WO1992008044A1 - Turbine a temperature elevee - Google Patents

Turbine a temperature elevee Download PDF

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Publication number
WO1992008044A1
WO1992008044A1 PCT/CA1991/000388 CA9100388W WO9208044A1 WO 1992008044 A1 WO1992008044 A1 WO 1992008044A1 CA 9100388 W CA9100388 W CA 9100388W WO 9208044 A1 WO9208044 A1 WO 9208044A1
Authority
WO
WIPO (PCT)
Prior art keywords
turbine
compressor
sector
fluid
stator sector
Prior art date
Application number
PCT/CA1991/000388
Other languages
English (en)
Inventor
Wallace B. Thomson
Original Assignee
Transalta Resources Investment Corporation
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 Transalta Resources Investment Corporation filed Critical Transalta Resources Investment Corporation
Publication of WO1992008044A1 publication Critical patent/WO1992008044A1/fr

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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
    • F02C3/045—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor having compressor and turbine passages in a single rotor-module
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/04—Blade-carrying members, e.g. rotors for radial-flow machines or engines
    • F01D5/043—Blade-carrying members, e.g. rotors for radial-flow machines or engines of the axial inlet- radial outlet, or vice versa, type
    • F01D5/046—Heating, heat insulation or cooling means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • 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]
    • 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
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This invention is concerned with a compressible fluid compressor/turbine.
  • the attainment of very high turbine inlet temperatures in gas turbines while at the same time maintaining very low turbine blade temperatures, is desirable.
  • a major problem in the gas turbine industry since its inception almost a century ago has been achieving turbine inlet temperatures high enough to arrive at a satisfactory cycle efficiency while maintaining adequate turbine blade lifetime.
  • the cooling of blades by having internal cooling flow passages has led to advances in cycle temperatures. In the past thirty years, however, the increase in allowable turbine inlet temperature has been very small.
  • Today 1500° K is about the practical upper limit when aviation fuels or natural gas fuels are used. With coal fuels, the upper limit is very much less.
  • the Rodgers patent seeks to achieve cooling by thermal conduction from the hot turbine blades to the cooler compressor vanes. Additionally, a small quantity of the compressor air is bled to help cool the turbine side.
  • the Shapiro patent is not particularly pertinent; it has separate axial flow compressor blades and turbine blades mounted on a single rotor. For cooling he depends upon conduction of heat from the hot turbine blades to the cooler compressor blades plus the compressor bleed air is used.
  • the Sanborn patent deals with a drag pump that has bleed ports to reduce unwanted circulation. It is unrelated to the present application but it is included merely for completeness of disclosure.
  • the Sabatiuk patent seeks to provide cooling by thermal conduction through metal walls. It is not pertinent.
  • the Oetliker patent shows a turbine with blades of different types of opposite sides of a single rotor. It is not particularly pertinent.
  • Patent 4,431,371, "Gas Turbine With Blade Temperature Control", issued March 1984 to Wallace B. Thomson, the original assignee being
  • This invention uses partial arc turbine cooling with radial cooling flow over the axial flow turbine blades.
  • the invention does not incorporate this flow with a centrifugal compressor but rather, makes use of axial flow compressors.
  • the 4,431,371 invention has a superior compressor efficiency.
  • the present invention is based on partial arc turbine cooling.
  • Part of the turbine arc is reserved for the axial passage of hot gases which deliver power to the rotor, while the remainder of the arc is used to permit cooling air to pass more or less radially over the external surfaces of the blades. Since the blades pass rapidly from hot to cool regions, they never become too hot but take on a temperature very much lower than that entering the turbine in the hot sector.
  • the advantages of partial arc turbine cooling over other cooling methods are: (1) the amount of cooling air is large so that the blades can be cooled to a lower temperature, (2) the cooling occurs directly on the external surfaces of the blades where cooling is most needed, (3) cooling air is not extracted from the main flow and then dumped overboard so there is no performance degradation from that source, (4) the entire blade surface is well cooled including such critical points as the leading edges, trailing edges and blade tops, (5) blade material is not removed for cooling passages within the blades so that the blades are stronger and stiffer and (6) there are no small coolant passages to become plugged.
  • the present invention seeks to use the air being compressed as the coolant for the turbine blades and to perform such cooling in the most direct way with a maximum of cooling effectiveness and a minimum of mechanical complexity and cycle thermodynamic losses.
  • the present invention seeks to provide a compressible fluid compressor/turbine of the kind in which a stage has a single rotor comprising a disc on which are mounted radial vanes as in a centrifugal compressor. Extending radially outward beyond the disc rim are turbine blades like those in an axial flow turbine. Most desirably, at least some of the turbine blades are formed as extensions of the compressor vanes.
