WO1992008044A1

WO1992008044A1 - High temperature turbine

Info

Publication number: WO1992008044A1
Application number: PCT/CA1991/000388
Authority: WO
Inventors: Wallace B. Thomson
Original assignee: Transalta Resources Corp
Current assignee: Transalta Resources Corp
Priority date: 1990-10-25
Filing date: 1991-10-25
Publication date: 1992-05-14
Anticipated expiration: 1993-04-25
Also published as: AU8856291A

Abstract

A compressible fluid compressor/turbine rotor has a disc (12) with radial vanes (16, 18) extending outward to the disc rim (14) and axial flow turbine blades (22) extending radially outward beyond the rim. The fluid to be compressed enters a compressor stator sector through the compressor inlet (28) that is located between the shaft and the rim. It then flows radially outward, guided by the rotor vanes, where some compression occurs. The fluid continues to flow radially outward over the turbine blades, cooling them in the process. Finally, the fluid enters the diffuser vanes (34) in the compressor stator sector where compression of the fluid is completed. The fluid may pass through one or more such compression stages including interstage cooling. The fully compressed fluid then is heated by the system heat source and is returned to a turbine stator sector, through stationary turbine nozzles, and passes over the rotating turbine blades delivering power to the rotor (36). These gases may then pass to other turbine stages and to other components in the system, if any, and then to the atmosphere.

Description

fflGH TEMPERATURE TURBINE

TECHNICAL FIELD

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.

BACKGROUND ART A search of the public records of the United States Patent and Trade Mark Office through class 60, subclass 39.43, revealed the patents listed below which relate to various efforts to cool gas turbines:

2,419,689 McClintock 3,685,287 Dooley 4,757,682 Bahniuk

2,611,241 Schulz 3,994,630 Rodgers 4,506,502 Shapiro 3,095,820 Sanborn et al 3,269,120 Sabatiuk 3,290,897 Wilkins 3,310,940 Oetliker The McClintock patent shows a rotor with blades of a single type to perform both compressor and turbine functions. In particular, he illustrates rotor blades with zero degrees of camber. Clearly, this arrangement would produce a prohibitive loss in compressor and turbine efficiency and would wipe out any gains resulting from an increase in the turbine inlet temperature.

Dooley similarly, uses one type of blading to perform both compressor and turbine functions; similarly, Bahniuk has a one type of blading performing both compressor and turbine functions. The Schulz patent shows an ingenious arrangement for cooling using a single type of blade for both compressor and turbine functions but in this instance, the blades are hollow and the compressor function is achieved by passing the supply air radially within the hollow blades.

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.

In addition to the above patents there exists United States

Patent 4,431,371, "Gas Turbine With Blade Temperature Control", issued March 1984 to Wallace B. Thomson, the original assignee being

Rockwell International. 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 idea of partial arc turbine cooling is not new. It was mentioned by Prof. A. Stodola in his classic text "Steam and Gas Turbines" in 1922. The German engineers, Leist and Knoernschild, performed studies on partial arc turbine cooling for jet engines just prior to World War π. Junkers Motor Works had almost completed building a partial arc cooled jet engine by the end of the war. General Electric Co. performed partial arc cooling tests on a J-47 at Evendale, OH in 1956.

DISCLOSURE OF THE INVENTION

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

compressor inlet to the tips of the turbine blades, the blades will have rotated about that many degrees.

While the fluid is being compressed in the vane and diffuser regions, it is at the same time acting as a coolant for the turbine blades. Depending on the application, the fluid being compressed may pass through several such compression (and cooling) stages. Eventually, the fluid reaches the heat source of the system where it is heated to a very high temperature. The fluid then is returned to the turbine where it enters the stationary nozzles in the hot or active sector of the turbine and expands, performing work on the rotating turbine blades. In this region the flow is basically that of an axial flow turbine except that the temperatures may be extremely high without overheating the blades. The expanded fluid then passes to other turbine stages, if any, and then to the rest of the cycle. 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. By suitable selection of the gas temperatures, pressures and flow rates, as well as the extent of the hot and cool arcs, very low blade temperatures can be achieved.

The stationary compressor and turbine sectors are sealed off from one another except where the blades rotate from one sector to the other. Mixing of gases at this location is minimized by designating for approximately equal static gas pressures across these interfaces.

According to one aspect of this invention there is provided 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.

According to another aspect of this invention there is provided 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.

According to a further aspect of the invention, there is provided 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.

According to yet another aspect there is provided a compressible fluid compressor/turbine for the purpose of operating at high turbine inlet temperatures, comprising:

c (a) a rotor having substantially radial compressor vanes over a radially inner portion and axial flow turbine blades over a radially outer portion;

(b) 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; and

(c) a turbine 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.

BRIEF DESCRIPTION OF THE DRAWING

Preferred embodiments of the present invention are shown in the following drawings in which: 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.

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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.

Formed on the disc are 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

r : - :• x ' in Figures 2 and 3. Over the remaining arc of the rotor the turbine blades are not shrouded but permit conventional axial turbine flow through turbine nozzles 36.

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.

In operation 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. When 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.

It will be noticed that up to 100% of the engine fluid flow can be used for cooling rather than just a small percentage that is typical of other turbine cooling methods. Thus the cooling is very effective

' f. P ' which can lead to either high allowable turbine inlet temperatures or very low blade operating temperatures or both of these.

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.

Air drawn into compressor sector Cl of stage 52 through an inlet shown as line 58 in Figure 4 is compressed and delivered through a scroll duct shown as line 60 to a heat exchanger 62 which acts as an intercooler for the gas exiting compressor Cl and, as described hereinafter, as a feedwater heater for the steam cycle. From the intercooler/heat

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.

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.

From the recuperator 74, 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,

C P?^" preheating the air that is headed for combustor 76. Following the recuperator, the combustion gases may deliver heat to a stack cooler 84 that further heats the feedwater, and finally the gases are exhausted in stack 96.

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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Claims

What is claimed is:

1. 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 portion and means for delivering drive fluid to said turbine stator sector, said turbine stator sector being circumferentially separated from said compressor stator sector.

2. A compressor/turbine as claimed in Claim 1 wherein at least some of said turbine blades are formed as extensions of said compressor vanes.

3. A compressor/turbine as claimed in Claim 1 wherein said compressor vanes are formed on a disc and said turbine blades project beyond the rim of the disc.

4. A compressor/turbine as claimed in Claim 2 wherein said compressor vanes are formed on a disc and said turbine blades project beyond the rim of the disc.

5. A compressor/turbine as claimed in Claim 3 wherein said turbine blades are shrouded outside said turbine stator sector.

6. A compressor/turbine as claimed in Claim 4 wherein said turbine blades are shrouded outside said turbine stator sector.

7. A compressor/turbine as claimed in Claim 1 wherein there is more than one compressor stator sector and an equal number of turbine stator sectors.

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8. A compressor/turbine system having multiple stages each comprising a turbine as claimed in Claim 1 wherein an intercooler is disposed between an outlet of a compressor stator sector of one stage and the inlet of a compressor stator sector of a subsequent stage.

9. 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.

10. A compressor/turbine as claimed in Claim 9 wherein means are provided for delivering compressed fluid from the compressor stator sector of a first stage to the compressor stator sector of a second stage and thence to a heat source.

11. A compressor/turbine as claimed in Claim 10 wherein one or more additional stages are disposed between the second stage and the heat source.

12. A compressor/turbine as claimed in Claim 10 wherein an intercooler is provided between the compressor stator sector of the first stage and the compressor stator sector of the second stage.

13. A compressor/turbine as claimed in Claim 10 wherein means are provided for directing drive fluid from said heat source to the turbine stator sector of the second stage and thereafter to the turbine stator sector of said first stage.

SUBSTITUTE SHE

14. 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 and radially 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.

15. A compressor/turbine as claimed in Claim 14 wherein at least some of said turbine blades are formed as extensions of said compressor vanes.

16. A compressor/turbine having at least two stages each as claimed in Claim 14 wherein compressed fluid from said diffuser vanes in said compressor stator sector of one stage is admitted to said rotating compressor vanes of a subsequent stage and hot gases from said rotating turbine blades of said subsequent stage are directed through an additional set of stationary turbine nozzles and expanded at high velocity into said rotating turbine blades of said first stage.

17. A compressor/turbine as claimed in Claim 16 wherein an intercooler is disposed between the compressor stator sector of one stage and the compressor stator sector of a subsequent stage.

18. A compressible fluid compressor/turbine for the purpose of operating at high turbine inlet temperatures, comprising:

! r-r-r m ( •^■ " ^r " ~ ^w to i li i ϋ ii e m-a- i (a) a rotor having substantially radial compressor vanes over a radially inner portion and axial flow turbine blades over a radially outer portion;

(c) a turbine 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 sector.

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