|Publication number||US3632221 A|
|Publication date||Jan 4, 1972|
|Filing date||Aug 3, 1970|
|Priority date||Aug 3, 1970|
|Also published as||CA951921A1, DE2121069A1|
|Publication number||US 3632221 A, US 3632221A, US-A-3632221, US3632221 A, US3632221A|
|Inventors||Donald E Uehling|
|Original Assignee||Gen Electric|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (6), Referenced by (19), Classifications (11)|
|External Links: USPTO, USPTO Assignment, Espacenet|
O ijited States Patent 11113,632,z21 Inventor 39 g z a 3,043,56l 6/1962 Schepcr, Jr 415/1 15 1n nna o pp No 60,320 FOREIGN PATENTS g 3 Netherlands 415 16  Patented Jan- 4, 1972 Primary Examiner-Henry F. Raduazo 1 Asslgnee General Eleetrle p y Attorneys-Derek P. Lawrence, Thomas J. Bird, 11., Lee H. Sachs, Frank L. Neuhauser, Oscar B. Waddell and Joseph B. F
54] GAS TURBINE ENGINE COOLING SYSTEM INCORPORATING A VORTEX SHAFT VALVE 14 Chill, 5 Drawing Figs- ABSTRACT: A gas turbine engine cooling system is disclosed 52 US. Cl 415/1 15, which has the eepehilhy of varying the amount of Cooling flew 4 5/D]G 1, 415,176 I delivered to the turbine components. Cooling flow is bled  Int. Cl Fold 5/18 from the compressor through a p Compressor which 50] Field of Search 415/176, forms a vortex Chamber through which the Cooling DIG 1 1 15 1 16, 175 passes. A plurality of small nozzles are positioned within this chamber with their outlets directed opposite the direction of  References Cited rotation of the split disc. Control flow is directed through UNITED STATES PATENTS these nozzles to counteract a natural vortex which is formed within the chamber due to rotation of the disc. The strength of l fgg z the vortex field is then utilized to control the amount of cool- 2'79l 5/1957 wheagey i 415 6 ing flow which passes through the chamber and thus ultimate- 2:9l0, 268 10/1959 Davies et al. 4l5/DIG. 1 to the turbine Swim iNVENT OR. DONALD E. usuuus GAS TURBINE ENGINE COOLING SYSTEM INCORPORATING A VORTEX SHAFT VALVE BACKGROUND OF THE INVENTION This invention relates generally to a gas turbine engine cooling system and, more particularly, to such a system which incorporates a vortex device to regulate the amount of cooling flow delivered to the turbine section of the engine.
In a desire to increase gas turbine engine performance levels, manufacturers are continually attempting to raise the overall operating temperature of such engines. In presently used engines, combustor gas temperatures of between l,850 and 2,550 F. have become commonly used. In spite of this fact, manufacturers are continually striving to further increase the operating temperature levels.
Gas turbine engines of the type generally referred to herein normally comprise a compressor, a primary combustion system, a turbine, a tailpipe (possibly including an augmenter combustion system), and a variable area exhaust nozzle. Turbofan engines, which include additional fan stages and fan turbine drives, utilize gas generators as described in the previous sentence to drive the fan section. Conventionally in such an engine or gas generator, air enters an inlet and is compressed within the compressor, ignited along with metered fuel in the primary combustion system to generate a high-energy gas system, performs work while expanding through the turbine, and exits through the variable area nozzle. The high energy associated with the gas exiting from the exhaust nozzle provides forward thrust to an aircraft powered by such an engine.
The problem generally associated with increasing the gas temperature of the primary airstream generated in the combustion system described above is that the components located downstream of the combustion system are incapable of withstanding such high temperatures. The components seeing the highest temperatures, other than the combustion system components themselves, are the turbine vanes and blades located immediately downstream of the combustors. In order to reduce the temperatures of these turbine vanes and/or blades to levels within material limitations, it has become common practice to provide a flow of relatively cool air through the blades. As a normal rule, this cooling air is taken from the primary airflow engine compressor somewhere along the axial length thereof and is piped to the turbine section. The cooling air is passed through the turbine blades through openings provided in a face of a rotor which supports the turbine blade and which is exposed to a cooling chamber pressurized by compressor discharge air. Altemately, the turbine wheel(s) or disc(s) may also be cooled by flowing the cooling air on the way to the turbine blades through radial passageways in the turbine wheel.