  • the fluid to be compressed is directed into the rotating vanes through a compressor inlet sector stationary relative to the engine frame. The fluid flows radially outward through the vane region, continues radially outward over the turbine blades and enters a diffiiser section where compression in that stage is completed.
  • the fluid flow is admitted to the rotating compressor impeller through a stationary sector near the engine shaft. This sector subtends an arc that has about the same angular size as the turbine blade cooling arc. However, the admission sector will be displaced possibly 180 degrees from the cooling arc because as the fluid travels from the
  • the blades are swept alternately by a radial flow of relatively cool compressor gases and by an axial flow of hot turbine gases. With this alternate cooling and heating the blades take on an intermediate temperature far below that of the gases from the heat source.
  • gas temperatures, pressures and flow rates, as well as the extent of the hot and cool arcs very low blade temperatures can be achieved.
  • a compressible fluid compressor/turbine comprising a rotor having compressor vanes over a radially inner portion and turbine blades over a radially outer portion, a compressor stator sector, means for admitting fluid to said compressor stator sector, a turbine stator sector in said radially outer portions and means for delivering drive fluid to said turbine stator sector, said turbine stator sector being circumferentially separated from said compressor stator sector.
  • a compressible fluid compressor/turbine comprising a rotor and a stator in which said stator is divided into a compressor sector and a turbine sector, said sectors being circumferentially separated by partition means, wherein said rotor has compressor vanes over a radially inner portion and axial flow turbine blades over a radially outer portion; said compressor stator sector comprises means for admitting fluid to be compressed to the rotating compressor vanes, means to recover the dynamic pressure of said fluid leaving the rotating turbine blades, and means to direct the compressed fluid to a heat source; and said turbine stator sector comprises means to direct gases from the heat source to said turbine stator sector, turbine nozzles to expand said gases at high velocity into the rotating turbine blades, and exit means for the expanded gases.
  • a plural stage compressor/turbine comprising a first stage rotor having compressor vanes over a radially inner portion and turbine blades over a radially outer portion, a compressor stator sector, means for admitting fluid to said compressor stator sector, a turbine stator sector in said radially outer portion, and means for delivering drive fluid to said turbine stator sector, said turbine stator sector being circumferentially separated from said compressor stator sector.
  • a compressible fluid compressor/turbine for the purpose of operating at high turbine inlet temperatures, comprising:
  • a compressor stator disposed in a first sector and having inlet means to admit the fluid to be compressed to the radially inner portion of the rotor, said vanes and blades forming fluid flow passages from the inlet means to the radially outermost tips of the turbine blades, means to recover the dynamic pressure of the fluid leaving the tips of said turbine blades in said compressor stator sector, duct means to lead the compressed fluid away from the rotor;
  • stator disposed in a second sector said stator having nozzle means to direct fluid substantially axially into the turbine blades in said second sector, duct means to lead the fluid from the turbine blades in said second sector and partition means to separate the turbine stator sector from said compressor stator assembly.
  • Figure 1 is a front view of a single stage of a compressible fluid turbine according to this invention
  • Figure 2 is a schematic side view of the turbine stage shown in Figure 1;
  • Figure 3 is a schematic side view cutting through the nozzles and blades and shows the stationary partition that separates the cool and hot sectors;
  • Figure 4 is a diagram of a multi-stage compressor/turbine according to this invention, shown as part of a combined cycle gas turbine/steam turbine central power station.
  • the turbine stage of Figure 1 comprises a shaft 10 to which rotor or impeller disc 12 is secured.
  • the rotor has radial compressor vanes of various lengths extending to the rim 14 of the disc. Beyond the rim the extensions of these vanes take on the form of axial flow turbine blades 22.
  • a plurality of generally radial vanes some of which at 16 extend from the junction of the rotor with the shaft, others of which at 18 extend outwardly from an imaginary circle indicated at 20, and still others at 19 extend outwardly from an imaginary circle at 24. All of those portions of the vanes within the periphery 14 of the disc are formed as compressor vanes. Each of these vanes projects radially outward beyond the periphery 14 and in this region is shaped in the form of an axial flow turbine blade. Obviously, there are many possible arrangements of vanes and blades similar to Figure 1 that could be devised by those skilled in the art.
  • An annular shroud 26 in Figure 2 extends from the compressor inlet 28 to the base of the blades 22 and is attached to the edges of vanes 16, 18 and 19.
  • the shroud is an optional but desirable feature that serves to support the turbine blades - especially those blades that are extensions of the shorter type vanes 19.
