US20100000217A1 - Turbine engine with interstage heat - Google Patents
Turbine engine with interstage heat Download PDFInfo
- Publication number
- US20100000217A1 US20100000217A1 US12/557,616 US55761609A US2010000217A1 US 20100000217 A1 US20100000217 A1 US 20100000217A1 US 55761609 A US55761609 A US 55761609A US 2010000217 A1 US2010000217 A1 US 2010000217A1
- Authority
- US
- United States
- Prior art keywords
- turbine
- working fluid
- section
- engine
- heat
- 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.)
- Abandoned
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
- 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
- F02C7/08—Heating air supply before combustion, e.g. by exhaust gases
-
- 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
- F02C1/00—Gas-turbine plants characterised by the use of hot gases or unheated pressurised gases, as the working fluid
- F02C1/04—Gas-turbine plants characterised by the use of hot gases or unheated pressurised gases, as the working fluid the working fluid being heated indirectly
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02E—REDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
- Y02E20/00—Combustion technologies with mitigation potential
- Y02E20/14—Combined heat and power generation [CHP]
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02E—REDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
- Y02E20/00—Combustion technologies with mitigation potential
- Y02E20/16—Combined cycle power plant [CCPP], or combined cycle gas turbine [CCGT]
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Engine Equipment That Uses Special Cycles (AREA)
Abstract
Efficiency and/or power are increased in a turbine engine by using a self-contained, passive heat transfer device, such as a heat pipe, to transfer heat from working fluid in one section of the engine to working fluid in another section of the engine.
Description
- This application is a Divisional of copending U.S. patent application Ser. No. 11/336,330, filed Jan. 20, 2006, and which claims the benefit of U.S. Provisional Application No. 60/701,010, filed Jul. 20, 2005.
- This invention relates generally to turbine engines, such as gas turbines and steam turbines, and more particularly to increasing the power density and/or efficiency of turbine engines.
- Turbine engines have long been used for a variety of purposes, including power generation and aircraft and marine propulsion. Briefly, a gas turbine engine includes a compressor, which provides pressurized air to a combustor, wherein it is mixed with fuel and ignited for generating hot combustion gases. These gases are expanded in a turbine that extracts energy therefrom for powering the compressor and providing useful work. A steam turbine includes a turbine that is driven by the expansion of superheated steam, which is produced by a boiler or the like.
- Various efforts have been used to enhance the efficiency and power output of these engines. For example, turbine reheat involves heating the gas or steam as it expands through the turbine to increase the turbine work output. The current practice for implementing reheat typically involves the injection of fuel, steam or any reheated working fluid at different stages. However, this practice requires use of generally massive and expensive auxiliary equipment such as valves, pumps, piping, etc.
- Another technique for increasing efficiency in gas turbines is compressor intercooling, which involves cooling air between stages in the compressor at a constant pressure. This reduces the work needed to achieve compression because a cooler gas is more easily compressed. Like turbine reheat, compressor intercooling typically requires the addition of massive and expensive auxiliary equipment.
- Accordingly, there is a need for a more compact and less expensive approach to increasing turbine engine power density and/or efficiency.
- The above-mentioned need is met by the present invention, which provides a turbine engine that includes a self-contained, passive heat transfer device, such as a heat pipe, arranged to transfer heat from working fluid in one section of the engine to working fluid in another section of the engine. In one embodiment, heat is transferred from the turbine to the compressor discharge air.
- In another embodiment, heat is transferred from a heat source to the turbine. In a further embodiment, heat is transferred from a forward portion of the turbine to an aft portion of the turbine. In yet another embodiment, heat is transferred from the compressor to a heat sink. Another possible embodiment applies to a steam turbine in which heat is transferred from the boiling device to the turbine.
- The present invention and its advantages over the prior art will be more readily understood upon reading the following detailed description and the appended claims with reference to the accompanying drawings.
- The subject matter that is regarded as the invention is particularly pointed out and distinctly claimed in the concluding part of the specification. The invention, however, may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which:
-
FIG. 1 is a sectional view of a gas turbine engine in accordance with a first embodiment of the invention. -
FIG. 2 is a fragmentary, sectional view taken along line 2-2 ofFIG. 1 . -
FIG. 3 is a sectional view of a gas turbine engine in accordance with a second embodiment of the invention. -
FIG. 4 is a sectional view of a gas turbine engine in accordance with a third embodiment of the invention. -
FIG. 5 is a sectional view of a gas turbine engine in accordance with a fourth embodiment of the invention. -
FIG. 6 is a sectional view of a steam turbine engine in accordance with a fifth embodiment of the invention. - The present invention generally relates to increasing efficiency and/or power output of turbine engines. As used herein, the term “turbine engine” includes gas turbines, steam turbines or any other device that utilizes a turbine. The term also includes systems that use one or more turbine engines, such as cogeneration or combined cycle power plants. Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,
FIG. 1 illustrates a longitudinal cross-sectional view of agas turbine engine 10. Theengine 10 includes, in serial axial flow communication about alongitudinal centerline axis 12, acompressor 14, aheat source 16, and aturbine 18 enclosed by acasing 20. Theturbine 18 is drivingly connected to thecompressor 14 via arotor shaft 22. - The
compressor 14 is a multi-stage, axial compressor configured for pressurizing air as it flows downstream therethrough. Each stage includes a plurality of circumferentially spaced apart rotor blades 24 (only one per stage shown inFIG. 1 ) extending radially outwardly from therotor shaft 22. A number of rows of circumferentially spaced apart stationary blades orstator vanes 26, which are fixed to thecasing 20, are interleaved with the rows ofcompressor rotor blades 24. Thus, for each stage, therotor blades 24 are located immediately downstream or aft of thestator vanes 26. In the illustrated embodiment, thecompressor 14 has six stages. The first-stage stator vanes 26 function as inlet guide vanes that provide flow conditioning for the first-stagecompressor rotor blades 24. - The
