turbine is a single stage turbine.
  • first compressor and the first turbine are connected to a first spool, and wherein the second compressor and the second turbine are connected to a second spool.
  • the first spool and the second spool are collinear.
  • the first compressor, the second compressor, the first turbine and the second turbine are connected to a single spool.
  • At least one of the first compressor and the second compressor is a direct drive compressor.
  • the first compressor and the second compressor are counter-rotating compressors.
  • the first compressor and the second compressor are co-rotating.
  • At least one of the first compressor and the second compressor is comprised of multiple rotors, each of the rotors being constructed of a lightweight high strength ceramic.
  • the lightweight high strength ceramic is a silicon based structural ceramic material.
  • the lightweight high strength ceramic comprises one of silicon nitride, silicon carbide, silicon carbide fiber reinforced ceramic composite, and carbon fiber reinforced silicon carbide composite.
  • An exemplary method for driving a turbine engine includes splitting an inlet flow between a first compressor and a second compressor, providing an output flow of each of the first compressor and the second compressor to a reverse flow combustor, and driving a first turbine and a second turbine to rotate by expanding combustion products generated in the reverse flow combustor across the first turbine and the second turbine.
  • Another example of any of the above described methods for driving a turbine engine further includes driving rotation of the first compressor via a shaft connecting the first compressor to the first turbine, and driving rotation of the second compressor via a shaft connecting the second compressor to the second turbine.
  • Another example of any of the above described methods for driving a turbine engine further includes driving rotation of the first compressor and the second compressor via a shaft connecting the first compressor and the second compressor to the first turbine and the second turbine.
  • a turbine engine in one exemplary embodiment includes a first compressor and a second compressor fluidly parallel to the first compressor, a combustor fluidly connected to the first compressor and the second compressor, and a turbine section comprising a first turbine and a second turbine downstream of the first turbine, the turbine section being fluidly connected to an output of the combustor.
  • each of the first turbine and the second turbine are single stage turbines.
  • FIG. 1 schematically illustrates a split compressor turbine engine architecture according to a first example.
  • FIG. 2 schematically illustrates a split compressor turbine engine architecture according to a second example.
  • FIG. 3 illustrates a method for operating a gas turbine engine according to either of the examples of FIGS. 1 and 2 .
  • FIG. 1 schematically illustrates an exemplary split compressor turbine engine architecture 100 .
  • the engine includes a pair of split compressors 112 , 114 within a compressor section 110 .
  • the compressors 112 , 114 share an inlet 116 , and operate in fluid parallel, with the compressed air output being merged into a single compressed airflow 122 downstream of both compressors 112 , 114 .
  • the inlet 116 draws ambient air from a surrounding atmosphere through a first inlet 102 on a forward end of the engine 100 and through a second inlet 104 , on a radially outward surface of the engine 100 .
  • Airflow into the engine follows flowpath 120 , and branches into the first inlet 102 along a first branch 124 and into the second inlet 104 along a second branch 126 .
  • the first and second branch 124 , 126 merge at the compressor section 110 inlet 116 , and the flow is then split again between inlets of the first compressor 112 and the second compressor 114 .
  • the split is approximately 50%, with each compressor 112 , 114 receiving approximately the same volume of air along its respective flowpath as the other compressor 112 , 114 .
  • Such an example can be achieved by providing each of the compressors 112 , 114 approximately equal sized inlets, thereby ensure that an approximately equal volume of air will enter the compressors 112 , 114 .
  • the compressors can be sized such that different volumes of air are received at their inlets, or a controlled or passive metering device can be incorporated at the inlet 116 .
  • Each compressor 112 , 114 includes multiple compressor stages 118 that sequentially compress the air resulting in a higher pressure at the compressor outlet than at the compressor inlet 116 .
  • Each stage includes a compressor rotor and a corresponding compressor stator, with the rotors being shaped to drive air along the compressor as the rotors rotate.
  • the compressor rotors are constructed of lightweight, high strength materials, such as a lightweight high strength ceramic material.
  • the light weight high strength ceramic material is a silicon based structural material, such as silicon nitride, silicon carbide, silicon carbide reinforced ceramic composite, or carbon fiber reinforced silicon carbide composite.
