US2626502A - Cooling system for gas turbine blading - Google Patents

Cooling system for gas turbine blading Download PDF

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US2626502A
US2626502A US751164A US75116447A US2626502A US 2626502 A US2626502 A US 2626502A US 751164 A US751164 A US 751164A US 75116447 A US75116447 A US 75116447A US 2626502 A US2626502 A US 2626502A
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turbine
pressure
cooling
air
blades
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Lagelbauer Ernest
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Lagelbauer Ernest
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, 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
    • F02C7/12Cooling of plants
    • F02C7/125Cooling of plants by partial arc admission of the working fluid or by intermittent admission of working and cooling fluid
    • YGENERAL 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
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/60Efficient propulsion technologies, e.g. for aircraft

Description

.lam 27, E953 E. LAGELBAUER COOLING SYSTEM FOR @As TURBINE BLADING 2 SHEETS-SHEET l Filed May 29. 1947 Kun-OOO IMLL:
BY @M1/@Mg Jan. 27, 1953 E. LAGELBAUER COOLING SYSTEM FOR GAS TURBINE BLADING 2 SHEETS-SHEET 2 Filed May 29. 1947 Patented Jan. 27, 1953 UNITED STATES PATENT OFFICE COOLING SYSTEM FOR GAS TURBINE BLADIN Ernest Lagelbauer, New York, N. Y.
Application May 29, 1947, Serial No. 751,164
2 Claims.
the efficient utilization of the heat that is abstracted in cooling the turbine blades.
A further feature is that the principal disadvantages of up-to-date gas turbine plants, namely the insuicient magnitude of the available adiabatic expansion gradient and the detrimental 1ow permissible heat intake temperature are to a large extent overcome by my method.
Additional features and advantages of my cooling method will be apparent as I proceedv with the description.
With reference to the drawings- Fig. l shows a schematic diagram of a dual gas turbine system incorporating my turbine blade cooling method;
Fig. 2 shows a partial sectional view through the primary turbine blading;
Fig. 3 shows a transverse sectional view through the turbine blading; and
Fig. 4 shows a view similar to Fig. 2 of a modiflcation of my invention.
In Fig. 1 I have shown a schematic diagram of a dual gas turbine system incorporating my turbine blading cooling method. Briefly, the component parts of this system, which may be adapted to either aircraft, locomotive or other similar types of load, are as follows:
The high speed single or two stage partial admission gas turbine I drives a low pressure stage air compressor Il and a high pressure stage air compressor I2. Between the low pressure compressor II and the high pressure compressor I2 is an inter-cooler I3 for maintaining the temperature of the air 60 taken in by the low pressure stage Il at a low level. The com- .bustion air 6I delivered by the high pressure stage I2 is conveyed through passage I4 through a regenerator l5 where it is preheated by the exhaust of turbine and thence to the continuous pressure combustion vessel I6. Fuel pump Il sprays fuel 66 at the suitable rate into the combustion vessel I6 and the formed combustion gases 62 are directed through the expansion nozzle I8 upon the turbine blading I9.
A low speed multistage complete admissionv turbine 20 which is energized by the exhaust 63 of the primary turbine IIJ drives a load 2I through a reduction gear 22. In addition, another turbine 23 of suitable design may be used for driving an additional low pressure stage air compressor 24 for supplementing the air 60 delivered by the low pressure compressor II. The intermediate pressurer gas discharge 63 of turbine III is conveyed through duct 25 to a diverter regulator 26 which has the function of apportioning the power gases 63 between turbines 20 and 23 or, in the event of emergency, to release the power gases 63 through duct 21 to the atmospheric exhaust 64. Diverter 28 has the purpose of regulating the air flow division between further compression by compressor I2 and employment in the secondary power cycle of turbine I0 in effecting the cooling of its blading I9 and the subsequent utilization of the acquired available energy content.
A pressure adjuster 29 is used to equalize the delivery pressure of compressor 24 with that of compressor II. Also indicated is a starting motor 30 for furnishing the initial rotating impulse to the turbine I0 and compressors II and I2.
This system is designed to allow operation with extraordinarily high turbine intake temperatures in allowing low combustion air rates, i. e. restricting air dilution of the combustion gases to the largest permissible extent, and with radically higher compression ratios. Such design characteristics are rendered possible owing to the application of a novel system of cooling those turbine blades which are exposed to gases of extreme temperatures, and to the division of the turbine system into a high speed Single stage partial admission turbine and one or more lower speed full admission turbines. In any event the high speed turbine I0 drives the air com-v pressors II and I2 which together are allowed to rotate at extremely high speeds, (over 20,000 R. P. M.) .l This latter feature besides tends towards space and weight economy of the power plant.