compressor 14 provides compressed air to theheat source 16 via adiffuser 27, which is located immediately downstream of thecompressor 14. The diffuser 27 conditions the compressor discharge air to be suitable for theheat source 16. As shown inFIG. 1 , theheat source 16 is a combustor having a generally annular hollow body defining acombustion chamber 28 therein. (A possible alternative to the annular combustor shown inFIG. 1 is the so-called cannular combustor, which includes a plurality of combustor cans arranged annularly about the engine, wherein each can is connected to the compressor and exhausts combustion products to the turbine.) Fuel is introduced into the forward end of thecombustor 16 by one or more fuel injectors 30 (only one shown inFIG. 1 ) and is mixed with the compressed air in any suitable fashion. The resulting fuel-air mixture is ignited and burned in thecombustion chamber 28 for generating hot combustion gases. Most of the compressed air discharged from thecompressor 14 passes into thecombustor 16 to support combustion. However, a portion of the compressed air can be bled off and used to cool combustor liners and turbomachinery further downstream. As another alternative, theheat source 16 could be a heat exchanger that transfers heat from an external source, such as a waste heat, to the compressor discharge air. - The hot combustion gases are discharged to the
turbine 18 located downstream of theheat source 16, where they are expanded so that energy is extracted. Theturbine 18 is a multi-stage, axial turbine wherein each stage includes a plurality of circumferentially spaced apart rotor blades 32 (only one per stage shown inFIG. 1 ) extending radially outwardly from therotor shaft 22. A number of rows of circumferentially spaced apart stationary blades ornozzles 34, which are fixed to thecasing 20, are interleaved with the rows ofrotor blades 32. Thus, for each stage, theturbine rotor blades 32 are located immediately downstream or aft of theturbine nozzles 34. Theturbine 18, as shown by way of example inFIG. 1 , has five stages. The expanding hot combustion gases cause therotor shaft 22 to rotate, which drives thecompressor 14 and produces usable work. - The
gas turbine engine 10 further includes one or moreheat transfer devices 36 positioned to transfer heat from one section of theengine 10 to a different and cooler section the ofengine 10. Although only one such heat transfer device is shown inFIG. 1 , it should be noted that more than one heat transfer device could be employed, with the plurality of devices preferably being equally spaced around the circumference of theengine 10. - One type of heat transfer device that can be employed is a heat pipe. A heat pipe generally comprises a sealed container containing a fluid at a pressure that allows the fluid to operate close to its liquid-gas phase change temperature. While various fluids including water can be used in the heat pipe, preferred fluids include liquid metals such as cesium, rubidium, potassium, sodium and mercury. One end of the container defines a cold side or condensing section and the other end defines a hot side or evaporative section. The heat pipe includes an internal capillary device, such as a wick, to draw condensed fluid from the cold side to the hot side. In the hot side of the heat pipe, the fluid absorbs heat from the surrounding environment until enough heat is absorbed to change the phase of the fluid from liquid to gas. Then the pressure differential caused by the temperature difference causes the gas to travel to the cold side where heat is removed and the fluid returns to the liquid phase, where it is again drawn to the hot side by the capillary device. The heat pipe thus uses the phase change of the fluid to remove heat from the hot evaporative section and deposit heat in the cooler location. The heat pipe is a self-contained device in that it uses its internal operating fluid, and not working fluid from the turbine engine, to transfer heat. The heat pipe is also a passive device, as its operation relies on the temperature difference and does not require any additional work input, such as an external pump.
- In the illustrated embodiment, the
heat transfer device 36 is a heat pipe having a hot side orevaporative section 38 positioned to be in thermal communication with the hot combustion gases in theturbine 18 and a cold side or condensingsection 40 positioned to be in thermal communication with the compressed air discharged from thecompressor 14. For instance, theevaporative section 38 is in thermal communication with one of the forward-most or first-stage turbine nozzles 34, and the condensingsection 40 is in thermal communication with thediffuser 27. While theevaporative section 38 is shown as being in thermal communication with a first-stage turbine nozzle, it should be noted that theevaporative section 38 could alternatively be in thermal communication with a latter-stage turbine nozzle. However, because the first-stage turbine nozzles are typically the hottest, and the closest to thediffuser 27, it is generally more effective to use a first-stage turbine nozzle. Also, theevaporative section 38 could be in thermal contact with an exterior surface of the turbine nozzle, or theevaporative section 38 could be incorporated into the structure of the turbine nozzle, essentially making the turbine nozzle the evaporative section. Furthermore, theevaporative section 38 need not necessarily be in physical contact with the nozzle structure. Theevaporative section 38 could be positioned anywhere, such as between adjacent turbine nozzles, so that it is in thermal communication with the hot gases discharged from theheat source 16. Similarly, the condensingsection 40 can be positioned within the flow in thediffuser 27, positioned in physical contact with a wall or walls of thediffuser 27, or incorporated into a wall or walls of thediffuser 27. - For example,
FIG. 2 shows one configuration in which theevaporative section 38 is incorporated into the structure of aturbine nozzle 34. The portion of theheat pipe container 41 that makes up theevaporative section 38 extends into the interior of theturbine nozzle 34 and is encased by the nozzle walls. Acapillary device 43, such as a wick, is disposed in the center of the heat pipe container to draw condensed fluid from the condensingsection 40 to theevaporative section 38. Agap 45 between the internal surface of thecontainer 41 and thecapillary device 43 defines a volume through which vapor can travel back to the condensingsection 40. Theheat pipe container 41 is preferably made of a material having high thermal conductivity so that heat will be efficiently transferred from the hot combustion gases, to the turbine nozzle walls, to the heat pipe container and to the heat pipe fluid. - As shown in
FIG. 1 , theheat pipe 36 penetrates thecasing 20 near its two ends so that the intermediate section between theevaporative section 38 and the condensingsection 40 is located outside of thecasing 20. Alternatively, the intermediate section could be attached to, or embedded in, the casing wall. - As mentioned above, more than one heat transfer device can be employed. For instance, there could be enough heat pipes so that each one of the first-stage turbine nozzles was in thermal communication with a respective evaporative section. Alternatively, the number of heat pipes could be such that only some portion of the turbine nozzles in the selected stage would be in thermal communication with a respective evaporative section, such as every other turbine nozzle or every third turbine nozzle.