  • the compressed airflow 122 is passed to a reverse flow combustor 130 , where the compressed air is mixed with a fuel and ignited according to known combustor techniques.
  • the resultant combustion products are passed along a combustion product flowpath 140 into a turbine section 150 .
  • each of the single stage turbines 152 , 154 includes a single rotor 151 , and a single stator vane 153 .
  • either or both of the turbines 152 , 154 within the turbine section 150 can include multiple turbine stages, instead of the illustrated single stage turbines 152 , 154 .
  • Each rotor 151 is connected to a corresponding shaft 160 , 170 .
  • the shafts 160 , 170 are alternately referred to as spools.
  • Each shaft 160 , 170 connects the turbine rotor 151 to a corresponding one of the compressors 112 , 114 , and drives the rotation of the corresponding compressor 112 , 114 .
  • the shafts 160 , 170 are collinear, with the shaft 160 that is connected to the forward turbine 154 being radially outward of the shaft 170 that is connected to the aft turbine 152 . While illustrated in FIG. 1 utilizing direct drive connections to the shafts 160 , 170 , one of skill in the art could adapt the engine architecture 100 to utilize a geared connection, and drive one or both of the compressors 112 , 114 via a geared connection.
  • the turbines 152 , 154 are sized such that rotation of the forward turbine 154 , and rotation of the aft turbine 152 , drive rotation of their corresponding compressors 112 , 114 at the same, or approximately the same, speed at any given time.
  • each of the compressors 112 , 114 are driven to rotate in the same direction about an engine centerline axis, and are referred to as co-rotating compressors 112 , 114 .
  • the compressors 112 , 114 rotate in opposite directions about the centerline axis, and are referred to as counter-rotating compressors 112 , 114 .
  • the turbine 152 , 154 corresponding to a given compressor 112 , 114 rotates in the same direction as the compressor 112 , 114 .
  • FIG. 2 schematically illustrates an alternate configuration split compressor turbine engine architecture 200 .
  • the architecture 200 includes two compressors 212 , 214 arranged in parallel with an inlet flow to a compressor section 210 being split between the compressors 212 , 214 .
  • the output of the compressor section 210 is provided to a reverse flow combustor 230 , where the compressed air from the compressor section 210 is mixed with a fuel and ignited.
  • the resultant combustion products are provided from the reverse flow combustor 230 to a turbine section 250 including a first turbine 254 and a second turbine 252 .
  • Each of the turbines 252 , 254 includes a single stage having a stator 253 and a rotor 251 .
  • the turbines 252 , 254 can include multiple stages and operate in a similar fashion.
  • Each of the turbines 252 , 254 is connected to a single shaft 260 .
  • the shaft 260 is, in turn, connected to both of the compressors 212 , 214 in either a direct drive (as illustrated) or a geared connection.
  • the shaft 260 translates rotation from the turbines 252 , 254 to the compressors 212 , 214 , thereby allowing the turbines 252 , 254 to drive rotation of the compressors 212 , 214 .
  • the engine architecture 200 of FIG. 2 operates, and is configured, in fundamentally the same manner as the engine architecture of FIG. 1 .
  • One of skill in the art, having the benefit of this disclosure will understand the necessary adjustments required to configure the engine architecture for a single shaft, as opposed to the two shaft example described above in greater detail.
  • FIG. 3 illustrates a method 300 for operating a split compressor turbine engine, such as the engine architectures 100 , 200 of FIGS. 1 and 2 .
  • a split compressor turbine engine such as the engine architectures 100 , 200 of FIGS. 1 and 2 .
  • an airflow is provided to a split inlet, and is divided between two parallel operating compressors within a compressor section in a “Split Inlet Flow to Compressors” step 310 .
  • the compressors operate in parallel to compress the air, and provide an output of compressed air.
  • the compressed air output from each compressor is rejoined into a single compressed airflow in a “Rejoin Compressed Airflow” step 320 .
  • the rejoined compressed air is provided to a reverse flow combustor and mixed with a fuel in the reverse flow combustor in a “Provide Compressed Air to Reverse Flow Combustor” step 330 .
  • the fuel/air mixture is ignited and the resultant combustion products are expelled from the reverse flow combustor according to known reverse flow combustor techniques.
  • the resultant combustion products are provided to a turbine section and expanded across multiple turbines within the turbine section in an “Expand Combustion Products Across Turbine” step 340 .
  • the expansion of the combustion products drives the turbines to rotate, and the rotation of the turbines is utilized to drive rotation of at least one corresponding compressor via a shaft connection.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