  • the vanes and blades are arranged so that continuous and substantially radial fluid flow paths are formed from the compressor inlet to the tips of the blades. This is a necessary feature that ensures that the compressor fluid flow continues over the blade surface in smooth aerodynamic flow to achieve a maximum cooling effect with a minimum pressure loss.
  • a scroll duct is formed at 30 and extends over a sector, typically 120 to 240 degrees, which defines the compressor sector of the stage. Within this sector the turbine blades are shrouded 32 as shown
  • a plurality of diffuser vanes 34 is disposed within the scroll duct, while within the unshrouded active sector of the turbine there is a plurality of turbine nozzles 36.
  • the stationary inlet to the compressor impeller is a sector that may subtend an angle that can range from 120 to 240 degrees (in Figure 1 it is shown as 180 degrees) and is typically located near the shaft. It is displaced roughly 180 degrees from the scroll 30 and diffuser vanes 34. The reason for this displacement is that the rotor turns through a considerable angle during the time it takes for the fluid to travel from the compressor inlet to the diffuser vanes.
  • a fluid such as air is admitted to the compressor in sector 28 and is centrifiiged radially outward to the diffuser vanes 34 in scroll duct 30 where the dynamic pressure is converted to static pressure as in a conventional centrifugal compressor.
  • the fluid then travels to the next compressor stage, if any, where more compression (and turbine cooling) occurs.
  • the fluid has passed through all compressor stages it eventually reaches the system heat source where it is heated to a very high temperature.
  • the hot combustion gases are then returned to the compressor/turbine but pass through the stationary nozzles 36 and the axial flow turbine blades 22 in the hot or active sectors of the various turbine stages.
  • the combustion gases then may pass through other system components and finally are exhausted to the atmosphere.
  • a key point in this type of cooling is to avoid significant pressure losses in the cooling mechanism, as pressure losses are very important in gas turbine cycles. Since the turbine blades can be made as extensions of the compressor vanes, the flow can be made aerodynamically smooth which leads to low pressure losses. Similarly, this smooth flow leads to good cooling heat transfer on all blade surfaces with a minimum of hot and cool spots. Since the blades are alternately heated and cooled it might be expected that the blade temperature would rise and fall with each revolution. Actually the bulk temperature of a typical blade does not measurably change at all. Only a very thin surface layer significantly changes in temperature - about 0.001 cm or so changes about 20° K in each revolution. This is about the same temperature fluctuation that occurs on the cylinder wall of an automobile engine. The resulting stresses are well below the metal elastic limit and no significant damage results.
  • Figure 4 shows schematically how a gas turbine using the present invention could be incorporated into a combined cycle gas turbine - steam turbine power plant.
  • the gas turbine 50 has three stages 52, 54 and 56 of the type shown in Figures 1, 2 and 3.
  • Stage 52 has a compressor sector Cl and a turbine sector T3;
  • stage 54 has a compressor sector C2 and turbine sector T2, and
  • stage 56 has compressor sector C3 and turbine sector Tl.
  • r exchanger 62 the cooled and partly pressurized gas is delivered through ducting 64 to the inlet of compressor C2 of stage 54 to be further compressed.
  • the air exiting compressor sector C2 is passed through a scroll duct in line 66 to a heat exchanger 68 which, as the heat exchanger 62, acts as an intercooler for the compressed air and as a feedwater heater for the steam cycle.
  • intercooler-feedwater heater arrangement There are several reasons for having the intercooler-feedwater heater arrangement described here.
  • the primary reason is to cool the compressor air flow so that cooling of the turbine blades in the manner of this invention is more effective - especially in coal fired plants where blade temperatures should be very low.
  • a second reason is to reduce the power required by the compressor(s) to compress the air.
  • the third reason is to use the heat rejected by the intercoolers to supply feedwater heating to the steam turbine cycle. From intercooler/heat exchanger 68 the air travels through ducting in line 70 to compressor sector C3 of stage 56 and thence through ducting 72 to recuperator 74 where it receives heat from the steam cycle boiler exit combustion gases.
  • the compressed and preheated air is passed to a combustor 76 to be combined with fuel and reacted to produce very high temperature combustion gases.
  • the combustor could, for this application, be one burning coal, oil or gas.
  • the gases may leave the combustor at temperatures in the order of 1900° K and enter turbine sector Tl of stage 56 thence to turbine sectors T2 and T3, to drive generator 80.
  • the high pressure combustion gases are delivered first to turbine stage 56 (Tl) because the compressor sector of that stage is at the highest pressure and thus the pressure difference across the compressor and turbine sectors in that stage and the gas leakage between these sectors are both minimized.
  • the combustion gases leaving the final turbine stage still at a relatively high temperature, are delivered to boiler 82 to generate steam. After exiting the boiler the combustion gases pass through recuperator 74,
  • Steam is delivered from boiler 82 through line 86 to steam turbine 88 which drives generator 90.
  • the condensate from the steam turbine and condenser 92 is delivered by feedwater pump 94 to the coolant sides of heat exchangers 62 and 68, passes through stack cooler 84 and is then returned to boiler 82.

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

Abstract

Un rotor de compresseur/turbine à fluide compressible comprend un disque (12) pourvu d'aubes radiales (16, 18) s'étendant vers l'extérieur, vers le rebord (14) du disque, et des ailettes (22) de turbine à flux axial s'étendant de manière radiale vers l'extérieur au-delà du rebord. Le fluide devant être comprimé pénètre dans le secteur de stator de compresseur à travers l'entrée (28) du compresseur, située entre l'arbre et le rebord. Il s'écoule ensuite de manière radiale vers l'extérieur, guidé par les aubes du rotor, où un certain niveau de compression a lieu. Le fluide continue à s'écouler radialement vers l'extérieur par-dessus les ailettes de turbine, les refroidissant par la même occasion. Finalement, le guide pénètre dans les aubes diffusantes (34) dans le secteur du stator du compresseur où la compression du fluide est complétée. Le fluide peut passer par un ou plusieurs de ces étages de compression, y compris un étage intermédiaire de refroidissement. Le fluide entièrement comprimé est alors chauffé par la source de chaleur du système, puis retourné à un secteur de stator de turbine à travers des tuyères de turbine fixes, et il passe sur les ailettes de turbine en rotation de manière à fournir de la puissance au rotor (36). Ces gaz peuvent ensuite se diriger vers d'autres étages de turbine et d'autres éléments du système, s'il en existe, puis vers l'atmosphère.
PCT/CA1991/000388 1990-10-25 1991-10-25 Turbine a temperature elevee WO1992008044A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US60322190A 1990-10-25 1990-10-25
US603,221 1990-10-25

Publications (1)

Publication Number Publication Date
WO1992008044A1 true WO1992008044A1 (fr) 1992-05-14

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AU (1) AU8856291A (fr)
WO (1) WO1992008044A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1998013584A1 (fr) * 1996-09-26 1998-04-02 Siemens Aktiengesellschaft Compensation de la perte de pression d'une conduite d'air de refroidissement dans une installation de turbine a gaz

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB196931A (en) * 1922-04-28 1924-01-17 Leon Dufour Improvements in combustion turbines
US2272676A (en) * 1938-12-23 1942-02-10 Leduc Rene Continuous flow gas turbine

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB196931A (en) * 1922-04-28 1924-01-17 Leon Dufour Improvements in combustion turbines
US2272676A (en) * 1938-12-23 1942-02-10 Leduc Rene Continuous flow gas turbine

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1998013584A1 (fr) * 1996-09-26 1998-04-02 Siemens Aktiengesellschaft Compensation de la perte de pression d'une conduite d'air de refroidissement dans une installation de turbine a gaz

Also Published As

Publication number Publication date
AU8856291A (en) 1992-05-26

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