- The
heat pipe 36 transfers heat from the hot combustion gases in theturbine 18 to the cooler compressed air being discharged from thecompressor 14. Therefore, the enthalpy of the working fluid (the hot combustion gases) in theturbine 18 is decreased, and the enthalpy of the working fluid (the compressed air) in the cooler area forward of thecombustion chamber 28 is increased. Engine efficiency is thus increased because the increased compressed air temperature results in higher firing temperatures. That is, because the compressed air discharged from thecompressor 14 is “preheated,” the temperature of the combustion products will be greater for a given amount of fuel. Another way to look at it is that less fuel is required to raise the temperature of the hot combustion products entering theturbine 18 to the desired level. Furthermore, removing heat from theturbine nozzles 34 reduces the cooling load needed for the turbine structure, thereby further increasing engine efficiency and/or power density. - Referring to
FIG. 3 , agas turbine engine 110 in accordance with a second embodiment of the present invention is shown. Like the engine of the first described embodiment, theengine 110 includes, in serial axial flow communication about alongitudinal centerline axis 112, acompressor 114, aheat source 116, and aturbine 118 enclosed by acasing 120. Theturbine 118 is drivingly connected to thecompressor 114 via arotor shaft 122. Thecompressor 114 is a multi-stage axial compressor having interleaved rows ofrotor blades 124 and stationary blades or stator vanes 126 (only one of each shown per stage). Theturbine 118 is a multi-stage turbine having interleaved rows ofrotor blades 132 and stationary blades or nozzles 134 (only one of each shown per stage). - In operation, the
compressor 114 provides compressed air to theheat source 116, typically via a diffuser (not shown inFIG. 3 ). In the illustrated embodiment, theheat source 116 is a combustor having a generally annular hollow body defining acombustion chamber 128 therein. Fuel is introduced into the forward end of thecombustor 116 by one or more fuel injectors 130 (only one shown inFIG. 3 ) and is mixed with the compressed air in any suitable fashion. The resulting fuel-air mixture is ignited and burned in thecombustion chamber 128 for generating hot combustion gases. The hot combustion gases are discharged to theturbine 118 located downstream of thecombustor 116, where they are expanded to drive theturbine 118 so that energy can be extracted. - Similarly to the previously described embodiment, the
gas turbine engine 110 includes one or moreheat transfer devices 136 positioned to transfer heat from one section of theengine 110 to a different and cooler section the ofengine 110. Although only one such heat transfer device is shown inFIG. 3 , it should be noted that more than one heat transfer device could be employed, with the plurality of devices preferably being equally spaced around the circumference of theengine 110. - In this second embodiment, the
heat transfer device 136 is a heat pipe having a hot side orevaporative section 138 positioned to be in thermal communication with the hot combustion gases in theheat source 116 and a cold side or condensingsection 140 positioned to be in thermal communication with the combustion gases in theturbine 118. Generally, the condensingsection 140 is located in the downstream portion of theturbine 118 where the expanding combustion gases are appreciably cooler than the gases in theheat source 116. Specifically, the condensingsection 140 is shown to be in thermal communication with the one of the fourth-stage turbine nozzles 134, although it should be noted that thecondensing section 140 could alternatively be in thermal communication with a turbine nozzle in one of the other turbine stages. However, because the temperature difference between the earlier stages of theturbine 118 and theheat source 116 is typically not large, it is generally preferred to use one of the latter-stage turbine nozzles, which has a more significant temperature differential with theheat source 116. Also, the condensingsection 140 could be in thermal contact with an exterior surface of the turbine nozzle, or thecondensing section 140 could be incorporated into the structure of the turbine nozzle, essentially making the turbine nozzle the condensing section. - The
evaporative section 138 of theheat pipe 136 can be suspended in thecombustion chamber 128 so as to be positioned away from the combustor walls. Alternatively,evaporative section 138 could be mounted to (as shown inFIG. 3 ), or even incorporated into, the combustor wall. - As mentioned above, more than one heat transfer device can be employed. For instance, there could be enough heat pipes so that each turbine nozzle of the selected stage is in thermal communication with a respective condensing section. Alternatively, the number of heat pipes could be such that only some portion of the turbine nozzles in the selected stage would be in thermal communication with a respective condensing section, such as every other turbine nozzle or every third turbine nozzle.