The ultra high speed turbine I0 driving the main air compressors II and I2 is designed to operate at a maximum rotational speed allowable with respect to mechanical stresses due to centrifugal forces generated. Consequently, design considerations dictate development of this turbine as a one or two pressure stage action, partial admission type; of course, its bladings I9 may be designed to have some reaction. The turbine nozzle I8 expands the combustion gases 62 as delivered by the continuous pressure combustion vessel I6 at a pressure of about twelve atmospherics or more down to the intermediary back pressure between high pressure turbine I and the low pressure turbine that is, to a pressure of 2 to 6 atmospheres, which is the intake pressure for turbine 2D. Turbine 20 drives the performance load 2|, for instance a locomotive, and in turn exhausts at atmospheric pressure.
The lowpressure turbine 20 is designed as a multistage, complete admission type containing several pressure stages and suitable velocity stages. In order to restrict the size and weight of this turbine to advantageous proportions in,
for instance, aerial application, a reduction gear 22 is interposed between the turbine 20 and the load 2I and the speed reduction ratio is one determining factor as to the number of stages used in the turbine 2%. In design, a suitablercompromise is to be struck between turbine diameter, number and kind of stages and the reduction ratio of the gear in accordance with the particular application.
The employment of an additional low pressure air compressor 24 driven either by turbine 2U directly or by a supplementary turbine 23 shown in Fig. l is optional but recommended for achieving highest eihciencies. This compressor simply relieves the load demand on the turbine l0.
An outstanding feature of the recommended turbine system is the internal cooling of the high pressure turbine I0 by compressed air.
This method affords several important advantages. It makes practicable the use of extraordinary high initial gas temperatures and allows efficient utilization of the heat abstracted in cooling the turbine blades I9. It also eliminates the detrimental eiect of partial admission adopted for the high pressure turbine I0. It provides complete employment of the otherwise only partially active impulse wheel S2 and consequently greater benets from` the respective machinery investment. It allows thev application of exceedingly large expansion gradients in the primary turbine cycle, owing to the small combustion air rate practicable in the primary cycle, and to some extent, in the secondary turbine cycle, owing to the dynamic boost imparted tothe cooling air 55 in the primary turbine I0. It shouldbe noted that in this process the heatreleased at combustion and taken into the thermodynamic process at extraordinary high temperature is subjectedl to a nearly continuous temperature gradient.
The irreversible temperature drop occurring in.
the primaryv turbine I0,is rather small and its detrimental effect thermodynamically insignificant, as can be visualized readily by inspection of a temperature-entropy diagram. The methodv also provides ample augmentationY of the mass flow through the low pressure turbine 20 in a thermodynamically highly efficient ,mannen This method of internal cooling is illustrated diagrammatically by Figs.,2, 3 and 4.1 It consists in passing precompressed air 65 through that sector SI of the turbine wheel 32 which is not occupied by the jet of power gases 62. Thus, while the blades I9 are heated in passing through the sector of Fig. 3 occupied by the exceedingly hot power gases 62, they are cooleddown again as they pass through the stream of cooling air 5 in sector 3l of Fig. 3. Consequently, the blades I?! are maintained at a temperature much lower than that of the combustion gases 62. In this process the cooling air 65 is energized by the turbine wheel 32 acquiring, in addition to heat, a boost in velocity as well as in pressure. With regard to the further mechanical utilization of the heat thus absorbed the air pressure rise eifected should be promoted by design and as far as possible to take place before the absorption of the heat.
Two modifications of this blade cooling system are available. The one of these illustrated by Figs. 2 and 3 is distinguished by simplicity. The secondary air taken from the low air compressor stages II and 24 of Fig. l enters the turbine blading I9 near the root and, due to appropriate guiding and to centrifugal action, passes over the blades I9 in approximately radial directions. Thereby it is heated in cooling the blades I 9 and brought up to a higher velocity potential. From the turbine wheel 32 the air enters a stage of stationary guide blades (not shown in Fig. 2) which guide the air flow smoothly into the duct 25 leading to the low pressure turbine 20 which utilizes the energy acquired by the cooling air in passing through the turbine wheel 32. In addition, the guide blades mentioned may be so shaped and positioned as to act as diffuser; that is, to convert the acquired velocity potential partially into pressure. In such a design an advantageous balance must be struck between losses related to the diffuser action and the subsequent re-expansion in turbine I0 on the one hand and the advantage gained by the accomplished pressure rise and reduction of flow velocity on the other. This advantage consists in that the cooling air 65 needs to be compressed by the air compressor only to a correspondingly lower initial pressure, for at any rate, the secondary Aair must leave turbine I0 at the same back pressurev as the power gases, which is the intermediary back pressure; or with the same air compression ratio, a higher intake pressure for turbine 20 is achieved. 'I'he secondary action of the cooling air amounts to a reduction of the power input to the air compression equipment and a correspondingly increased power flow to the turbine 20. This circumstance aiords an important means for effecting load apportionment between turbine I 0 and turbine 20, namely, by Varying the air flow rate in the cooling air 65 cycle. It should be noted that the pressure Within the turbine casing 35 of turbine III is very nearly that of the intermediary pressure or back pressure against which the turbine I0 discharges.