As previously discussed, engine gas temperatures are increased in order to improve gas turbine engine performance levels. The peak temperatures previously mentioned, however, are normally not seen during an entire engine flight mission. That is, as engine thrust levels are varied by the pilot, the gas temperatures within the engine also vary. (A representative plot of engine rotor speed versus turbine inlet temperature is shown in FIG. 1.) As previously mentioned, turbine cooling air is normally bleed air extracted from the compressor. In order to optimize engine efficiency, however, it is necessary to minimize the amount of cooling air which is extracted from the primary airflow during engine operation. Thus, it is highly desirable to have the capability to vary the amount of cooling air delivered to the turbine and turbine disc as the temperatures thereof vary.
The hardware associated with prior attempts to vary the amount of cooling air with turbine blade cooling requirements has been rather complicated and heavy. Additionally, many prior art devices have required a number of moving components in order to vary the amount of cooling air. In fact, no such device to modulate turbine cooling air is known to be in common use.
SUMMARY OF THE INVENTION It is an object of the present invention, therefore, to provide a simple, lightweight and reliable apparatus for varying the amount of cooling flow delivered to a turbine section of a gas turbine engine.
It is a further object of this invention to provide such an apparatus which operates on a purely fluidic principle and which is capable of being positioned entirely within a compressor drum.
Briefly stated, the objects of this invention are carried out by providing an aircraft gas turbine engine with at least one compressor disc which is separated or split to provide a passageway for the flow of cooling air from the circumference to the bore thereof. Located within the passageway and connected to the compressor disc are a plurality of air jets having outlets directed in a direction opposite the direction of rotation of the compressor disc. These air jets are fluidically connected to a source of higher pressure air than that associated with the passageway. Valve means are provided to vary the amount of pressurized air delivered to the air jets. The air jets are utilized to vary the strength of a natural vortex which tends to be generated in the cooling flow within the cooling passageway as the compressor disc rotates. The strength of I this vortex field is then utilized to control the amount of cooling flow which passes through the passageway, and thus the amount of flow which passes to the turbine section located downstream of the compressor.
DESCRIPTION OF THE DRAWINGS While the specification concludes with a series of claims which particularly point out and distinctly claim the subject matter which applicant regards as his invention, an understanding of this invention will be gained from the following detailed description of a preferred embodiment, which is given in light of the accompanying drawings, in which:
FIG. 1 is a graphical plot of turbine inlet temperature versus engine rotor speed;
FIG. 2 is a schematic view, in axial cross section, of a portion of a gas turbine engine constructed in accordance with this invention;
FIG. 3 is an enlarged view of portions of FIG. 2;
FIG. 4 is a view taken along lines 4-4 of FIG. 3; and
FIG. 5 is an axial cross-sectional view of an alternative embodiment.
DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT Referring now to the drawings wherein like numerals correspond to like elements, reference is specificaliy made to FIG. 1 wherein a plot of turbine inlet temperature versus engine speed for a typical gas turbine engine or turbofan engine is shown. This plot shows that the temperature seen by a turbine components varies greatly with engine speed. In order to assure safe operation of the engine, of course, the amount of cooling flow delivered to the turbine components must be capable of maintaining the components at an acceptable temperature when turbine inlet temperature is at its maximum (generally at maximum engine speed). This amount of cooling flow is much higher than that required, however, when the turbine inlet temperature has substantially decreased (at lower engine speeds). Therefore, in order to maintain engine efficiency it is desirable to provide some means for varying the amount of cooling flow delivered to the turbine.
In light of the above, reference is made to FIG. 2 wherein a gas turbine engine 10 is shown to include an outer casing 12 in which air is pressurized by a compressor 14 for delivery to a combustor 16. As is generally well known, fuel is injected into the combustor 16 through a plurality of fuel nozzles 18 (only one of which is shown), and the resultant fuel/air mixture is ignited to generate a high-energy gas stream.
The high-energy gas stream, which is designated by the numeral 20, then passes to a turbine 22 through a stationary inlet nozzle 24 which is positioned at the downstream end of the combustor 16. The high-energy gas stream then passes through an engine exhaust nozzle (not shown) to provide a propulsive force to the engine 10.
Air flows through the compressor 14 within an annular flow path which is defined at its inner bounds by the circumference of a series of interconnected discs 26 having circumferentially mounted compressor blades 28 extending into the gas stream. The outer bounds of the annular flow path are defined by the outer engine casing 12 which has circumferentially mounted stator vanes 30 extending into the gas stream between each row of rotating blades 28.
"The annular flow path continues through an outlet guide vane 32 positioned downstream of the compressor 14. Just aft of the outlet guide vane 32 a snout assembly 34, which forms the upstream end of the combustor l6, breaks the annular flow path into cooling passageways 36 and 38 and diffuser passageway 40. Thus, a portion of the compressor discharge air flows around the combustor 16 to cool the same, while the remaining air enters the combustor 16 through the diffuser passage 40 and is ignited therein to form the high-energy gas stream 20.
The high-energy gas stream thus provided next flows through an annular flow path for the turbine 22, which flow path is defined by the circumferences of a pair of discs 42 which are interconnected for rotation by a heat shield 44 and an annular torque member 46. Circumferentially mounted, hollow turbine blades 45 extend from the discs 42 into the annular flow path thus described. The compressor discs 26 and the turbine discs 42 are interconnected for rotation by opposing conical rotor elements 52, 54, and a tubular shaft 56, which cooperate to form what is normally referred to as the engine rotor. The engine rotor is joumaled for rotation by means which are not shown as they form no part of the present invention but which would include thrust bearings, frame members, etc. These members would be lubricated in any manner known in the art.
As previously described, the temperature of the high-energy gas stream 20 is sufficiently high that it is desirable to provide cooling air to certain components of the turbine section. For this reason, as shown clearly in FIG. 3, one or more of the compressor discs 26 is separated into two separated disc portions 62 in order to provide a cooling air passageway 64 therebetween. The cooling passageway 64 is fluidically connected to the compressor annular flow path by means of a plurality of radial holes 66 bored through the circumference of the compressor disc. The cooling passageway 64 thus delivers a certain percentage of the gas flow from the compressor annular flow path to a plenum chamber 68 (FIG. 2) formed partially by the conical rotor element 52, the tubular shaft 56, and the circumferential portions of the compressor discs 26.
The cooling air next flows into a second plenum chamber 70 through a plurality of airholes 72 positioned within the conical rotor element 52. The plenum chamber 70 is formed by the conical rotor elements 52 and 54 and by a generally cylindrical shaped flow member 74 which is supported on its opposite ends by the outlet guide vanes 32 and the turbine inlet nozzle 24. The wall member 74 further defines the cooling passage 38 previously discussed.
As clearly shown in FIG. 2, the cooling air which is generally designated by the numerals 76 next passes into a third plenum chamber 78 formed by the conical rotor element 54, the tubular shaft 56, and a rotor support element 80 connected to the downstream end of one of the turbine discs 42. Air flowing from the plenum chamber 70 into the plenum chamber 78 does so through one or more airholes 82 positioned in the conical rotor element 54.
The cooling air 76 next flows between the turbine discs 42, thus cooling the outer surfaces thereof. The air continues to flow through an opening 84 fonned in the annular torque member 46 and thus into a chamber 88 formed by the annular torque member 46 and the turbine heat shield 44. The cooling air is then delivered to the interior portions of the hollow turbine blades 45 by means of radial passageways 88 extending through the turbine discs 42. The cooling air enters the radial passageways 88 through holes 90 formed in the turbine heat shield 44. After this cooling air has performed its function of cooling the turbine blades, it is allowed to mix with the highenergy gas stream 20 prior to its flowing through the exhaust nozzle (not shown) as is well known in the art.
As previously mentioned, the principal object of the present invention is to provide a simple. and reliable means for varying the amount of cooling flow which is delivered to the turbine blades. For this purpose, as is clearly shown in FIGS. 3 and 4, a plurality of small jet nozzles 92 are positioned within the passageway 64 along the inner wall of one of the separated disc portions 62. The jet nozzles 92 are positioned along the separated disc portion 62 in such a manner that outlets 94 associated with each jet nozzle 92 are adapted to deliver pressurized air in a direction opposite the direction of rotation of the separated disc portion 62.
Each jet nozzle 92 is connected with a source of pressurized air by means of a tube 96 which is connected to the outside of the separated disc portion 62 as clearly shown in FIGS. 2 and 3. The tubing 96 is, in turn, connected to a control valve 98, which may be attached to the tubular shaft 56 for rotation therewith. The control valve 98 varies the amount of pressurized air delivered to the small jet nozzles 92 as will presently be discussed.
As shown in FIG. 3, the control valve 98 is connected to a source of pressurized air by a second tubing 100. If desirable, multiple valves 98 and lines 100 may be provided. While the pressurized air may be taken from any location, in the present instance the tubing 100 is connected to the compressor rotor stage which includes the separated disc portions 62. For this purpose, an air passageway consisting of one or more holes 102 is provided in the circumference of one of the downstream compressor discs 26. Should multiple holes 102 be provided, the control flow therefrom could be manifolded thereby balancing circumferential extraction pressure. (As is generally known in the art, the pressure of the gas stream increases the air flows downstream in an axial flow compressor. Therefore, the pressure of the air flowing through the airholes 102 is higher than that flowing through the airholes 66.) Depending upon the application, air from the tubing 96 could be delivered through a manifold formed within one of the separated disc portions 62 for delivery to the jet nozzles 92 or, in the alternative, each radial row of jet nozzles 92 could be provided with a separate tube 96 and control valve 98.
As previously discussed, the compressor discs 26 and thus the separated disc portions 62 form part of the compressor rotor. That is, the compressor discs 26 rotate in order to provide rotation of the compressor blades 28. Because of this rotation, the cooling air flowing through the cooling passageway 64 tends to form a natural vortex within the passageway 64. In order to overcome this vortex, prior art devices were equipped with radially extending paddles or vanes which prevented the formation of the vortex. The present device not only eliminates the need for the radial vanes, but also utilizes the natural vortex to control the amount of cooling flow which flows through the passageway 64 and thus to control the amount of cooling air which is delivered to the turbine discs 42 and to the interiors of the turbine blades 45, as required.
Modulation of the amount of cooling air delivered to the turbine blades is accomplished in the following manner. When the control valve 98 is open, control air flows through the airholes 102, through the tubing 100, through the control valve 98, through the tubing 96, and thus out of the small jet nozzles 92. When air is flowing from these jet nozzles 92, an aerodynamic bafile is fonned within the cooling passageway 64 to preclude formation of the natural vortex in much the same manner as radially extending vanes would do. The number of the jet nozzles 92, their radial position, and the magnitude of the angle 9 (FIG. 4) would depend upon the design speed and the size of the separated disc portions 62. The number and position would best be determined by analytical and experimental development.
In a preferred embodiment, the control valve 98 responds to centrifugal force such that maximum control flow occurs at maximum compressor rotor speed. Thus, at maximum rotor speed (and in a normal situation thus at maximum turbine inlet temperature as shown in FIG. 1) maximum control flow passes through the jet nozzles 92 thus assuring prevention of vortex buildup within the passageway 64. Cooling flow through the radial holes 66 is thus unimpeded by a vortex, and maximum flow through the radial holes 66 occurs. The actual amount of cooling flow would, of course, depend on the size and number of the radial holes 66, the pressure drop across the holes 66, and normal losses associated with such a flow system.
As compressor rotor speed (and turbine inlet temperature) decrease, the control valve 98 allows less control flow therethrough; and thus the flow through the jet nozzles 92 decreases. As the flow through the nozzle 92 decreases, a partial vortex forms within the passageway 64 and the pressure buildup associated with such a vortex results in decreased cooling flow through the radial holes 66. That is, as the vortex builds up within the passageway 64, the pressure drop across the radial holes 66 decreases and thus the cooling flow through the radial holes 66 decreases.
It should be pointed out that with the device shown in FIGS. 2 through 4, when maximum cooling flow is required the bleed flow through the airholes 102 used for vortex flow control actually augments the main cooling flow through the radial holes 66 after it has performed its function of preventing vortex formation. When minimum cooling flow through the radial holes 66 is required, the flow through the airholes 102 is essentially zero. While the ratio of control flow through the airholes 102 to the cooling flow through the radial holes 66 would vary from system to system, a desirable ratio may be on the order of l:5.
While any type valve could be utilized for the control valve 98, as previously discussed, a desirable type valve would be a centrifugally actuated type. In any case, the control valve 98 would be designed to fall open, so that in the event of failure, maximum cooling flow is provided to the turbine.
As shown in the alternative embodiment pictured in FIG. 5, the control valve or valves 98 could be positioned at the periphery of one of the compressor discs 26. In this case, the control flow line(s) 100 runs along the outer portion of the disc 26 to a manifold 104 positioned either externally or internally of one of the separated disc portions 62'. The air then flows from the manifold 104 to the jet nozzles 92 to prevent formation of a vortex within the passageway 64 as described in conjunction with FIGS. 2 through 4.
In both of the embodiments shown and described above, applicant has provided a control system which is entirely contained within the compressor drum. In both cases, the only movable parts in the entire system are those associated with the control valve or valves 98. From the above description, it should be readily apparent that applicant has provided a simple, reliable, and lightweight apparatus for controlling the amount of cooling flow delivered to a turbine portion of a gas turbine engine.
While certain embodiments of applicants invention have been described in detail, it should be apparent from reading the above description that changes could be made in applicants basic device without departing from the broader scope of applicants invention. For example, while the above description speaks solely of an aircraft gas turbine engine, it should be readily apparent that applicants device would be applicable to any turbojet, turbofan, or turboshaft engine which utilizes a gas generator or core engine which requires a variable cooling pattern. Likewise, while the above description was limited to a case wherein the outlets of the nozzles 96 were directed opposite the rotational direction of the comprcssor discs, the apparatus would likewise operate with nozzles directed in the opposite direction to augment the natural vortex which is formed in the chamber. It is intended, therefore, that the appended claims cover these and any other changes which come within the broader scope of applicants invention.
What I claim is:
1. In a gas turbine engine having an annular gas stream, a rotating member including a compressor rotor having a plurality of compressor blades projecting into said gas stream and a turbine rotor having a plurality of turbine blades projecting from turbine discs into said gas stream, said turbine blades having passageways for cooling purposes, and said compressor rotor including means for forming a vortex chamber,
first passageway means for providing a flow path from said gas stream to said vortex chamber,
second passageway means for providing a flow path from said vortex chamber to said turbine disc and blade passageways, and
means for decreasing the strength of a natural vortex formed within said vortex chamber due to rotation of said rotating member, whereby the amount of flow to said turbine disc and blade passageways is controlled.
2. A gas turbine engine as recited in claim 1 further characterized in that said compressor rotor includes a plurality of discs for supporting said compressor blades, and at least one of said compressor discs comprises separated disc portions defining said vortex chamber.
3. A gas turbine engine as recited in claim 2 wherein said vortex strength decreasing means comprising a plurality of jet nozzles positioned within said vortex chamber, and further including means for delivering control fluid to said jet nozzles.
4. A gas turbine engine as recited in claim 3 wherein said jet nozzles are radially positioned along one of said separated disc portions.
5. A gas turbine engine as recited in claim 4 wherein said means for delivering control fluid comprising tubing fluidically connected to a source of pressurized fluid and valve means for varying the amount of pressurized fluid flowing through said tubing.
6. A gas turbine engine as recited in claim 5 wherein said source of pressurized fluid is said gas stream.
7. A gas turbine engine as recited in claim6 wherein said valve means are centrifugally controlled.
8. A gas turbine engine as recited in claim 7 wherein said valve means permit maximum fluid passage through said tubing at maximum rotational speed of said rotating member.
9. A gas turbine engine as recited in claim 8 wherein said first passageway means comprises one or more radial airholes positioned within a peripheral portion of the compressor disc which includes said separated disc portions.
10. A gas turbine engine as recited in claim 9 further characterized in that said rotating member includes a tubular shaft, a first conical rotor element connected on one end to said tubular shaft and on its opposite end to said compressor discs, and a second conical rotor element connected on one end to said tubular shaft and on its opposite end to said turbine rotor.
ll. A gas turbine engine as recited in claim 10 wherein said first conical rotor element cooperates with said tubular shaft to form a first plenum chamber, said first and second conical rotor elements cooperate to form a second plenum chamber, and said second conical rotor element cooperates with said tubular shaft to form a third plenum chamber.
12. A gas turbine engine as recited in claim 11 wherein said second passageway means comprises one or more airholes positioned in said first conical rotor element and one or more airholes positioned within said second conical rotor element, whereby cooling air is capable of passing from said vortex chamber into said first plenum chamber, from said first plenum chamber into the said second plenum chamber, from said second chamber into said third plenum chamber, and from said third plenum chamber into said turbine blade passageways.
at least one or more radial airholes positioned within said peripheral portion and defining a flow path from said gas stream to said vortex chamber,
passageway means for providing a flow path from said vortex chamber to said turbine blade passageways, and
fluidic means for controlling the strength of a natural vortex formed within said vortex chamber due to rotation of said rotating members, said fluidic means comprising a plurality of jet nozzles positioned within said vortex chamber, and means for providing control fluid to said jet nozzles.
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|U.S. Classification||415/115, 416/193.00R, 416/95, 415/176, 416/193.00A, 415/914|
|Cooperative Classification||Y02T50/675, F02C7/18, Y10S415/914|