- The
heat pipe 136 transfers heat from the hot combustion gases in theheat source 116 to the cooler combustion gases in the downstream portion of theturbine 118, thereby “reheating” the combustion gases. Therefore, the enthalpy of the working fluid (the hot combustion gases) in theheat source 116 is decreased, and the enthalpy of the working fluid (the cooler combustion gases) being expanded in the downstream portion of theturbine 118 is increased. This use of reheat increases the work output of the engine 110 (compared to a simple cycle not using reheat) without increasing fuel input or work input to the compressor, thereby making theengine 110 more efficient and/or power dense. - Turning to
FIG. 4 , agas turbine engine 210 in accordance with a third embodiment of the present invention is shown. Like the engines of the prior embodiments, theengine 210 includes, in serial axial flow communication about alongitudinal centerline axis 212, acompressor 214, aheat source 216, and aturbine 218 enclosed by acasing 220. Theturbine 218 is drivingly connected to thecompressor 214 via arotor shaft 222. Thecompressor 214 is a multi-stage axial compressor having interleaved rows ofrotor blades 224 and stationary blades or stator vanes 226 (only one of each shown per stage). Theturbine 218 is a multi-stage turbine having interleaved rows ofrotor blades 232 and stationary blades or nozzles 234 (only one of each shown per stage). - In operation, the
compressor 214 provides compressed air to theheat source 216, typically via a diffuser (not shown inFIG. 4 ). In the illustrated embodiment, theheat source 216 is a combustor having a generally annular hollow body defining acombustion chamber 228 therein. Fuel is introduced into the forward end of thecombustor 216 by one or more fuel injectors 230 (only one shown inFIG. 4 ) and is mixed with the compressed air in any suitable fashion. The resulting fuel-air mixture is ignited and burned in thecombustion chamber 228 for generating hot combustion gases. The hot combustion gases are discharged to theturbine 218 located downstream of thecombustor 216, where they are expanded to drive theturbine 218 so that energy can be extracted. - Similarly to the previously described embodiments, the
gas turbine engine 210 includes one or moreheat transfer devices 236 positioned to transfer heat from one section of theengine 210 to a different and cooler section the ofengine 210. Although only one such heat transfer device is shown inFIG. 4 , it should be noted that more than one heat transfer device could be employed, with the plurality of devices preferably being equally spaced around the circumference of theengine 210. - In this embodiment, the
heat transfer device 236 is a heat pipe having a hot side orevaporative section 238 positioned to be in thermal communication with the hot combustion gases in the forward or upstream portion of portion of theturbine 218 and a cold side or condensingsection 240 positioned to be in thermal communication with the combustion gases in the aft or downstream portion of theturbine 218. The expanding combustion gases are appreciably cooler in the aft portion of theturbine 218. Theevaporative section 238 is in thermal communication with one of theturbine nozzles 234 of a forward stage, such as the second stage, as shown inFIG. 4 , and thecondensing section 240 is in thermal communication with the one of theturbine nozzles 234 of an aft stage, such as the aft-most or fifth stage as shown inFIG. 4 . The greatest temperature differential is realized by using the forward-most and aft-most stages, but it should be noted that turbine nozzles in other stages could also be used. Also, theevaporative section 238 and thecondensing section 240 could be in thermal contact with an exterior surface of the respective turbine nozzles, or alternatively could be incorporated into the structure of the respective turbine nozzles. - As mentioned above, more than one heat transfer device can be employed. For instance, there could be enough heat pipes so that each turbine nozzle of the selected stages is in thermal communication with a respective evaporative or condensing section, as the case may be. Alternatively, the number of heat pipes could be such that only some portion of the turbine nozzles in the selected stages would be in thermal communication with a respective evaporative or condensing section, such as every other turbine nozzle or every third turbine nozzle.
- The
heat pipe 236 transfers heat from the hot combustion gases in the forward portion of theturbine 218 to the cooler combustion gases in the aft portion of theturbine 218, thereby “reheating” the downstream combustion gases. Therefore, the enthalpy of the working fluid (the hot combustion gases) in the forward portion of theturbine 218 is decreased, and the enthalpy of the working fluid (the cooler combustion gases) being expanded in the aft portion of theturbine 218 is increased. This use of reheat increases the work output of the engine 210 (compared to a simple cycle not using reheat) without increasing fuel input or work input to the compressor, thereby making theengine 210 more efficient. - Referring to
FIG. 5 , agas turbine engine 310 in accordance with a fourth embodiment of the present invention is shown. Like the engines of the prior embodiments, theengine 310 includes, in serial axial flow communication about alongitudinal centerline axis 312, acompressor 314, aheat source 316, and aturbine 318 enclosed by acasing 320. Theturbine 318 is drivingly connected to thecompressor 314 via arotor shaft 322. Thecompressor 314 is a multi-stage axial compressor having interleaved rows ofrotor blades 324 and stationary blades or stator vanes 326 (only one of each shown per stage). Theturbine 318 is a multi-stage turbine having interleaved rows ofrotor blades 332 and stationary blades or nozzles 334 (only one of each shown per stage). - In operation, the
compressor 314 provides compressed air to theheat source 316, typically via a diffuser (not shown inFIG. 5 ). In the illustrated embodiment, theheat source 316 is a combustor having a generally annular hollow body defining acombustion chamber 328 therein. Fuel is introduced into the forward end of thecombustor 316 by one or more fuel injectors 330 (only one shown inFIG. 5 ) and is mixed with the compressed air in any suitable fashion. The resulting fuel-air mixture is ignited and burned in thecombustion chamber 328 for generating hot combustion gases. The hot combustion gases are discharged to theturbine 318 located downstream of thecombustor 316, where they are expanded to drive theturbine 318 so that energy can be extracted. - Similarly to the previously described embodiments, the
gas turbine engine 310 includes one or moreheat transfer devices 336 positioned to transfer heat from one section of theengine 310 to a different and cooler section the ofengine 310. Although only one such heat transfer device is shown inFIG. 5 , it should be noted that more than one heat transfer device could be employed, with the plurality of devices preferably being equally spaced around the circumference of theengine 310. - In this embodiment, the