A modification of my cooling method is shown in Fig. 4. This method employs a distinct compressor stage 40-4I, integral with the turbine I0. The impeller blading 42 is mounted on the turbine shaft 43 ahead of the turbine wheel 44 and imparts mechanical energy, derived from the turbine cycle through the turbine shaft 43 to the cooling air 65 delivered at intermediary pressure level by the regular air compressor II through duct I4. The acquired velocity potential is then partially converted into pressure energy in passing through the appropriately designed blading 46 provided for the cooling air 6 5 and which is a part of the compressor stage 40-4I the blades 46 also guide the air 65 so as to enter the turbine wheel 44 closely in the direction of the leading edge of the blades 45 which, of course, is of importance with regard to avoiding shock losses. In cooling the blading 45 the secondary air 65 therefore absorbs heat at a considerably higher pressure level than in the alternative method described above. Again, the pressure level of the heat intake decidedly determines the thermaleiciency of utilization of the heat.
accesos The primary advantage of this latter method is that it enables absorption of the heat by the cooling air 95 at a higher pressure level throughout, that is, considerably above the intermedi-ate back pressure between turbine i9 and turbine 2Q. It Will be noted that in the more primitive method indicated above, most of the cooling heat is absorbed merely at the intermediate back pressure, if indeed, not slightly below it, and, particularly that the heat due to cooling of the turbine blades is absorbed before any occurring pressure rise, whereas, it is imperative thermodynamicaly that the pressure rise precede the heating.
r1`he second important advantage or" this alternative is the gas dynamic refinement achieved. Farticularly, it should be observed that in the more primitive method described above, there is some interference between the iet oi combustion gases 92 traversing through the turbine blading i5 essentially in axial direction and the stream of cooling air 55 which flows through the turbine blading @l5 in more or less radial directions. Besides, the difference of the absolute velocity of the combustion gases 92 and of the cooling is greater. The interference results in exchange oi momentum between the combustion gases and the cooling air` 95 with the consequence of corresponding energy degeneration. Such losses increase with the involved e velocity difference.
Another possibility or arrangement is similar modification is marked by greater simplicity but lower einciency as compared with the second method.
Necessarily, the combustion gases 62 and the heated cooling air E5 leave turbine it at a common intermediary exhaust pressure although they diier considerably in temperature. Nevertheless, since the content of available energy is not affected materially during the intermixture of gases of equal pressure but oi moderately dir"- ferent temperature level, maintenance of the separation of the combustion gases 32 and of tde heated cooling air 95 is not warranted, particularly in view of the machinery complications which would be involved.
The gas discharge of turbine i5 which is or" a pressure level of 2 to 6 atmospheres is conducted into turbine 29 and possibly to an eventually employed turbine 23, Fig. 1. The employment of regeneration, such as indicated in Fig. 1 by regenerator l5 and duct id is an obvious device for improving the heat economy through preheating the combustion air 5l by waste heat. Yet, in instances of turbine applications where extreme machinery weight reduction is a supreme factor, these devices may be omitted, of course at the cost ci a corresponding impairment oi the overall fuel economy.
lt is emphasized that the described dual gas turbine system with cooling the blades of a primary ultra speed gas turbine is applicable to whatever working substance employed, -for instance steam, and furthermore may be applied for closed gas or steam turbine cycles such as become or" extraordinary importance for ericent utilization of atomic energy for power generation,
While the invention has been described in detail with respect to a present preferred form which it may assume, it is not to be limited to such details and form since many changes and modifications may be made in the invention without departing from the spirit and scope of the invention in its broadest aspects. Hence, it is desired to cover any and all forms and modiiications of the invention which may come within the language or scope of any one or more of the appended claims.
l claim:
1. In a power System, a rotatable shaft, a turbine vvheel thereon having a plurality of blades rotatable in an annular path, an impeller wheel on said shaft adjacent the turbine wheel and having blades thereon, means to pass compressed cooling air into contact with said blades of the impeller wheel, means to pass combustion gases into contact with the exterior surfaces of said turbine blades throughout a predetermined segment or" said annular path, fixed blades positioned between the impeller wheel and turbine wheel to guide the compressed air into contact with said exterior surfaces ci the turbine blades throughout the remaining segment or said annular path.