A turbine engine includes a first compressor and a second compressor fluidly parallel to the first compressor. A reverse flow combustor is fluidly connected to the first compressor and the second compressor. A first turbine and a second turbine are fluidly connected in series, and fluidly connected to an output of the reverse flow combustor.

Description

TECHNICAL FIELD
The present disclosure relates generally to turbine engines, and more specifically to a split compressor turbine engine.
BACKGROUND
Turbine engines generally compress air in a compressor section, and provide the compressed air to a combustor. The compressed air is mixed with a fuel, and ignited within the combustor. The resultant combustion products are passed to a turbine section, and are expanded across the turbine section. The expansion of the combustion products drives rotation of the turbine section. The turbine section is connected to the compressor section via one or more shafts, and the rotation of the turbine section, in turn, drives rotation of the compressor section.
In a typical example, the compressor section and the turbine section each include multiple compressors and turbines, respectively. The first compressor, referred to as a low pressure compressor, compresses ambient air, and provides the compressed air to the second compressor, referred to as the high pressure compressor. This arrangement is referred to as the compressors being in series, and provides compressed air to the combustor section from a single output source in the compressor section.
SUMMARY OF THE INVENTION
In one exemplary embodiment a turbine engine includes a first compressor and a second compressor fluidly parallel to the first compressor, a reverse flow combustor fluidly connected to the first compressor and the second compressor, and a first turbine and a second turbine fluidly in series, and fluidly connected to an output of the reverse flow combustor.
In another example of the above described turbine engine a fluid inlet of the first compressor and a fluid inlet of the second compressor are approximately equal sized, such that fluid flow into each of the first compressor and the second compressor is approximately equal.
In another example of any of the above described turbine engines at least one of the first turbine and the second turbine is a single stage turbine.
In another example of any of the above described turbine engines each of the first turbine and the second turbine is a single stage turbine.
In another example of any of the above described turbine engines the first compressor and the first turbine are connected to a first spool, and wherein the second compressor and the second turbine are connected to a second spool.
In another example of any of the above described turbine engines the first spool and the second spool are collinear.
In another example of any of the above described turbine engines the first compressor, the second compressor, the first turbine and the second turbine are connected to a single spool.
In another example of any of the above described turbine engines at least one of the first compressor and the second compressor is a direct drive compressor.
In another example of any of the above described turbine engines the first compressor and the second compressor are counter-rotating compressors.
In another example of any of the above described turbine engines the first compressor and the second compressor are co-rotating.
In another example of any of the above described turbine engines at least one of the first compressor and the second compressor is comprised of multiple rotors, each of the rotors being constructed of a lightweight high strength ceramic.
In another example of any of the above described turbine engines the lightweight high strength ceramic is a silicon based structural ceramic material.
In another example of any of the above described turbine engines the lightweight high strength ceramic comprises one of silicon nitride, silicon carbide, silicon carbide fiber reinforced ceramic composite, and carbon fiber reinforced silicon carbide composite.
An exemplary method for driving a turbine engine includes splitting an inlet flow between a first compressor and a second compressor, providing an output flow of each of the first compressor and the second compressor to a reverse flow combustor, and driving a first turbine and a second turbine to rotate by expanding combustion products generated in the reverse flow combustor across the first turbine and the second turbine.
In another example of the above described method for driving a turbine engine splitting an inlet flow between the first compressor and the second compressor, comprises splitting the inlet flow approximately evenly.
In another example of any of the above described methods for driving a turbine engine expanding the combustion products across the first turbine and the second turbine comprises expanding an output of the first turbine across the second turbine.
Another example of any of the above described methods for driving a turbine engine further includes driving rotation of the first compressor via a shaft connecting the first compressor to the first turbine, and driving rotation of the second compressor via a shaft connecting the second compressor to the second turbine.
Another example of any of the above described methods for driving a turbine engine further includes driving rotation of the first compressor and the second compressor via a shaft connecting the first compressor and the second compressor to the first turbine and the second turbine.
In one exemplary embodiment a turbine engine includes a first compressor and a second compressor fluidly parallel to the first compressor, a combustor fluidly connected to the first compressor and the second compressor, and a turbine section comprising a first turbine and a second turbine downstream of the first turbine, the turbine section being fluidly connected to an output of the combustor.
In another example of the above described turbine engine each of the first turbine and the second turbine are single stage turbines.
These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 schematically illustrates a split compressor turbine engine architecture according to a first example.
FIG. 2 schematically illustrates a split compressor turbine engine architecture according to a second example.
FIG. 3 illustrates a method for operating a gas turbine engine according to either of the examples of FIGS. 1 and 2.
 DETAILED DESCRIPTION OF AN EMBODIMENT
FIG. 1 schematically illustrates an exemplary split compressor turbine engine architecture 100. The engine includes a pair of split compressors 112, 114 within a compressor section 110. The compressors 112, 114 share an inlet 116, and operate in fluid parallel, with the compressed air output being merged into a single compressed airflow 122 downstream of both compressors 112, 114. The inlet 116 draws ambient air from a surrounding atmosphere through a first inlet 102 on a forward end of the engine 100 and through a second inlet 104, on a radially outward surface of the engine 100.