heat transfer device 336 is a heat pipe having a hot side orevaporative section 338 positioned to be in thermal communication with the air flow in thecompressor 314 and a cold side or condensingsection 340 positioned to be in thermal communication with a cooler working fluid that functions as a heat sink. Specifically, theevaporative section 338 is in thermal communication with the one of the compressor stator vanes 326. While theevaporative section 338 can be in thermal communication with any of thecompressor stator vane 326, it is preferably in thermal communication with a mid-stage compressor stator vane, such as a fourth-stagecompressor stator vane 326 as shown inFIG. 5 . Because the compressor air is approximately halfway to the full pressurization in the mid-stages, placing theevaporative section 338 in thermal communication with a mid-stage compressor stator vane is generally more effective in providing the intercooling effect described in more detail below. Furthermore, theevaporative section 338 could be in thermal contact with an exterior surface of thecompressor stator vane 326, or theevaporative section 338 could be incorporated into the structure of the compressor stator vane, essentially making the compressor stator vane the evaporative section. - In the illustrated embodiment, the condensing
section 340 is in thermal communication with a bypass airflow (represented by arrow A) located outside of thecasing 320. For example, one type of gas turbine engine commonly used for aircraft propulsion is the bypass turbofan engine. A bypass turbofan engine includes a fan (not shown) located forward of thecompressor 314 and driven by the turbine via a dual shaft arrangement. The air exiting the fan is split so that a portion of the air flows into thecompressor 314 and the rest of the air (i.e., the bypass airflow A) bypasses the engine core outside of thecasing 320. The bypass airflow A flows through a duct defined between thecasing 320 and a nacelle (not shown) and provides most of the engine thrust. WhileFIG. 5 shows using the bypass airflow as a heat sink, it should be noted that other fluid flows in or about theengine 310 could be used as a heat sink. - The condensing
section 340 is located in the bypass airflow A so that heat from the condensingsection 340 is transferred to the relatively cool bypass airflow. The condensingsection 340 can be configured withexternal fins 342 to facilitate the heat transfer. - As mentioned above, more than one heat transfer device can be employed. For instance, there could be enough heat pipes so that each compressor stator vane of the selected stage is in thermal communication with a respective evaporative section. Alternatively, the number of heat pipes could be such that only some portion of the compressor stator vanes in the selected stage would be in thermal communication with a respective evaporative section, such as every other compressor stator vane or every third compressor stator vane.
- The
heat pipe 336 transfers heat from the air flow in thecompressor 314 to the bypass airflow, thereby providing “intercooling” to thecompressor 314. Therefore, the enthalpy of the working fluid (the air flow) in thecompressor 314 is decreased, and the enthalpy of the working fluid (the bypass flow) of the heat sink is increased. This use of intercooling decreases the amount of work input to the compressor 314 (compared to a simple cycle not using intercooling) needed to achieve a given pressure ratio. This results in a net increase in the work output of theengine 310, which increases the overall efficiency of theengine 310. - Referring to
FIG. 6 , asteam turbine engine 410 in accordance with a fifth embodiment of the present invention is shown. The steam engine includes amulti-stage turbine 418 and a boilingdevice 419 comprising aboiler 444 and aheat source 446. Theturbine 418 includes interleaved rows of circumferentially spacedrotor blades 432 and circumferentially spaced stationary blades ornozzles 434 enclosed by acasing 420. Therotor blades 432 extend radially outwardly from arotor shaft 422, and thenozzles 434 are fixed to the casing 420 (only one of each shown per stage). Therotor shaft 422 is mounted to rotate about alongitudinal centerline axis 412. - The
heat source 446 heatswater 448 in theboiler 444 to produce high-pressure, superheated steam that is supplied to the forward end of theturbine 418 via aconduit 450. The superheated steam is expanded in theturbine 418 to drive theturbine 418 and therotor shaft 422 so that work can be produced. Theheat source 446 could be a burner, hot exhaust products from a gas turbine engine, or other waste heat sources, thereby enabling this embodiment to be used in conjunction with cogeneration or combined cycle plants. - The
steam turbine engine 410 includes one or moreheat transfer devices 436 positioned to transfer heat from one section of theengine 410 to a different and cooler section the ofengine 410. Although only one such heat transfer device is shown inFIG. 6 , it should be noted that more than one heat transfer device could be employed. - In the illustrated embodiment, the
heat transfer device 436 is a heat pipe having a hot side orevaporative section 438 positioned to be in thermal communication with the hot gas from theheat source 446 that is heating theboiler 444 and a cold side or condensingsection 440 positioned to be in thermal communication with the expanding vapor in theturbine 418. As an alternative, theevaporative section 438 positioned to be in thermal communication with the superheated steam in theboiler 444 or theconduit 450. Generally, the condensingsection 440 is located in the downstream or aft portion of theturbine 418 where the vapor is appreciably cooler than the heat source gas. Specifically, the condensingsection 440 is shown to be in thermal communication with the one of the fourth-stage turbine nozzles 434, although it should be noted that thecondensing section 440 could alternatively be in thermal communication with a turbine nozzle in one of the other turbine stages. However, it is generally preferred to use one of the latter-stage turbine nozzles because the temperature difference between the latter stages of theturbine 418 and theheat source 446 will be greater. Also, the condensingsection 440 could be in thermal contact with an exterior surface of the turbine nozzle, or thecondensing section 440 could be incorporated into the structure of the turbine nozzle, essentially making the turbine nozzle the condensing section. - As mentioned above, more than one heat transfer device can be employed. For instance, there could be enough heat pipes so that each turbine nozzle of the selected stage is in thermal communication with a respective condensing section. Alternatively, the number of heat pipes could be such that only some portion of the turbine nozzles in the selected stage would be in thermal communication with a respective condensing section, such as every other turbine nozzle or every third turbine nozzle.