2. A power system comprising `a high pressure 'turbine and a low-pressure turbine, said high pressure turbine being mounted on a rotatable shaft and having a plurality oi blades rotatable an annular path, a low pressure compressor and a high pressure compressor mounted on said shaft, e, combustion vessel connected to the high pressure compressor, means for passing fuel and high pressure air into the combustion vessel to iorm combustion gases and then to the high pressure turbine into Contact with the exterior surfaces oi said turbine blades throughout a predetermined segment of said annular path, means to pass compressed cooling air from said low pressure compressor into contact with said exterior surfaces or the turbine blades throughout the remaining segment of said annular path, thereby cooling the blades and heating the compressed air, and means to pass the mixture of combustion gases and compressed cooling air to said low pressure turbine to serve yas a thermodynamic working substance.
ERNEST LAGELBAUER.
CES CITED The following reierences are of record in the nie of this patent:
UNITED STATES PATENTS Number Name Date 1,375,931 Rateau Apr. 26, 1921 1,421,087 Johnson June 27, 1922 2,168,726 Whittle Aug. 8, 1939 2,216,731 Birmann Oct. 8, 1940 2,320,391 Vfaleeld June l, 1943 2,364,139 Bchi Dec. 5, 1944 2,419,639 McClintock Apr. 29, 1947 2,442,019 Ray May 25, 1948 2,462,600 Boestad Feb. 22, 1949 2,511,385 Udale June 13, 1950 2,536,962 liano Jan. 2, 1951 FOREIGN PATENTS er Country Date 19,929 Great Britain Got. 5, 1901 375,543 Great Britain June 30, 1932 502,414 Great Britain Mar. 13, 1939 346,599 Germany Jan. 5, 1922 346,611 France Dec. 3, 1904
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Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2960825A (en) * 1955-10-19 1960-11-22 Thompson Ramo Wooldridge Inc Reexpansion gas turbine engine with differential compressor drive
US2976684A (en) * 1951-05-10 1961-03-28 Wirth Emil Richard Improvements in gas turbines
US3312056A (en) * 1964-03-09 1967-04-04 Lagelbauer Ernest Super temperature dual flow turbine system
US3531934A (en) * 1967-11-07 1970-10-06 Charles David Hope-Gill Gas turbine power plant
US3645096A (en) * 1969-01-23 1972-02-29 Georg S Mittelstaedt Peripheral suction openings in gas turbine engines
US3771312A (en) * 1970-11-19 1973-11-13 Automotive turbine engine
EP0206683A2 (en) * 1985-06-17 1986-12-30 University Of Dayton Internal bypass gas turbine engines with blade cooling
US4872675A (en) * 1987-02-17 1989-10-10 Horace Crowden Baseball pitching device
US5347806A (en) * 1993-04-23 1994-09-20 Cascaded Advanced Turbine Limited Partnership Cascaded advanced high efficiency multi-shaft reheat turbine with intercooling and recuperation
US5680752A (en) * 1992-08-28 1997-10-28 Abb Carbon Ab Gas turbine plant with additional compressor
US20070280400A1 (en) * 2005-08-26 2007-12-06 Keller Michael F Hybrid integrated energy production process

Citations (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE346599C (en) *
GB190119920A (en) * 1901-10-05 1902-10-04 Hugh Francis Fullagar Improved Method of and Means or Apparatus for Producing Motive Power from the Combustion of Fuel.