Airflow into the engine follows flowpath 120, and branches into the first inlet 102 along a first branch 124 and into the second inlet 104 along a second branch 126. The first and second branch 124, 126 merge at the compressor section 110 inlet 116, and the flow is then split again between inlets of the first compressor 112 and the second compressor 114. In some examples, the split is approximately 50%, with each compressor 112, 114 receiving approximately the same volume of air along its respective flowpath as the other compressor 112, 114. Such an example can be achieved by providing each of the compressors 112, 114 approximately equal sized inlets, thereby ensure that an approximately equal volume of air will enter the compressors 112, 114. In alternative examples, the compressors can be sized such that different volumes of air are received at their inlets, or a controlled or passive metering device can be incorporated at the inlet 116.
Each compressor 112, 114 includes multiple compressor stages 118 that sequentially compress the air resulting in a higher pressure at the compressor outlet than at the compressor inlet 116. Each stage includes a compressor rotor and a corresponding compressor stator, with the rotors being shaped to drive air along the compressor as the rotors rotate. In some examples, the compressor rotors are constructed of lightweight, high strength materials, such as a lightweight high strength ceramic material. In further examples, the light weight high strength ceramic material is a silicon based structural material, such as silicon nitride, silicon carbide, silicon carbide reinforced ceramic composite, or carbon fiber reinforced silicon carbide composite.
The compressed airflow 122 is passed to a reverse flow combustor 130, where the compressed air is mixed with a fuel and ignited according to known combustor techniques. The resultant combustion products are passed along a combustion product flowpath 140 into a turbine section 150.
Within the turbine section 150 are two single stage turbines 152, 154 arranged in fluid series. Each of the single stage turbines 152, 154 includes a single rotor 151, and a single stator vane 153. In alternative examples, either or both of the turbines 152, 154 within the turbine section 150 can include multiple turbine stages, instead of the illustrated single stage turbines 152, 154.
Each rotor 151 is connected to a corresponding shaft 160, 170. The shafts 160, 170 are alternately referred to as spools. Each shaft 160, 170 connects the turbine rotor 151 to a corresponding one of the compressors 112, 114, and drives the rotation of the corresponding compressor 112, 114. In the example of FIG. 1, the shafts 160, 170 are collinear, with the shaft 160 that is connected to the forward turbine 154 being radially outward of the shaft 170 that is connected to the aft turbine 152. While illustrated in FIG. 1 utilizing direct drive connections to the shafts 160, 170, one of skill in the art could adapt the engine architecture 100 to utilize a geared connection, and drive one or both of the compressors 112, 114 via a geared connection.
Further, in the example of FIG. 1, the turbines 152, 154 are sized such that rotation of the forward turbine 154, and rotation of the aft turbine 152, drive rotation of their corresponding compressors 112, 114 at the same, or approximately the same, speed at any given time.
In some examples, each of the compressors 112, 114 are driven to rotate in the same direction about an engine centerline axis, and are referred to as co-rotating compressors 112, 114. In alternative examples, the compressors 112, 114 rotate in opposite directions about the centerline axis, and are referred to as counter-rotating compressors 112, 114. In either example, the turbine 152, 154 corresponding to a given compressor 112, 114 rotates in the same direction as the compressor 112, 114.
With continued reference to FIG. 1, and with like numerals indicating like elements, FIG. 2 schematically illustrates an alternate configuration split compressor turbine engine architecture 200. As with the first example, the architecture 200 includes two compressors 212, 214 arranged in parallel with an inlet flow to a compressor section 210 being split between the compressors 212, 214.
The output of the compressor section 210 is provided to a reverse flow combustor 230, where the compressed air from the compressor section 210 is mixed with a fuel and ignited. The resultant combustion products are provided from the reverse flow combustor 230 to a turbine section 250 including a first turbine 254 and a second turbine 252. Each of the turbines 252, 254 includes a single stage having a stator 253 and a rotor 251. In alternative examples, the turbines 252, 254 can include multiple stages and operate in a similar fashion.
Each of the turbines 252, 254 is connected to a single shaft 260. The shaft 260 is, in turn, connected to both of the compressors 212, 214 in either a direct drive (as illustrated) or a geared connection. The shaft 260 translates rotation from the turbines 252, 254 to the compressors 212, 214, thereby allowing the turbines 252, 254 to drive rotation of the compressors 212, 214.
Aside from the alternate utilization of a single shaft 260, in place of the two shaft 260, 270 arrangement of FIG. 1, the engine architecture 200 of FIG. 2 operates, and is configured, in fundamentally the same manner as the engine architecture of FIG. 1. One of skill in the art, having the benefit of this disclosure will understand the necessary adjustments required to configure the engine architecture for a single shaft, as opposed to the two shaft example described above in greater detail.
With continued reference to FIGS. 1 and 2, FIG. 3 illustrates a method 300 for operating a split compressor turbine engine, such as the engine architectures 100, 200 of FIGS. 1 and 2. Initially, an airflow is provided to a split inlet, and is divided between two parallel operating compressors within a compressor section in a “Split Inlet Flow to Compressors” step 310.
The compressors operate in parallel to compress the air, and provide an output of compressed air. The compressed air output from each compressor is rejoined into a single compressed airflow in a “Rejoin Compressed Airflow” step 320.
The rejoined compressed air is provided to a reverse flow combustor and mixed with a fuel in the reverse flow combustor in a “Provide Compressed Air to Reverse Flow Combustor” step 330. The fuel/air mixture is ignited and the resultant combustion products are expelled from the reverse flow combustor according to known reverse flow combustor techniques.
The resultant combustion products are provided to a turbine section and expanded across multiple turbines within the turbine section in an “Expand Combustion Products Across Turbine” step 340. The expansion of the combustion products drives the turbines to rotate, and the rotation of the turbines is utilized to drive rotation of at least one corresponding compressor via a shaft connection.
It is further understood that any of the above described concepts can be used alone or in combination with any or all of the other above described concepts. Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.

Claims (21)

The invention claimed is:
1. A turbine engine comprising:
a first compressor and a second compressor in fluid parallel with the first compressor, each of the first compressor and the second compressor including multiple compressor stages;
a reverse flow combustor fluidly connected to said first compressor and said second compressor; and
a first turbine and a second turbine in fluid series, and fluidly connected to an output of the reverse flow combustor.
2. The turbine engine of claim 1, wherein a fluid inlet of the first compressor and a fluid inlet of the second compressor are equal sized, such that fluid flow into each of the first compressor and the second compressor is equal.
3. The turbine engine of claim 1, wherein at least one of said first turbine and said second turbine is a single stage turbine.
4. The turbine engine of claim 3, wherein each of said first turbine and said second turbine is a single stage turbine.
5. The turbine engine of claim 1, wherein said first compressor and said first turbine are connected to a first spool, and wherein said second compressor and said second turbine are connected to a second spool.
6. The turbine engine of claim 5, wherein said first spool and said second spool are collinear.
7. The turbine engine of claim 1, wherein said first compressor, said second compressor, said first turbine and said second turbine are connected to a single spool.
8. The turbine engine of claim 1, wherein at least one of said first compressor and said second compressor is a direct drive compressor.
9. The turbine engine of claim 1, wherein said first compressor and said second compressor are counter-rotating compressors, relative to each other.
10. The turbine engine of claim 1, wherein said first compressor and said second compressor are co-rotating.
11. The turbine engine of claim 1, wherein at least one of said first compressor and said second compressor is comprised of multiple rotors, each of said rotors being constructed of a lightweight high strength ceramic.
12. The turbine engine of claim 11, wherein the lightweight high strength ceramic is a silicon based structural ceramic material.
13. The turbine engine of claim 12, wherein the lightweight high strength ceramic comprises one of silicon nitride, silicon carbide, silicon carbide fiber reinforced ceramic composite, and carbon fiber reinforced silicon carbide composite.
14. A method for driving a turbine engine comprising:
splitting an inlet flow between a first compressor and a second compressor such that the second compressor is in fluid parallel with the first compressor, each of the first compressor and the second compressor including multiple compressor stages;
providing an output flow of each of said first compressor and said second compressor to a reverse flow combustor; and
driving a first turbine and a second turbine to rotate by expanding combustion products generated in said reverse flow combustor across the first turbine and the second turbine.
15. The method of claim 14, wherein splitting an inlet flow between the first compressor and the second compressor, comprises splitting the inlet flow evenly.
16. The method of claim 14, wherein expanding the combustion products across the first turbine and the second turbine comprises expanding an output of the first turbine across the second turbine.
17. The method of claim 14, further comprising driving rotation of the first compressor via a shaft connecting the first compressor to the first turbine, and driving rotation of the second compressor via a shaft connecting the second compressor to the second turbine.
18. The method of claim 14, further comprising driving rotation of the first compressor and the second compressor via a shaft connecting the first compressor and the second compressor to the first turbine and the second turbine.
19. A turbine engine comprising:
a first compressor and a second compressor in fluid parallel with the first compressor, each of the first compressor and the second compressor including multiple compressor stages;
a combustor fluidly connected to said first compressor and said second compressor; and
a turbine section comprising a first turbine and a second turbine downstream of the first turbine, the turbine section being fluidly connected to an output of the combustor.
20. The turbine engine of claim 19, wherein each of said first turbine and said second turbine are single stage turbines.
21. The turbine engine of claim 1, wherein the first turbine and the second turbine are sized such that the first compressor and the second compressor are driven to rotate at the same speed.
US15/268,713 2016-09-19 2016-09-19 Split compressor turbine engine Active 2037-10-26 US10385774B2 (en)

Priority Applications (2)

Application Number Priority Date Filing Date Title
US15/268,713 US10385774B2 (en) 2016-09-19 2016-09-19 Split compressor turbine engine
EP17191922.8A EP3296542B1 (en) 2016-09-19 2017-09-19 Split compressor turbine engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US15/268,713 US10385774B2 (en) 2016-09-19 2016-09-19 Split compressor turbine engine

Publications (2)

Publication Number Publication Date
US20180080373A1 US20180080373A1 (en) 2018-03-22
US10385774B2 true US10385774B2 (en) 2019-08-20

Family

ID=60019671

Family Applications (1)

Application Number Title Priority Date Filing Date
US15/268,713 Active 2037-10-26 US10385774B2 (en) 2016-09-19 2016-09-19 Split compressor turbine engine

Country Status (2)

Country Link
US (1) US10385774B2 (en)
EP (1) EP3296542B1 (en)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11002146B1 (en) 2020-10-26 2021-05-11 Antheon Research, Inc. Power generation system
US11530617B2 (en) 2020-10-26 2022-12-20 Antheon Research, Inc. Gas turbine propulsion system

Citations (26)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2405919A (en) * 1940-03-02 1946-08-13 Power Jets Res & Dev Ltd Fluid flow energy transformer
US2435836A (en) * 1944-12-13 1948-02-10 Gen Electric Centrifugal compressor
US2695499A (en) * 1949-08-22 1954-11-30 Power Jets Res & Dev Ltd Gas turbine power unit
US3163003A (en) * 1954-10-25 1964-12-29 Garrett Corp Gas turbine compressor
US3486328A (en) 1967-01-11 1969-12-30 Snecma Multiflow turbojet engine
DE1626033A1 (en) 1967-12-30 1970-08-20 Daimler Benz Ag Gas turbine plant
US3625003A (en) 1970-09-08 1971-12-07 Gen Motors Corp Split compressor gas turbine
US3779486A (en) * 1970-09-26 1973-12-18 Great Britain & Northern Irela Gas turbine engines
US3981616A (en) * 1974-10-22 1976-09-21 The United States Of America As Represented By The Secretary Of The Air Force Hollow composite compressor blade
US4171614A (en) 1976-04-17 1979-10-23 Motoren- Und Turbinen-Union Munchen Gmbh Gas turbine engine
US5131223A (en) 1989-07-13 1992-07-21 Sundstrand Corporation Integrated booster and sustainer engine for a missile
US5619854A (en) * 1994-05-25 1997-04-15 Gec Alsthom Diesels, Ltd. Turbocharged internal combustion engine
US7424805B2 (en) 2005-04-29 2008-09-16 General Electric Company Supersonic missile turbojet engine
US7628018B2 (en) * 2008-03-12 2009-12-08 Mowill R Jan Single stage dual-entry centriafugal compressor, radial turbine gas generator
US20100319343A1 (en) * 2009-06-23 2010-12-23 Arnold Steven D Turbocharger with two-stage compressor, including a twin-wheel parallel-flow first stage
US8176725B2 (en) * 2009-09-09 2012-05-15 United Technologies Corporation Reversed-flow core for a turbofan with a fan drive gear system
EP2551485A2 (en) 2011-07-26 2013-01-30 United Technologies Corporation Gas turbine engine with aft core driven fan section
US20130219860A1 (en) * 2012-02-29 2013-08-29 Gabriel L. Suciu Counter-rotating low pressure turbine without turbine exhaust case
US20130255224A1 (en) * 2012-03-27 2013-10-03 Daniel Bernard Kupratis Reverse core gear turbofan
US8745989B2 (en) 2009-04-09 2014-06-10 Pratt & Whitney Canada Corp. Reverse flow ceramic matrix composite combustor
US8935912B2 (en) * 2011-12-09 2015-01-20 United Technologies Corporation Gas turbine engine with variable overall pressure ratio
US9140212B2 (en) * 2012-06-25 2015-09-22 United Technologies Corporation Gas turbine engine with reverse-flow core having a bypass flow splitter
US9297311B2 (en) * 2011-03-22 2016-03-29 Alstom Technology Ltd Gas turbine power plant with flue gas recirculation and oxygen-depleted cooling gas
US20160169102A1 (en) * 2014-12-12 2016-06-16 United Technologies Corporation Reverse core flow gas turbine engine
US20160237894A1 (en) * 2015-02-13 2016-08-18 United Technologies Corporation Turbine engine with a turbo-compressor
US9460110B2 (en) 2011-09-21 2016-10-04 Kevin Mark Klughart File system extension system and method

Patent Citations (27)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2405919A (en) * 1940-03-02 1946-08-13 Power Jets Res & Dev Ltd Fluid flow energy transformer
US2435836A (en) * 1944-12-13 1948-02-10 Gen Electric Centrifugal compressor
US2695499A (en) * 1949-08-22 1954-11-30 Power Jets Res & Dev Ltd Gas turbine power unit
US3163003A (en) * 1954-10-25 1964-12-29 Garrett Corp Gas turbine compressor
US3486328A (en) 1967-01-11 1969-12-30 Snecma Multiflow turbojet engine
DE1626033A1 (en) 1967-12-30 1970-08-20 Daimler Benz Ag Gas turbine plant
US3625003A (en) 1970-09-08 1971-12-07 Gen Motors Corp Split compressor gas turbine
US3779486A (en) * 1970-09-26 1973-12-18 Great Britain & Northern Irela Gas turbine engines
US3981616A (en) * 1974-10-22 1976-09-21 The United States Of America As Represented By The Secretary Of The Air Force Hollow composite compressor blade
US4171614A (en) 1976-04-17 1979-10-23 Motoren- Und Turbinen-Union Munchen Gmbh Gas turbine engine
US5131223A (en) 1989-07-13 1992-07-21 Sundstrand Corporation Integrated booster and sustainer engine for a missile
US5619854A (en) * 1994-05-25 1997-04-15 Gec Alsthom Diesels, Ltd. Turbocharged internal combustion engine
US7424805B2 (en) 2005-04-29 2008-09-16 General Electric Company Supersonic missile turbojet engine
US7628018B2 (en) * 2008-03-12 2009-12-08 Mowill R Jan Single stage dual-entry centriafugal compressor, radial turbine gas generator
US8745989B2 (en) 2009-04-09 2014-06-10 Pratt & Whitney Canada Corp. Reverse flow ceramic matrix composite combustor
US20100319343A1 (en) * 2009-06-23 2010-12-23 Arnold Steven D Turbocharger with two-stage compressor, including a twin-wheel parallel-flow first stage
US8176725B2 (en) * 2009-09-09 2012-05-15 United Technologies Corporation Reversed-flow core for a turbofan with a fan drive gear system
US9297311B2 (en) * 2011-03-22 2016-03-29 Alstom Technology Ltd Gas turbine power plant with flue gas recirculation and oxygen-depleted cooling gas
EP2551485A2 (en) 2011-07-26 2013-01-30 United Technologies Corporation Gas turbine engine with aft core driven fan section
US20130025286A1 (en) * 2011-07-26 2013-01-31 Kupratis Daniel B Gas turbine engine with aft core driven fan section
US9460110B2 (en) 2011-09-21 2016-10-04 Kevin Mark Klughart File system extension system and method
US8935912B2 (en) * 2011-12-09 2015-01-20 United Technologies Corporation Gas turbine engine with variable overall pressure ratio
US20130219860A1 (en) * 2012-02-29 2013-08-29 Gabriel L. Suciu Counter-rotating low pressure turbine without turbine exhaust case
US20130255224A1 (en) * 2012-03-27 2013-10-03 Daniel Bernard Kupratis Reverse core gear turbofan
US9140212B2 (en) * 2012-06-25 2015-09-22 United Technologies Corporation Gas turbine engine with reverse-flow core having a bypass flow splitter
US20160169102A1 (en) * 2014-12-12 2016-06-16 United Technologies Corporation Reverse core flow gas turbine engine
US20160237894A1 (en) * 2015-02-13 2016-08-18 United Technologies Corporation Turbine engine with a turbo-compressor

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
European Search Report for Application No. 17191922.8 dated Feb. 13, 2018.

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11002146B1 (en) 2020-10-26 2021-05-11 Antheon Research, Inc. Power generation system
US11448083B2 (en) 2020-10-26 2022-09-20 Antheon Research, Inc. Power generation system
US11530617B2 (en) 2020-10-26 2022-12-20 Antheon Research, Inc. Gas turbine propulsion system
US20230121431A1 (en) * 2020-10-26 2023-04-20 Francis O'Neill Gas turbine propulsion system
US11821323B2 (en) 2020-10-26 2023-11-21 Antheon Research, Inc. Power generation system
US11970947B2 (en) 2020-10-26 2024-04-30 Antheon Research, Inc. Power generation system

Also Published As

Publication number Publication date
EP3296542B1 (en) 2019-05-22
EP3296542A1 (en) 2018-03-21
US20180080373A1 (en) 2018-03-22

Similar Documents

Publication Publication Date Title
US9163521B2 (en) Gas turbine engine with supersonic compressor
US9752510B2 (en) Gas turbine engine driving multiple fans
EP3354876B1 (en) Gas turbine engine architecture with split compressor system
EP3318743A1 (en) Intercooled cooled cooling integrated air cycle machine
US10378438B2 (en) Reverse flow gas turbine engine with radially outward turbine
US20090191045A1 (en) Low pressure turbine with counter-rotating drives for single spool
US20060090451A1 (en) Counter-rotating gas turbine engine and method of assembling same
AU2013273476B2 (en) Combination of two gas turbines to drive a load
US9657643B2 (en) Energy efficient pump system
US11286885B2 (en) External core gas turbine engine assembly
US10094281B2 (en) Gas turbine engine with twin offset gas generators
US11519337B2 (en) Gas turbine auxiliary power unit
US10287991B2 (en) Gas turbine engine with paired distributed fan sets
US10385774B2 (en) Split compressor turbine engine
US20170184020A1 (en) Method and system for in-line distributed propulsion
US9897001B2 (en) Compressor areas for high overall pressure ratio gas turbine engine
US10480519B2 (en) Hybrid compressor
US9915199B2 (en) Bi-directional compression fan rotor for a gas turbine engine
EP2505791A1 (en) Securing system and corresponding gas turbine engine
US20180163664A1 (en) Concentric shafts driving adjacent fans for aircraft propulsion
RU2379523C2 (en) Aircraft propfan engine
US20160160761A1 (en) Gas Turbine Engine With Single Turbine Driving Two Compressors
US10174680B2 (en) Gas turbine engine with distributed fans and bypass air mixer

Legal Events

Date Code Title Description
AS Assignment

Owner name: UNITED TECHNOLOGIES CORPORATION, CONNECTICUT

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:KUPRATIS, DANIEL BERNARD;KASHYAP, TANIA BHATIA;ZELESKY, MARK F.;AND OTHERS;SIGNING DATES FROM 20160913 TO 20160916;REEL/FRAME:039776/0138

STPP Information on status: patent application and granting procedure in general

Free format text: NON FINAL ACTION MAILED

STPP Information on status: patent application and granting procedure in general

Free format text: RESPONSE TO NON-FINAL OFFICE ACTION ENTERED AND FORWARDED TO EXAMINER

STPP Information on status: patent application and granting procedure in general

Free format text: NOTICE OF ALLOWANCE MAILED -- APPLICATION RECEIVED IN OFFICE OF PUBLICATIONS

STPP Information on status: patent application and granting procedure in general

Free format text: PUBLICATIONS -- ISSUE FEE PAYMENT VERIFIED

STCF Information on status: patent grant

Free format text: PATENTED CASE

AS Assignment

Owner name: RAYTHEON TECHNOLOGIES CORPORATION, MASSACHUSETTS

Free format text: CHANGE OF NAME;ASSIGNOR:UNITED TECHNOLOGIES CORPORATION;REEL/FRAME:054062/0001

Effective date: 20200403

AS Assignment

Owner name: RAYTHEON TECHNOLOGIES CORPORATION, CONNECTICUT

Free format text: CORRECTIVE ASSIGNMENT TO CORRECT THE AND REMOVE PATENT APPLICATION NUMBER 11886281 AND ADD PATENT APPLICATION NUMBER 14846874. TO CORRECT THE RECEIVING PARTY ADDRESS PREVIOUSLY RECORDED AT REEL: 054062 FRAME: 0001. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF ADDRESS;ASSIGNOR:UNITED TECHNOLOGIES CORPORATION;REEL/FRAME:055659/0001

Effective date: 20200403

MAFP Maintenance fee payment

Free format text: PAYMENT OF MAINTENANCE FEE, 4TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1551); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY

Year of fee payment: 4

AS Assignment

Owner name: RTX CORPORATION, CONNECTICUT

Free format text: CHANGE OF NAME;ASSIGNOR:RAYTHEON TECHNOLOGIES CORPORATION;REEL/FRAME:064714/0001

Effective date: 20230714