- The
heat pipe 436 transfers heat from the hot gas of theheat source 446 to the expanding vapor in the aft portion of theturbine 418, thereby “reheating” the vapor. Therefore, the enthalpy of the working fluid (the hot gas) from theheat source 446 is decreased, and the enthalpy of the working fluid (the vapor) in the aft portion of theturbine 418 is increased. This use of reheat increases the work output of the engine 410 (compared to a simple cycle not using reheat) without increasing fuel input or work input to the compressor, thereby making theengine 410 more efficient. - Although all of the illustrated embodiments describe a heat pipe for the heat transfer device, it should be noted that other types of heat transfer devices could be employed. For example, the heat transfer device may also comprise a material that has a very high thermal conductivity, such as solid crystal diamond, formed into a high thermal conductivity conduit. Such a high thermal conductivity conduit would be a self-contained, passive heat transfer device because it does not utilize the turbine engine's working fluid or an additional work input to transfer heat. Another alternative is to use a self-contained device that utilizes internal nanopumps for an internal heat transfer circuit.
- While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention as defined in the appended claims.
Claims (20)
1. A turbine engine comprising:
a turbine;
means for delivering a hot, pressurized working fluid to said turbine; and
a self-contained, passive heat transfer device arranged to transfer heat from working fluid in one section of said engine to working fluid in another section of said engine.
2. The turbine engine of claim 1 wherein said heat transfer device is a heat pipe having a condensing section and an evaporative section.
3. The turbine engine of claim 2 wherein said means for delivering a hot, pressurized working fluid to said turbine includes a compressor, and wherein said condensing section is in thermal communication with working fluid discharged from said compressor and said evaporative section is in thermal communication with working fluid downstream of said means for delivering a hot, pressurized working fluid to said turbine.
4. The turbine engine of claim 3 further comprising a diffuser located downstream of said compressor, said condensing section being in thermal communication with said diffuser.
5. The turbine engine of claim 4 wherein said condensing section is incorporated into said diffuser.
6. The turbine engine of claim 2 wherein said means for delivering a hot, pressurized working fluid to said turbine includes a heat source, and wherein said condensing section is in thermal communication with working fluid in said turbine and said evaporative section is in thermal communication with working fluid in said heat source.
7. The turbine engine of claim 2 wherein said condensing section is in thermal communication with working fluid in an aft portion of said turbine and said evaporative section is in thermal communication with working fluid in a forward portion of said turbine.
8. The turbine engine of claim 2 wherein said means for delivering a hot, pressurized working fluid to said turbine includes a compressor, and wherein said condensing section is in thermal communication with working fluid in a heat sink and said evaporative section is in thermal communication with working fluid in said compressor.
9. The turbine engine of claim 8 wherein said heat sink is a bypass flow.
10. The turbine engine of claim 2 wherein said means for delivering a hot, pressurized working fluid to said turbine includes a boiling device, and wherein said condensing section is in thermal communication with working fluid in said turbine and said evaporative section is in thermal communication with working fluid in said boiling device.
11. The turbine engine of claim 2 wherein said turbine is an axial turbine having multiple stages of turbine nozzles.
12. The turbine engine of claim 11 wherein said condensing section is in thermal communication with an aft stage turbine nozzle.
13. The turbine engine of claim 12 wherein said means for delivering a hot, pressurized working fluid to said turbine includes a combustor, and wherein said evaporative section is in thermal communication with working fluid in said combustor.
14. The turbine engine of claim 12 wherein said evaporative section is in thermal communication with a forward stage turbine nozzle.
15. The turbine engine of claim 12 wherein said means for delivering a hot, pressurized working fluid to said turbine includes a boiling device, and wherein said evaporative section is in thermal communication with working fluid in said boiling device.
16. The turbine engine of claim 11 wherein said condensing section is incorporated into one of said turbine nozzles.
17. The turbine engine of claim 11 wherein said evaporative section is incorporated into one of said turbine nozzles.
18. The turbine engine of claim 1 further comprising one or more additional a heat transfer devices arranged to transfer heat from working fluid in one section of said engine to working fluid in another section of said engine.
19. A method of operating a turbine engine comprising a turbine, said method comprising:
expanding a hot, pressurized working fluid in said turbine; and
using at least one self-contained, passive heat transfer device to transfer heat from working fluid in one section of said engine to working fluid in another section of said engine.