FR346611A (en) * 1904-08-15 1905-02-01 Pierre Rambal Combined gas turbine and steam turbine
US1375931A (en) * 1917-11-06 1921-04-26 Rateau Auguste Camille Edmond Pertaining to internal-combustion aircraft-motors
US1421087A (en) * 1920-03-09 1922-06-27 Johnson Herbert Stone Internal-combustion turbine
GB375643A (en) * 1931-06-09 1932-06-30 Fabrication Des Moteurs Et Tur Improvements in and relating to combustion turbines
GB502414A (en) * 1937-09-13 1939-03-13 Karl Leist Improvements in turbines
US2168726A (en) * 1936-03-04 1939-08-08 Whittle Frank Propulsion of aircraft and gas turbines
US2216731A (en) * 1937-12-02 1940-10-08 Turbo Engineering Corp Turbosupercharger mounting
US2320391A (en) * 1938-09-06 1943-06-01 George H Wakefield Explosion turbine motor
US2364189A (en) * 1940-09-21 1944-12-05 Buchi Alfred Cooling device for turbine rotors
US2419689A (en) * 1942-11-05 1947-04-29 Raymond K Mcclintock Gas turbine
US2442019A (en) * 1943-06-11 1948-05-25 Allis Chalmers Mfg Co Turbine construction
US2462600A (en) * 1943-01-16 1949-02-22 Jarvis C Marble Turbine
US2511385A (en) * 1945-03-14 1950-06-13 George M Holley Two-stage gas turbine
US2536062A (en) * 1948-12-30 1951-01-02 Kane Saul Allan System of blade cooling and power supply for gas turbines

Patent Citations (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE346599C (en) *
GB190119920A (en) * 1901-10-05 1902-10-04 Hugh Francis Fullagar Improved Method of and Means or Apparatus for Producing Motive Power from the Combustion of Fuel.
FR346611A (en) * 1904-08-15 1905-02-01 Pierre Rambal Combined gas turbine and steam turbine
US1375931A (en) * 1917-11-06 1921-04-26 Rateau Auguste Camille Edmond Pertaining to internal-combustion aircraft-motors
US1421087A (en) * 1920-03-09 1922-06-27 Johnson Herbert Stone Internal-combustion turbine
GB375643A (en) * 1931-06-09 1932-06-30 Fabrication Des Moteurs Et Tur Improvements in and relating to combustion turbines
US2168726A (en) * 1936-03-04 1939-08-08 Whittle Frank Propulsion of aircraft and gas turbines
GB502414A (en) * 1937-09-13 1939-03-13 Karl Leist Improvements in turbines
US2216731A (en) * 1937-12-02 1940-10-08 Turbo Engineering Corp Turbosupercharger mounting
US2320391A (en) * 1938-09-06 1943-06-01 George H Wakefield Explosion turbine motor
US2364189A (en) * 1940-09-21 1944-12-05 Buchi Alfred Cooling device for turbine rotors
US2419689A (en) * 1942-11-05 1947-04-29 Raymond K Mcclintock Gas turbine
US2462600A (en) * 1943-01-16 1949-02-22 Jarvis C Marble Turbine
US2442019A (en) * 1943-06-11 1948-05-25 Allis Chalmers Mfg Co Turbine construction
US2511385A (en) * 1945-03-14 1950-06-13 George M Holley Two-stage gas turbine
US2536062A (en) * 1948-12-30 1951-01-02 Kane Saul Allan System of blade cooling and power supply for gas turbines

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2976684A (en) * 1951-05-10 1961-03-28 Wirth Emil Richard Improvements in gas turbines
US2960825A (en) * 1955-10-19 1960-11-22 Thompson Ramo Wooldridge Inc Reexpansion gas turbine engine with differential compressor drive
US3312056A (en) * 1964-03-09 1967-04-04 Lagelbauer Ernest Super temperature dual flow turbine system
US3531934A (en) * 1967-11-07 1970-10-06 Charles David Hope-Gill Gas turbine power plant
US3645096A (en) * 1969-01-23 1972-02-29 Georg S Mittelstaedt Peripheral suction openings in gas turbine engines
US3771312A (en) * 1970-11-19 1973-11-13 Automotive turbine engine
EP0206683A2 (en) * 1985-06-17 1986-12-30 University Of Dayton Internal bypass gas turbine engines with blade cooling
EP0206683A3 (en) * 1985-06-17 1988-08-03 University Of Dayton Internal bypass gas turbine engines with blade cooling
US4872675A (en) * 1987-02-17 1989-10-10 Horace Crowden Baseball pitching device
US5680752A (en) * 1992-08-28 1997-10-28 Abb Carbon Ab Gas turbine plant with additional compressor
US5347806A (en) * 1993-04-23 1994-09-20 Cascaded Advanced Turbine Limited Partnership Cascaded advanced high efficiency multi-shaft reheat turbine with intercooling and recuperation
US5386688A (en) * 1993-04-23 1995-02-07 Cascaded Advanced Turbine Limited Partnership Method of generating power with high efficiency multi-shaft reheat turbine with interccooling and recuperation
US20070280400A1 (en) * 2005-08-26 2007-12-06 Keller Michael F Hybrid integrated energy production process
US7961835B2 (en) 2005-08-26 2011-06-14 Keller Michael F Hybrid integrated energy production process
US8537961B2 (en) 2005-08-26 2013-09-17 Michael Keller Hybrid integrated energy production process

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