20. The method of claim 19 wherein heat is transferred away from said turbine.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US12/557,616 US20100000217A1 (en) | 2005-07-20 | 2009-09-11 | Turbine engine with interstage heat |
Applications Claiming Priority (3)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US70101005P | 2005-07-20 | 2005-07-20 | |
US11/336,330 US7600382B2 (en) | 2005-07-20 | 2006-01-20 | Turbine engine with interstage heat transfer |
US12/557,616 US20100000217A1 (en) | 2005-07-20 | 2009-09-11 | Turbine engine with interstage heat |
Related Parent Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US11/336,330 Division US7600382B2 (en) | 2005-07-20 | 2006-01-20 | Turbine engine with interstage heat transfer |
Publications (1)
Publication Number | Publication Date |
---|---|
US20100000217A1 true US20100000217A1 (en) | 2010-01-07 |
Family
ID=37677799
Family Applications (2)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US11/336,330 Expired - Fee Related US7600382B2 (en) | 2005-07-20 | 2006-01-20 | Turbine engine with interstage heat transfer |
US12/557,616 Abandoned US20100000217A1 (en) | 2005-07-20 | 2009-09-11 | Turbine engine with interstage heat |
Family Applications Before (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US11/336,330 Expired - Fee Related US7600382B2 (en) | 2005-07-20 | 2006-01-20 | Turbine engine with interstage heat transfer |
Country Status (2)
Country | Link |
---|---|
US (2) | US7600382B2 (en) |
WO (1) | WO2007015847A2 (en) |
Cited By (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO2014099482A1 (en) * | 2012-12-18 | 2014-06-26 | United Technologies Corporation | Oscillating heat pipe for thermal management of gas turbine engines |
US20150008611A1 (en) * | 2012-02-17 | 2015-01-08 | Konica Minolta, Inc. | Method and apparatus for production of an obliquely stretched long film |
US9790893B2 (en) | 2013-03-14 | 2017-10-17 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine flow duct having integrated heat exchanger |
Families Citing this family (20)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US7966807B2 (en) * | 2007-01-17 | 2011-06-28 | United Technologies Corporation | Vapor cooled static turbine hardware |
FR2915520B1 (en) * | 2007-04-30 | 2012-05-25 | Snecma | ENGINE ASSEMBLY COMPRISING ONE OR MORE CALODUCES FOR COOLING A HIGH-PRESSURE COMPRESSOR |
DE102007026455A1 (en) * | 2007-06-05 | 2008-12-11 | Rolls-Royce Deutschland Ltd & Co Kg | Jet engine with compressor air circulation and method of operating the same |
EP2148045A1 (en) * | 2008-07-25 | 2010-01-27 | Siemens Aktiengesellschaft | Casing section for a gas turbine |
US8677763B2 (en) * | 2009-03-10 | 2014-03-25 | General Electric Company | Method and apparatus for gas turbine engine temperature management |
US9033648B2 (en) * | 2010-12-24 | 2015-05-19 | Rolls-Royce North American Technologies, Inc. | Cooled gas turbine engine member |
US10151243B2 (en) | 2013-02-23 | 2018-12-11 | Rolls-Royce Corporation | Cooled cooling air taken directly from combustor dome |
US20160290231A1 (en) * | 2015-04-02 | 2016-10-06 | General Electric Company | Heat pipe intercooling system for a turbomachine |
US20160290232A1 (en) * | 2015-04-02 | 2016-10-06 | General Electric Company | Heat pipe cooling system for a turbomachine |
US20160290235A1 (en) * | 2015-04-02 | 2016-10-06 | General Electric Company | Heat pipe temperature management system for a turbomachine |
US9797310B2 (en) * | 2015-04-02 | 2017-10-24 | General Electric Company | Heat pipe temperature management system for a turbomachine |
JP6585073B2 (en) * | 2015-04-02 | 2019-10-02 | ゼネラル・エレクトリック・カンパニイ | Heat pipe temperature management system for wheels and buckets in turbomachinery |
US10400675B2 (en) * | 2015-12-03 | 2019-09-03 | General Electric Company | Closed loop cooling method and system with heat pipes for a gas turbine engine |
US20170159675A1 (en) * | 2015-12-03 | 2017-06-08 | General Electric Company | Closed loop cooling method for a gas turbine engine |
US10309242B2 (en) * | 2016-08-10 | 2019-06-04 | General Electric Company | Ceramic matrix composite component cooling |
US10450957B2 (en) * | 2017-01-23 | 2019-10-22 | United Technologies Corporation | Gas turbine engine with heat pipe system |
US10392968B2 (en) | 2017-04-24 | 2019-08-27 | United Technologies Corporation | Turbine casing cooling structure |
US10450892B2 (en) | 2017-04-24 | 2019-10-22 | United Technologies Corporation | Thermal management of turbine casing using varying working mediums |
US11022037B2 (en) * | 2018-01-04 | 2021-06-01 | General Electric Company | Gas turbine engine thermal management system |
US11459945B1 (en) | 2021-09-10 | 2022-10-04 | Hamilton Sundstrand Corporation | Micro-turbine generator multi-stage turbine with integrated reheat cycle |
Citations (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3429122A (en) * | 1966-11-07 | 1969-02-25 | Martin Marietta Corp | Heat pipe regenerator for gas turbine engines |
US4207027A (en) * | 1976-08-12 | 1980-06-10 | Rolls-Royce Limited | Turbine stator aerofoil blades for gas turbine engines |
US5178514A (en) * | 1983-05-26 | 1993-01-12 | Rolls-Royce Plc | Cooling of gas turbine shroud rings |
US5439351A (en) * | 1977-07-22 | 1995-08-08 | Rolls-Royce, Plc | Heat pipes |
US5975841A (en) * | 1997-10-03 | 1999-11-02 | Thermal Corp. | Heat pipe cooling for turbine stators |
US6109019A (en) * | 1996-11-29 | 2000-08-29 | Mitsubishi Heavy Industries, Ltd. | Steam cooled gas turbine system |
US20020012588A1 (en) * | 2000-05-31 | 2002-01-31 | Minoru Matsunaga | Gas turbine engine |
US20120195744A1 (en) * | 2010-12-30 | 2012-08-02 | Naik Subhash K | Engine hot section component and method for making the same |
Family Cites Families (14)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3355883A (en) * | 1966-01-24 | 1967-12-05 | Gen Motors Corp | Closed loop heat exchanger for a gas turbine engine |
GB1302036A (en) * | 1969-06-26 | 1973-01-04 | ||
US3621908A (en) * | 1970-09-04 | 1971-11-23 | Dynatherm Corp | Transporting thermal energy through a rotating device |
US3999377A (en) * | 1974-01-16 | 1976-12-28 | Oklejas Robert A | Tesla-type turbine with alternating spaces on the rotor of cooling air and combustion gases |
US4186559A (en) * | 1976-06-07 | 1980-02-05 | Decker Bert J | Heat pipe-turbine |
SU1521284A3 (en) * | 1985-02-02 | 1989-11-07 | Проф.Др.-Инж.Др.-Инж. Е.Х.Клаус Книциа (Фирма) | Power plant |
JPS62170731A (en) * | 1986-01-22 | 1987-07-27 | Kawasaki Heavy Ind Ltd | Gas turbine provided with heat exchanger |
US5082050A (en) * | 1990-05-29 | 1992-01-21 | Solar Turbines Incorporated | Thermal restraint system for a circular heat exchanger |
US5249921A (en) * | 1991-12-23 | 1993-10-05 | General Electric Company | Compressor outlet guide vane support |
JPH06280797A (en) * | 1993-03-30 | 1994-10-04 | Mitsubishi Heavy Ind Ltd | Cooling device for gas turbine |
JPH11117810A (en) * | 1997-10-16 | 1999-04-27 | Ishikawajima Harima Heavy Ind Co Ltd | Rotor system curvature preventing device for gas turbine engine |
DE10029060A1 (en) * | 2000-06-13 | 2002-01-24 | Rolls Royce Deutschland | Turbo air jet engine with heat exchanger |
US20020124569A1 (en) * | 2001-01-10 | 2002-09-12 | Treece William D. | Bimetallic high temperature recuperator |
US6711902B2 (en) * | 2001-06-04 | 2004-03-30 | Richard E. Douglas | Integrated cycle power system and method |
-
2006
- 2006-01-20 US US11/336,330 patent/US7600382B2/en not_active Expired - Fee Related
- 2006-07-14 WO PCT/US2006/027818 patent/WO2007015847A2/en active Application Filing
-
2009
- 2009-09-11 US US12/557,616 patent/US20100000217A1/en not_active Abandoned
Patent Citations (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3429122A (en) * | 1966-11-07 | 1969-02-25 | Martin Marietta Corp | Heat pipe regenerator for gas turbine engines |
US4207027A (en) * | 1976-08-12 | 1980-06-10 | Rolls-Royce Limited | Turbine stator aerofoil blades for gas turbine engines |
US5439351A (en) * | 1977-07-22 | 1995-08-08 | Rolls-Royce, Plc | Heat pipes |
US5178514A (en) * | 1983-05-26 | 1993-01-12 | Rolls-Royce Plc | Cooling of gas turbine shroud rings |
US6109019A (en) * | 1996-11-29 | 2000-08-29 | Mitsubishi Heavy Industries, Ltd. | Steam cooled gas turbine system |
US5975841A (en) * | 1997-10-03 | 1999-11-02 | Thermal Corp. | Heat pipe cooling for turbine stators |
US20020012588A1 (en) * | 2000-05-31 | 2002-01-31 | Minoru Matsunaga | Gas turbine engine |
US20120195744A1 (en) * | 2010-12-30 | 2012-08-02 | Naik Subhash K | Engine hot section component and method for making the same |
Cited By (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20150008611A1 (en) * | 2012-02-17 | 2015-01-08 | Konica Minolta, Inc. | Method and apparatus for production of an obliquely stretched long film |
WO2014099482A1 (en) * | 2012-12-18 | 2014-06-26 | United Technologies Corporation | Oscillating heat pipe for thermal management of gas turbine engines |
US9790893B2 (en) | 2013-03-14 | 2017-10-17 | Rolls-Royce North American Technologies, Inc. | Gas turbine engine flow duct having integrated heat exchanger |
Also Published As
Publication number | Publication date |
---|---|
WO2007015847A2 (en) | 2007-02-08 |
US20070017208A1 (en) | 2007-01-25 |
WO2007015847A3 (en) | 2009-04-16 |
US7600382B2 (en) | 2009-10-13 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US7600382B2 (en) | Turbine engine with interstage heat transfer | |
US6295803B1 (en) | Gas turbine cooling system | |
US11047307B2 (en) | Hybrid expander cycle with intercooling and turbo-generator | |
US11143106B2 (en) | Combustion section heat transfer system for a propulsion system | |
CN110529256B (en) | Air cycle assembly for a gas turbine engine assembly | |
US9410451B2 (en) | Gas turbine engine with integrated bottoming cycle system | |
JP6702636B2 (en) | Power generation system and method for generating power | |
JP6739956B2 (en) | Turbine engine with integrated heat recovery and cooling cycle system | |
JPH05340269A (en) | Gas turbine, heat transfer apparatus and cooling system for gas turbine | |
WO2005003533A2 (en) | High compression gas turbine with superheat enhancement | |
EP2825749A2 (en) | Structures and methods for intercooling aircraft gas turbine engines | |
JP2009185813A (en) | Device and method for starting of power generation plant | |
US5287694A (en) | Fluid channeling system | |
EP3683421A1 (en) | Work recovery system for a gas turbine engine utilizing a recuperated supercritical co2 cycle driven by cooling air waste heat | |
RU2478811C2 (en) | Ventilation and supercharging of turbo-machine components | |
EP3851652B1 (en) | Supercritical co2 cycle for gas turbine engines using partial core exhaust flow | |
US9995216B1 (en) | Disc turbine engine | |
US20180202360A1 (en) | Rotor Shaft Cooling | |
US8448447B2 (en) | Gas turbine engine with fuel booster | |
US11428162B2 (en) | Supercritical CO2 cycle for gas turbine engines using powered cooling flow | |
RU2735880C1 (en) | Method of using gas-air thermodynamic cycle for increasing efficiency of small turbo-engine | |
RU2349775C1 (en) | Nuclear gas-turbine aviation engine | |
JP7184474B2 (en) | Wheelspace temperature control system and method | |
US20140216045A1 (en) | Gas turbine with improved power output | |
RU2391516C2 (en) | Steam-gas installation |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
STCB | Information on status: application discontinuation |
Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION |