Science.gov

Sample records for air cooled turbine

  1. Modulated gas turbine cooling air

    SciTech Connect

    Schwarz, F.M.; Candelori, D.J.; Brooke, R.D.

    1993-07-06

    In an axial flow gas turbine engine in an aircraft, the engine having an annular flow of hot working fluid passing sequentially through a first bladed rotor stage, a vaned stator assembly having a plurality of hollow vanes, and a second bladed rotor stage; a flow resistant labyrinth seal comprised of an annular seal runner sealingly secured to the first and second rotor stages and a seal shroud surrounding and secured to the seal runner, forming a labyrinth flow passage therebetween; an upstream plenum in restricted fluid communication with the annular flow upstream of the vaned stator assembly and with the labyrinth flow passage; a downstream plenum in fluid communication with the labyrinth flow passage and in restricted flow communication with the annular flow downstream of the vaned stator assembly; a compressor; a conduit network connected to deliver a cooling airflow from the compressor to the upstream plenum, and a modulatable control valve means located in the conduit network, the method of operation comprising: measuring the temperature of gas passing through the labyrinth flow passage; sensing aircraft speed and comparing the sensed speed to a preselected air craft speed range; holding the valves open any item the sensed aircraft speed is less than the preselected aircraft speed range; and modulating he quantity of the cooling airflow in response to the measurement of the temperature of the gas passing through the labyrinth flow passage to keep the temperature at a substantially constant maximum value when the sensed aircraft speed is greater than the aircraft speed range.

  2. Closed-loop air cooling system for a turbine engine

    DOEpatents

    North, William Edward

    2000-01-01

    Method and apparatus are disclosed for providing a closed-loop air cooling system for a turbine engine. The method and apparatus provide for bleeding pressurized air from a gas turbine engine compressor for use in cooling the turbine components. The compressed air is cascaded through the various stages of the turbine. At each stage a portion of the compressed air is returned to the compressor where useful work is recovered.

  3. Closed loop air cooling system for combustion turbines

    DOEpatents

    Huber, D.J.; Briesch, M.S.

    1998-07-21

    Convective cooling of turbine hot parts using a closed loop system is disclosed. Preferably, the present invention is applied to cooling the hot parts of combustion turbine power plants, and the cooling provided permits an increase in the inlet temperature and the concomitant benefits of increased efficiency and output. In preferred embodiments, methods and apparatus are disclosed wherein air is removed from the combustion turbine compressor and delivered to passages internal to one or more of a combustor and turbine hot parts. The air cools the combustor and turbine hot parts via convection and heat is transferred through the surfaces of the combustor and turbine hot parts. 1 fig.

  4. Closed loop air cooling system for combustion turbines

    DOEpatents

    Huber, David John; Briesch, Michael Scot

    1998-01-01

    Convective cooling of turbine hot parts using a closed loop system is disclosed. Preferably, the present invention is applied to cooling the hot parts of combustion turbine power plants, and the cooling provided permits an increase in the inlet temperature and the concomitant benefits of increased efficiency and output. In preferred embodiments, methods and apparatus are disclosed wherein air is removed from the combustion turbine compressor and delivered to passages internal to one or more of a combustor and turbine hot parts. The air cools the combustor and turbine hot parts via convection and heat is transferred through the surfaces of the combustor and turbine hot parts.

  5. Integrated turbine-compressor provides air flow for cooling

    NASA Technical Reports Server (NTRS)

    Ferri, A.

    1970-01-01

    Modified supersonic turbine cycle provides cooling air to surrounding structures. Simplified mechanical design assures correct balance of air flow, allows direct issue of cool air to the structure, and assists in matching turbine work output to work input required by the compressor.

  6. Effect of Chord Size on Weight and Cooling Characteristics of Air-Cooled Turbine Blades

    NASA Technical Reports Server (NTRS)

    Esgar, Jack B; Schum, Eugene F; Curren, Arthur N

    1958-01-01

    An analysis has been made to determine the effect of chord size on the weight and cooling characteristics of shell-supported, air-cooled gas-turbine blades. In uncooled turbines with solid blades, the general practice has been to design turbines with high aspect ratio (small blade chord) to achieve substantial turbine weight reduction. With air-cooled blades, this study shows that turbine blade weight is affected to a much smaller degree by the size of the blade chord.

  7. Air cooling of disk of a solid integrally cast turbine rotor for an automotive gas turbine

    NASA Technical Reports Server (NTRS)

    Gladden, H. J.

    1977-01-01

    A thermal analysis is made of surface cooling of a solid, integrally cast turbine rotor disk for an automotive gas turbine engine. Air purge and impingement cooling schemes are considered and compared with an uncooled reference case. Substantial reductions in blade temperature are predicted with each of the cooling schemes studied. It is shown that air cooling can result in a substantial gain in the stress-rupture life of the blade. Alternatively, increases in the turbine inlet temperature are possible.

  8. Air cooled turbine component having an internal filtration system

    DOEpatents

    Beeck, Alexander R.

    2012-05-15

    A centrifugal particle separator is provided for removing particles such as microscopic dirt or dust particles from the compressed cooling air prior to reaching and cooling the turbine blades or turbine vanes of a turbine engine. The centrifugal particle separator structure has a substantially cylindrical body with an inlet arranged on a periphery of the substantially cylindrical body. Cooling air enters centrifugal particle separator through the separator inlet port having a linear velocity. When the cooling air impinges the substantially cylindrical body, the linear velocity is transformed into a rotational velocity, separating microscopic particles from the cooling air. Microscopic dust particles exit the centrifugal particle separator through a conical outlet and returned to a working medium.

  9. Turbine inter-disk cavity cooling air compressor

    DOEpatents

    Chupp, Raymond E.; Little, David A.

    1998-01-01

    The inter-disk cavity between turbine rotor disks is used to pressurize cooling air. A plurality of ridges extend radially outwardly over the face of the rotor disks. When the rotor disks are rotated, the ridges cause the inter-disk cavity to compress air coolant flowing through the inter-disk cavity en route to the rotor blades. The ridges eliminate the need for an external compressor to pressurize the air coolant.

  10. Turbine inter-disk cavity cooling air compressor

    DOEpatents

    Chupp, R.E.; Little, D.A.

    1998-01-06

    The inter-disk cavity between turbine rotor disks is used to pressurize cooling air. A plurality of ridges extend radially outwardly over the face of the rotor disks. When the rotor disks are rotated, the ridges cause the inter-disk cavity to compress air coolant flowing through the inter-disk cavity en route to the rotor blades. The ridges eliminate the need for an external compressor to pressurize the air coolant. 5 figs.

  11. Ambient air cooling arrangement having a pre-swirler for gas turbine engine blade cooling

    DOEpatents

    Lee, Ching-Pang; Tham, Kok-Mun; Schroeder, Eric; Meeroff, Jamie; Miller, Jr., Samuel R; Marra, John J

    2015-01-06

    A gas turbine engine including: an ambient-air cooling circuit (10) having a cooling channel (26) disposed in a turbine blade (22) and in fluid communication with a source (12) of ambient air: and an pre-swirler (18), the pre-swirler having: an inner shroud (38); an outer shroud (56); and a plurality of guide vanes (42), each spanning from the inner shroud to the outer shroud. Circumferentially adjacent guide vanes (46, 48) define respective nozzles (44) there between. Forces created by a rotation of the turbine blade motivate ambient air through the cooling circuit. The pre-swirler is configured to impart swirl to ambient air drawn through the nozzles and to direct the swirled ambient air toward a base of the turbine blade. The end walls (50, 54) of the pre-swirler may be contoured.

  12. Turbine inter-disk cavity cooling air compressor

    DOEpatents

    Little, David Allen

    2001-01-01

    A combustion turbine may have a cooling circuit for directing a cooling medium through the combustion turbine to cool various components of the combustion turbine. This cooling circuit may include a compressor, a combustor shell and a component of the combustion turbine to be cooled. This component may be a rotating blade of the combustion turbine. A pressure changing mechanism is disposed in the combustion turbine between the component to be cooled and the combustor shell. The cooling medium preferably flows from the compressor to the combustor shell, through a cooler, the component to the cooled and the pressure changing mechanism. After flowing through the pressure changing mechanism, the cooling medium is returned to the combustor shell. The pressure changing mechanism preferably changes the pressure of the cooling medium from a pressure at which it is exhausted from the component to be cooled to approximately that of the combustor shell.

  13. Cyclic stress analysis of an air-cooled turbine vane

    NASA Technical Reports Server (NTRS)

    Kaufman, A.; Gauntner, D. J.; Gauntner, J. W.

    1975-01-01

    The effects of gas pressure level, coolant temperature, and coolant flow rate on the stress-strain history and life of an air-cooled vane were analyzed using measured and calculated transient metal temperatures and a turbine blade stress analysis program. Predicted failure locations were compared to results from cyclic tests in a static cascade and engine. The results indicate that a high gas pressure was detrimental, a high coolant flow rate somewhat beneficial, and a low coolant temperature the most beneficial to vane life.

  14. Cooling air recycling for gas turbine transition duct end frame and related method

    DOEpatents

    Cromer, Robert Harold; Bechtel, William Theodore; Sutcu, Maz

    2002-01-01

    A method of cooling a transition duct end frame in a gas turbine includes the steps of a) directing cooling air into the end frame from a region external of the transition duct and the impingement cooling sleeve; and b) redirecting the cooling air from the end frame into the annulus between the transition duct and the impingement cooling sleeve.

  15. Cooling circuit for steam and air-cooled turbine nozzle stage

    DOEpatents

    Itzel, Gary Michael; Yu, Yufeng

    2002-01-01

    The turbine vane segment includes inner and outer walls with a vane extending therebetween. The vane includes leading and trailing edge cavities and intermediate cavities. An impingement plate is spaced from the outer wall to impingement-cool the outer wall. Post-impingement cooling air flows through holes in the outer wall to form a thin air-cooling film along the outer wall. Cooling air is supplied an insert sleeve with openings in the leading edge cavity for impingement-cooling the leading edge. Holes through the leading edge afford thin-film cooling about the leading edge. Cooling air is provided the trailing edge cavity and passes through holes in the side walls of the vane for thin-film cooling of the trailing edge. Steam flows through a pair of intermediate cavities for impingement-cooling of the side walls. Post-impingement steam flows to the inner wall for impingement-cooling of the inner wall and returns the post-impingement cooling steam through inserts in other intermediate cavities for impingement-cooling the side walls of the vane.

  16. Cooled snubber structure for turbine blades

    DOEpatents

    Mayer, Clinton A; Campbell, Christian X; Whalley, Andrew; Marra, John J

    2014-04-01

    A turbine blade assembly in a turbine engine. The turbine blade assembly includes a turbine blade and a first snubber structure. The turbine blade includes an internal cooling passage containing cooling air. The first snubber structure extends outwardly from a sidewall of the turbine blade and includes a hollow interior portion that receives cooling air from the internal cooling passage of the turbine blade.

  17. Internally coated air-cooled gas turbine blading

    NASA Technical Reports Server (NTRS)

    Hsu, L.; Stevens, W. G.; Stetson, A. R.

    1979-01-01

    Ten candidate modified nickel-aluminide coatings were developed using the slip pack process. These coatings contain additives such as silicon, chromium and columbium in a nickel-aluminum coating matrix with directionally solidified MAR-M200 + Hf as the substrate alloy. Following a series of screening tests which included strain tolerance, dynamic oxidation and hot corrosion testing, the Ni-19A1-1Cb (nominal composition) coating was selected for application to the internal passages of four first-stage turbine blades. Process development results indicate that a dry pack process is suitable for internal coating application resulting in 18 percent or less reduction in air flow. Coating uniformity, based on coated air-cooled blades, was within + or - 20 percent. Test results show that the presence of additives (silicon, chromium or columbium) appeared to improve significantly the ductility of the NiA1 matrix. However, the environmental resistance of these modified nickel-aluminides were generally inferior to the simple aluminides.

  18. Cold air performance of a 12.766-centimeter-tip-diameter axial-flow cooled turbine. 2: Effect of air ejection on turbine performance

    NASA Technical Reports Server (NTRS)

    Haas, J. E.; Kofskey, M. G.

    1977-01-01

    An air cooled version of a single-stage, axial-flow turbine was investigated to determine aerodynamic performance with and without air ejection from the stator and rotor blades surfaces to simulate the effect of cooling air discharge. Air ejection rate was varied from 0 to 10 percent of turbine mass flow for both the stator and the rotor. A primary-to-air ejection temperature ratio of about 1 was maintained.

  19. The design of an air-cooled metallic high temperature radial turbine

    NASA Technical Reports Server (NTRS)

    Snyder, Philip H.; Roelke, Richard J.

    1988-01-01

    Recent trends in small advanced gas turbine engines call for higher turbine inlet temperatures. Advances in radial turbine technology have opened the way for a cooled metallic radial turbine capable of withstanding turbine inlet temperatures of 2500 F while meeting the challenge of high efficiency in this small flow size range. In response to this need, a small air-cooled radial turbine has been designed utilizing internal blade coolant passages. The coolant flow passage design is uniquely tailored to simultaneously meet rotor cooling needs and rotor fabrication constraints. The rotor flow-path design seeks to realize improved aerodynamic blade loading characteristics and high efficiency while satisfying rotor life requirements. An up-scaled version of the final engine rotor is currently under fabrication and, after instrumentation, will be tested in the warm turbine test facility at the NASA Lewis Research Center.

  20. Analysis of spanwise temperature distribution in three types of air-cooled turbine blade

    NASA Technical Reports Server (NTRS)

    Livingood, John N B; Brown, W Byron

    1950-01-01

    Methods for computing spanwise blade-temperature distributions are derived for air-cooled hollow blades, air-cooled hollow blades with inserts, and air-cooled blades containing internal cooling fins. Individual and combined effects on spanwise blade-temperature distributions of cooling-air and radial heat conduction are determined. In general, the effects of radiation and radial heat conduction were found to be small and the omission of these variations permitted the construction of nondimensional charts for use in determining spanwise temperature distribution through air-cooled turbine blades. An approximate method for determining the allowable stress-limited blade-temperature distribution is included, with brief accounts of a method for determining the maximum allowable effective gas temperatures and the cooling-air requirements. Numerical examples that illustrate the use of the various temperature-distribution equations and of the nondimensional charts are also included.

  1. Performance and economic enhancement of cogeneration gas turbines through compressor inlet air cooling

    NASA Astrophysics Data System (ADS)

    Delucia, M.; Bronconi, R.; Carnevale, E.

    1994-04-01

    Gas turbine air cooling systems serve to raise performance to peak power levels during the hot months when high atmospheric temperatures cause reductions in net power output. This work describes the technical and economic advantages of providing a compressor inlet air cooling system to increase the gas turbine's power rating and reduce its heat rate. The pros and cons of state-of-the-art cooling technologies, i.e., absorption and compression refrigeration, with and without thermal energy storage, were examined in order to select the most suitable cooling solution. Heavy-duty gas turbine cogeneration systems with and without absorption units were modeled, as well as various industrial sectors, i.e., paper and pulp, pharmaceuticals, food processing, textiles, tanning, and building materials. The ambient temperature variations were modeled so the effects of climate could be accounted for in the simulation. The results validated the advantages of gas turbine cogeneration with absorption air cooling as compared to other systems without air cooling.

  2. Preliminary analysis of problem of determining experimental performance of air-cooled turbine II : methods for determining cooling-air-flow characteristics

    NASA Technical Reports Server (NTRS)

    Ellerbrock, Herman H , Jr

    1950-01-01

    In the determination of the performance of an air-cooled turbine, the cooling-air-flow characteristics between the root and the tip of the blades must be evaluated. The methods, which must be verified and the unknown functions evaluated, that are expected to permit the determination of pressure, temperature, and velocity through the blade cooling-air passages from specific investigation are presented.

  3. Effect of Air Cooling of Turbine Disk on Power and Efficiency of Turbine from Turbo Engineering Corporation TT13-18 Turbosupercharger.

    NASA Technical Reports Server (NTRS)

    Berkey, William E.

    1949-01-01

    An investigation was conducted to determine the effect of turbine-disk cooling with air on the efficiency and the power output of the radial-flow turbine from the Turbo Engineering Corporation TT13-18 turbosupercharger. The turbine was operated at a constant range of ratios of turbine-inlet total pressure to turbine-outlet static pressure of 1,5 and 2.0, turbine-inlet total pressure of 30 inches mercury absolute, turbine-inlet total temperature of 12000 to 20000 R, and rotor speeds of 6000 to 22,000 rpm, Over the normal operating range of the turbine, varying the corrected cooling-air weight flow from approximately 0,30 to 0.75 pound per second produced no measurable effect on the corrected turbine shaft horsepower or the turbine shaft adiabatic efficiency. Varying the turbine-inlet total temperature from 12000 to 20000 R caused no measurable change in the corrected cooling-air weight flow. Calculations indicated that the cooling-air pumping power in the disk passages was small and was within the limits of the accuracy of the power measurements. For high turbine power output, the power loss to the compressor for compressing the cooling air was approximately 3 percent of the total turbine shaft horsepower.

  4. Composite casting/bonding construction of an air-cooled, high temperature radial turbine wheel

    NASA Technical Reports Server (NTRS)

    Hammer, A. N.; Aigret, G.; Rodgers, C.; Metcalfe, A. G.

    1983-01-01

    A composite casting/bonding technique has been developed for the fabrication of a unique air-cooled, high temperature radial inflow turbine wheel design applicable to auxilliary power units with small rotor diameters and blade entry heights. The 'split blade' manufacturing procedure employed is an alternative to complex internal ceramic coring. Attention is given to both aerothermodynamic and structural design, of which the latter made advantageous use of the exploration of alternative cooling passage configurations through CAD/CAM system software modification.

  5. Effect of Gas/Steam Turbine Inlet Temperatures on Combined Cycle Having Air Transpiration Cooled Gas Turbine

    NASA Astrophysics Data System (ADS)

    Kumar, S.; Singh, O.

    2012-10-01

    Worldwide efforts are being made for further improving the gas/steam combined cycle performance by having better blade cooling technology in topping cycle and enhanced heat recovery in bottoming cycle. The scope of improvement is possible through turbines having higher turbine inlet temperatures (TITs) of both gas turbine and steam turbine. Literature review shows that a combined cycle with transpiration cooled gas turbine has not been analyzed with varying gas/steam TITs. In view of above the present study has been undertaken for thermodynamic study of gas/steam combined cycle with respect to variation in TIT in both topping and bottoming cycles, for air transpiration cooled gas turbine. The performance of combined cycle with dual pressure heat recovery steam generator has been evaluated for different cycle pressure ratios (CPRs) varying from 11 to 23 and the selection diagrams presented for TIT varying from 1,600 to 1,900 K. Both the cycle efficiency and specific work increase with TIT for each pressure ratio. For each TIT there exists an optimum pressure ratio for cycle efficiency and specific work. For the CPR of 23 the best cycle performance is seen at a TIT of 1,900 K for maximum steam temperature of 570 °C, which gives the cycle efficiency of 60.9 % with net specific work of 909 kJ/kg.

  6. Contingency power for a small turboshaft engine by using water injection into turbine cooling air

    NASA Technical Reports Server (NTRS)

    Biesiadny, Thomas J.; Klann, Gary A.

    1992-01-01

    Because of one-engine-inoperative (OEI) requirements, together with hot-gas reingestion and hot-day, high-altitude take-off situations, power augmentation for multiengine rotorcraft has always been of critical interest. However, power augmentation by using overtemperature at the turbine inlet will shorten turbine life unless a method of limiting thermal and mechanical stress is found. A possible solution involves allowing the turbine inlet temperature to rise to augment power while injecting water into the turbine cooling air to limit hot-section metal temperatures. An experimental water injection device was installed in an engine and successfully tested. Although concern for unprotected subcomponents in the engine hot section prevented demonstration of the technique's maximum potential, it was still possible to demonstrate increases in power while maintaining nearly constant turbine rotor blade temperature.

  7. Contingency power for small turboshaft engines using water injection into turbine cooling air

    NASA Technical Reports Server (NTRS)

    Biesiadny, Thomas J.; Berger, Brett; Klann, Gary A.; Clark, David A.

    1987-01-01

    Because of one engine inoperative requirements, together with hot-gas reingestion and hot day, high altitude takeoff situations, power augmentation for multiengine rotorcraft has always been of critical interest. However, power augmentation using overtemperature at the turbine inlet will shorten turbine life unless a method of limiting thermal and mechanical stresses is found. A possible solution involves allowing the turbine inlet temperature to rise to augment power while injecting water into the turbine cooling air to limit hot-section metal temperatures. An experimental water injection device was installed in an engine and successfully tested. Although concern for unprotected subcomponents in the engine hot section prevented demonstration of the technique's maximum potential, it was still possible to demonstrate increases in power while maintaining nearly constant turbine rotor blade temperature.

  8. Internal coating of air cooled gas turbine blades

    NASA Technical Reports Server (NTRS)

    Ahuja, P. L.

    1979-01-01

    Six coating systems were evaluated for internal coating of decent stage (DS) eutectic high pressure turbine blades. Sequential deposition of electroless Ni by the hydrazine process, slurry Cr, and slurry Al, followed by heat treatment provided the coating composition and thickness for internal coating of DS eutectic turbine blades. Both NiCr and NiCrAl coating compositions were evaluated for strain capability and ductile to brittle transition temperature.

  9. Internal coating of air-cooled gas turbine blades

    NASA Technical Reports Server (NTRS)

    Hsu, L. L.; Stetson, A. R.

    1980-01-01

    Four modified aluminide coatings were developed for IN-792 + Hf alloy using a powder pack method applicable to internal surfaces of air-cooled blades. The coating compositions are Ni-19Al-1Cb, Ni-19Al-3Cb, Ni-17Al-20Cr, and Ni-12Al-20Cr. Cyclic burner rig hot corrosion (900 C) and oxidation (1050 C) tests indicated that Ni-Al-Cb coatings provided better overall resistance than Ni-Al-Cr coatings. Tensile properties of Ni-19Al-1Cb and Ni-12Al-20Cr coated test bars were fully retained at room temperature and 649 C. Stress rupture results exhibited wide scatter around uncoated IN-792 baseline, especially at high stress levels. High cycle fatigue lives of Ni-19Al-1Cb and Ni-12Al-20Cr coated bars (as well as RT-22B coated IN-792) suffered approximately 30 percent decrease at 649 C. Since all test bars were fully heat treated after coating, the effects of coating/processing on IN-792 alloy were not recoverable. Internally coated Ni-19Al-1Cb, Ni-19Al-3Cb, and Ni-12Al-20Cr blades were included in 500-hour endurance engine test and the results were similar to those obtained in burner rig oxidation testing.

  10. Study of Ram-air Heat Exchangers for Reducing Turbine Cooling-air Temperature of a Supersonic Aircraft Turbojet Engine

    NASA Technical Reports Server (NTRS)

    Diaguila, Anthony J; Livingood, John N B; Eckert, Ernst R G

    1956-01-01

    The sizes and weights of the cores of heat exchangers were determined analytically for possible application for reducing turbine cooling-air temperatures of an engine designed for a Mach number of 2.5 and an altitude The sizes and weights of the cores of heat exchangers were determined analytically for possible application for reducing turbine cooling-air temperatures of an engine designed for a Mach number of 2.5 and an altitude of 70,000 feet. A compressor-bleed-air weight flow of 2.7 pounds per second was assumed for the coolant; ram air was considered as the other fluid. Pressure drops and inlet states of both fluids were prescribed, and ranges of compressor-bleed-air temperature reductions and of the ratio of compressor-bleed to ram-air weight flows were considered.

  11. Summary of research and development effort on air and water cooling of gas turbine blades

    SciTech Connect

    Fraas, A.P.

    1980-03-01

    The review on air- and water-cooled gas turbines from the 1904 Lemale-Armengaud water-cooled gas turbine, the 1948 to 1952 NACA work, and the program at GE indicates that the potential of air cooling has been largely exploited in reaching temperatures of 1100/sup 0/C (approx. 2000/sup 0/F) in utility service and that further increases in turbine inlet temperature may be obtained with water cooling. The local heat flux in the first-stage turbine rotor with water cooling is very high, yielding high-temperature gradients and severe thermal stresses. Analyses and tests indicate that by employing a blade with an outer cladding of an approx. 1-mm-thick oxidation-resistant high-nickel alloy, a sublayer of a high-thermal-conductivity, high-strength, copper alloy containing closely spaced cooling passages approx. 2 mm in ID to minimize thermal gradients, and a central high-strength alloy structural spar, it appears possible to operate a water-cooled gas turbine with an inlet gas temperature of 1370/sup 0/C. The cooling-water passages must be lined with an iron-chrome-nickel alloy must be bent 90/sup 0/ to extend in a neatly spaced array through the platform at the base of the blade. The complex geometry of the blade design presents truly formidable fabrication problems. The water flow rate to each of many thousands of coolant passages must be metered and held to within rather close limits because the heat flux is so high that a local flow interruption of only a few seconds would lead to a serious failure.Heat losses to the cooling water will run approx. 10% of the heat from the fuel. By recoverying this waste heat for feedwater heating in a command cycle, these heat losses will give a degradation in the power plant output of approx. 5% relative to what might be obtained if no cooling were required. However, the associated power loss is less than half that to be expected with an elegant air cooling system.

  12. Benefits of compressor inlet air cooling for gas turbine cogeneration plants

    SciTech Connect

    De Lucia, M.; Lanfranchi, C.; Boggio, V.

    1996-07-01

    Compressor inlet air cooling is an effective method for enhancing the performance of gas turbine plants. This paper presents a comparative analysis of different solutions for cooling the compressor inlet air for the LM6000 gas turbine in a cogeneration plant operated in base load. Absorption and evaporative cooling systems are considered and their performance and economic benefits compared for the dry low-NO{sub x} LM6000 version. Reference is made to two sites in Northern and Southern Italy, whose climate data series for modeling the variations in ambient temperature during the single day were used to account for the effects of climate in the simulation. The results confirmed the advantages of inlet air cooling systems. In particular, evaporative cooling proved to be cost effective, though capable of supplying only moderate cooling, while absorption systems have a higher cost but are also more versatile and powerful in base-load operation. An integration of the two systems proved to be able to give both maximum performance enhancement and net economic benefit.

  13. Engine investigation of an air-cooled turbine rotor blade incorporating impingement-cooled leading edge, chordwise passages, and a slotted trailing edge

    NASA Technical Reports Server (NTRS)

    Dengler, R. P.; Yeh, F. C.; Gauntner, J. W.; Fallon, G. E.

    1972-01-01

    Experimental temperatures are presented for an air-cooled turbine rotor blade tested in an engine. The data were obtained for turbine stator inlet temperatures from 2000 to 2500 F and for turbine-inlet gas pressures from 32 to 46 psia. Average and local blade heat-transfer data are correlated. Potential allowable increases in gas temperature are also discussed.

  14. JT8D revised high-pressure turbine cooling and other outer air seal program

    NASA Technical Reports Server (NTRS)

    Gaffin, W. O.

    1979-01-01

    The JT8D high pressure turbine was revised to reduce leakage between the blade tip shrouds and the outer air seal, and engine testing was performed to determine the effect on performance. The addition of a second knife-edge on the blade tip shroud, the extension of the honeycomb seal land to cover the added knife-edge and an existing spoiler on the shroud, and a material substitution in the seal support ring to improve thermal growth characteristics are included. A relocation of the blade cooling air discharge to insure adequate cooling flow is required. Significant specific fuel consumption and exhaust gas temperature improvements were demonstrated with the revised turbine in sea level and simulated altitude engine tests. Inspection of the revised seal hardware after these tests showed no unusual wear or degradation.

  15. Flow measurement in base cooling air passages of a rotating turbine blade

    NASA Technical Reports Server (NTRS)

    Liebert, C. H.; Pollack, F. G.

    1974-01-01

    The operational performance is decribed of a shaft-mounted system for measuring the air mass flow rate in the base cooling passages of a rotating turbine blade. Shaft speeds of 0 to 9000 rpm, air mass flow rates of 0.0035 to 0.039 kg/sec (0.0077 to 0.085 lbm/sec), and blade air temperatures of 300 to 385 K (80 to 233 F) were measured. Comparisons of individual rotating blade flows and corresponding stationary supply orifice flows agreed to within 10 percent.

  16. Review and status of heat-transfer technology for internal passages of air-cooled turbine blades

    NASA Technical Reports Server (NTRS)

    Yeh, F. C.; Stepka, F. S.

    1984-01-01

    Selected literature on heat-transfer and pressure losses for airflow through passages for several cooling methods generally applicable to gas turbine blades is reviewed. Some useful correlating equations are highlighted. The status of turbine-blade internal air-cooling technology for both nonrotating and rotating blades is discussed and the areas where further research is needed are indicated. The cooling methods considered include convection cooling in passages, impingement cooling at the leading edge and at the midchord, and convection cooling in passages, augmented by pin fins and the use of roughened internal walls.

  17. Structural Design and Preliminary Evaluation of a Lightweight, Brazed, Air-Cooled Turbine Rotor Assembly

    NASA Technical Reports Server (NTRS)

    Meyer, Andre J., Jr.; Morgan, William C.

    1958-01-01

    A lightweight turbine rotor assembly was devised, and components were evaluated in a full-scale jet engine. Thin sheet-metal airfoils were brazed to radial fingers that were an integral part of a number of thin disks composing the turbine rotor. Passages were provided between the disks and in the blades for air cooling. The computed weight of the assembly was 50 percent less than that of a similar turbine of normal construction used in a conventional turbojet engine. Two configurations of sheet-metal test blades simulating the manner of attachment were fabricated and tested in a turbojet engine at rated speed and temperature. After 8-1/2 hours of operation pieces broke loose from the tip sections of the better blades. Severe cracking produced by vibration was determined as the cause of failure. Several methods of overcoming the vibration problem are suggested.

  18. Computing Cooling Flows in Turbines

    NASA Technical Reports Server (NTRS)

    Gauntner, J.

    1986-01-01

    Algorithm developed for calculating both quantity of compressor bleed flow required to cool turbine and resulting decrease in efficiency due to cooling air injected into gas stream. Program intended for use with axial-flow, air-breathing, jet-propulsion engines with variety of airfoil-cooling configurations. Algorithm results compared extremely well with figures given by major engine manufacturers for given bulk-metal temperatures and cooling configurations. Program written in FORTRAN IV for batch execution.

  19. Thermal and flow analysis of a convection air-cooled ceramic coated porous metal concept for turbine vanes

    NASA Technical Reports Server (NTRS)

    Stepka, F. S.

    1981-01-01

    The heat transfer and pressure drop through turbine vanes made of a sintered, porous metal coated with a thin layer of ceramic and convection cooled by spanwise flow of cooling air were analyzed. The analysis was made to determine the feasibility of using this concept for cooling very small turbines, primarily for short duration applications such as in missile engines. The analysis was made for gas conditions of approximately 10 and 40 atm and 1644 K and with turbine vanes made of felt type porous metals with relative densities from 0.2 to 0.6 and ceramic coating thicknesses of 0.076 to 0.254 mm.

  20. Analysis of Coolant-flow Requirements for an Improved, Internal-strut-supported, Air-cooled Turbine-rotor Blade

    NASA Technical Reports Server (NTRS)

    Schramm, Wilson B; Nachtigall, Alfred J

    1952-01-01

    An analytical evaluation of a new typ An analytical evaluation of a new type of air-cooled turbine-rotor-blade design, based on the principle of submerging the load-carrying element in cooling air within a thin high-temperature sheel, indicates that this principle of blade design permits the load carrying element to be operated at considerably lower temperature than that of the enveloping shell. Comparison with an air-cooled shell-supported air-cooled blade has greater potentiality to withstand increased stresses that can be anticipated in future engines.

  1. Effects of a ceramic coating on metal temperatures of an air-cooled turbine vane

    NASA Astrophysics Data System (ADS)

    Gladden, H. J.; Liebert, C. H.

    1980-02-01

    The metal temperatures of air cooled turbine vanes both uncoated and coated with the NASA thermal barrier system were studied experimentally. Current and advanced gas turbine engine conditions were simulated at reduced temperatures and pressures. Airfoil metal temperatures were significantly reduced, both locally and on the average, by use of the the coating. However, at low gas Reynolds number, the ceramic coating tripped a laminar boundary layer on the suction surface, and the resulting higher heat flux increased the metal temperatures. Simulated coating loss was also investigated and shown to increase local metal temperatures. However, the metal temperatures in the leading edge region remained below those of the uncoated vane tested at similar conditions. Metal temperatures in the trailing edge region exceeded those of the uncoated vane.

  2. Effects of a ceramic coating on metal temperatures of an air-cooled turbine vane

    NASA Technical Reports Server (NTRS)

    Gladden, H. J.; Liebert, C. H.

    1980-01-01

    The metal temperatures of air cooled turbine vanes both uncoated and coated with the NASA thermal barrier system were studied experimentally. Current and advanced gas turbine engine conditions were simulated at reduced temperatures and pressures. Airfoil metal temperatures were significantly reduced, both locally and on the average, by use of the the coating. However, at low gas Reynolds number, the ceramic coating tripped a laminar boundary layer on the suction surface, and the resulting higher heat flux increased the metal temperatures. Simulated coating loss was also investigated and shown to increase local metal temperatures. However, the metal temperatures in the leading edge region remained below those of the uncoated vane tested at similar conditions. Metal temperatures in the trailing edge region exceeded those of the uncoated vane.

  3. Radial turbine cooling

    NASA Astrophysics Data System (ADS)

    Roelke, Richard J.

    The technology of high temperature cooled radial turbines is reviewed. Aerodynamic performance considerations are described. Heat transfer and structural analysis are addressed, and in doing so the following topics are covered: cooling considerations, hot side convection, coolant side convection, and rotor mechanical analysis. Cooled rotor concepts and fabrication are described, and the following are covered in this context: internally cooled rotor, hot isostatic pressure bonded rotor, laminated rotor, split blade rotor, and the NASA radial turbine program.

  4. Radial turbine cooling

    NASA Technical Reports Server (NTRS)

    Roelke, Richard J.

    1992-01-01

    The technology of high temperature cooled radial turbines is reviewed. Aerodynamic performance considerations are described. Heat transfer and structural analysis are addressed, and in doing so the following topics are covered: cooling considerations, hot side convection, coolant side convection, and rotor mechanical analysis. Cooled rotor concepts and fabrication are described, and the following are covered in this context: internally cooled rotor, hot isostatic pressure bonded rotor, laminated rotor, split blade rotor, and the NASA radial turbine program.

  5. Heat transfer technology for internal passages of air-cooled blades for heavy-duty gas turbines.

    PubMed

    Weigand, B; Semmler, K; von Wolfersdorf, J

    2001-05-01

    The present review paper, although far from being complete, aims to give an overview about the present state of the art in the field of heat transfer technology for internal cooling of gas turbine blades. After showing some typical modern cooled blades, the different methods to enhance heat transfer in the internal passages of air-cooled blades are discussed. The complicated flows occurring in bends are described in detail, because of their increasing importance for modern cooling designs. A short review about testing of cooling design elements is given, showing the interaction of the different cooling features as well. The special focus of the present review has been put on the cooling of blades for heavy-duty gas turbines, which show several differences compared to aero-engine blades. PMID:11460627

  6. Preliminary analysis of problem of determining experimental performance of air-cooled turbine III : methods for determining power and efficiency

    NASA Technical Reports Server (NTRS)

    Ellerbrock, Herman H , Jr; Ziemer, Robert R

    1950-01-01

    Suggested formula are given for determining air-cooled turbine-performance characteristics, such as power and efficiency, as functions of certain parameters. These functions, generally being unknown, are determined from experimental data obtained from specific investigations. Special plotting methods for isolating the effect of each parameter are outlined.

  7. Durability of zirconia thermal-barrier ceramic coatings on air-cooled turbine blades in cyclic jet engine operation

    NASA Technical Reports Server (NTRS)

    Liebert, C. H.; Jacobs, R. E.; Stecura, S.; Morse, C. R.

    1976-01-01

    Thermal barrier ceramic coatings of stabilized zirconia over a bond coat of Ni Cr Al Y were tested for durability on air cooled turbine rotor blades in a research turbojet engine. Zirconia stabilized with either yttria, magnesia, or calcia was investigated. On the basis of durability and processing cost, the yttria stabilized zirconia was considered the best of the three coatings investigated.

  8. Performance comparison between transpiration air cooled turbine 3000 F (1649 C) stator vanes and solid uncooled vanes

    NASA Astrophysics Data System (ADS)

    Manning, G. B.; Moskowitz, S.; Cole, R.

    1984-06-01

    Testing was conducted to compare the aerodynamic performance of a turbine vane using transpiration air-cooling capable of operation at 3000 F (1649 C) gas temperature with a vane of identical profile with no cooling provisions to determine the effect of cooling on vane kinetic energy efficiency and loss coefficient. The test configuration was a 10-vane section of full scale first stagae turbien stator annulus designed for 1.6 pressure ratio, cooling air flow equal to 6.1 percent of primary flow, 3000 F (1649 C) turbine inlet temperature and primary-to-coolant temperature ratio of 2.7. To enable comparison with other investigations, tests were conducted at three pressure ratios from 1.4 to 1.6, three coolant flows from 75 to 120 percent of design, and three primary-to-coolant temperature ratios from 2.70 to 1.15. Efficiency, loss coefficent and flow capacity test results were in good agreement with predicted values for both the transpiration air cooled and uncooled vanes. The testing demonstrated that it is necessary to conduct test evaluations of transpiration air-cooled components at or near design coolant-to-gas stream temperature ratio in order to achieve correct results.

  9. Experimental Investigation of Air-Cooled Turbine Blades in Turbojet Engine. 7: Rotor-Blade Fabrication Procedures

    NASA Technical Reports Server (NTRS)

    Long, Roger A.; Esgar, Jack B.

    1951-01-01

    An experimental investigation was conducted to determine the cooling effectiveness of a wide variety of air-cooled turbine-blade configurations. The blades, which were tested in the turbine of a - commercial turbojet engine that was modified for this investigation by replacing two of the original blades with air-cooled blades located diametrically opposite each other, are untwisted, have no aerodynamic taper, and have essentially the same external profile. The cooling-passage configuration is different for each blade, however. The fabrication procedures were varied and often unique. The blades were fabricated using methods most suitable for obtaining a small number of blades for use in the cooling investigations and therefore not all the fabrication procedures would be directly applicable to production processes, although some of the ideas and steps might be useful. Blade shells were obtained by both casting and forming. The cast shells were either welded to the blade base or cast integrally with the base. The formed shells were attached to the base by a brazing and two welding methods. Additional surface area was supplied in the coolant passages by the addition of fins or tubes that were S-brazed. to the shell. A number of blades with special leading- and trailing-edge designs that provided added cooling to these areas were fabricated. The cooling effectiveness and purposes of the various blade configurations are discussed briefly.

  10. A method for measuring cooling air flow in base coolant passages of rotating turbine blades

    NASA Technical Reports Server (NTRS)

    Liebert, C. H.; Pollack, F. G.

    1975-01-01

    Method accurately determines actual coolant mass flow rate in cooling passages of rotating turbine blades. Total and static pressures are measured in blade base coolant passages. Mass flow rates are calculated from these measurements of pressure, measured temperature and known area.

  11. Radial turbine cooling

    NASA Technical Reports Server (NTRS)

    Roelke, Richard J.

    1992-01-01

    Radial turbines have been used extensively in many applications including small ground based electrical power generators, automotive engine turbochargers and aircraft auxiliary power units. In all of these applications the turbine inlet temperature is limited to a value commensurate with the material strength limitations and life requirements of uncooled metal rotors. To take advantage of all the benefits that higher temperatures offer, such as increased turbine specific power output or higher cycle thermal efficiency, requires improved high temperature materials and/or blade cooling. Extensive research is on-going to advance the material properties of high temperature superalloys as well as composite materials including ceramics. The use of ceramics with their high temperature potential and low cost is particularly appealing for radial turbines. However until these programs reach fruition the only way to make significant step increases beyond the present material temperature barriers is to cool the radial blading.

  12. Turbine airfoil with ambient cooling system

    DOEpatents

    Campbell, Jr, Christian X.; Marra, John J.; Marsh, Jan H.

    2016-06-07

    A turbine airfoil usable in a turbine engine and having at least one ambient air cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels configured to receive ambient air at about atmospheric pressure. The ambient air cooling system may have a tip static pressure to ambient pressure ratio of at least 0.5, and in at least one embodiment, may include a tip static pressure to ambient pressure ratio of between about 0.5 and about 3.0. The cooling system may also be configured such that an under root slot chamber in the root is large to minimize supply air velocity. One or more cooling channels of the ambient air cooling system may terminate at an outlet at the tip such that the outlet is aligned with inner surfaces forming the at least one cooling channel in the airfoil to facilitate high mass flow.

  13. Conjugate heat transfer investigation on the cooling performance of air cooled turbine blade with thermal barrier coating

    NASA Astrophysics Data System (ADS)

    Ji, Yongbin; Ma, Chao; Ge, Bing; Zang, Shusheng

    2016-08-01

    A hot wind tunnel of annular cascade test rig is established for measuring temperature distribution on a real gas turbine blade surface with infrared camera. Besides, conjugate heat transfer numerical simulation is performed to obtain cooling efficiency distribution on both blade substrate surface and coating surface for comparison. The effect of thermal barrier coating on the overall cooling performance for blades is compared under varied mass flow rate of coolant, and spatial difference is also discussed. Results indicate that the cooling efficiency in the leading edge and trailing edge areas of the blade is the lowest. The cooling performance is not only influenced by the internal cooling structures layout inside the blade but also by the flow condition of the mainstream in the external cascade path. Thermal barrier effects of the coating vary at different regions of the blade surface, where higher internal cooling performance exists, more effective the thermal barrier will be, which means the thermal protection effect of coatings is remarkable in these regions. At the designed mass flow ratio condition, the cooling efficiency on the pressure side varies by 0.13 for the coating surface and substrate surface, while this value is 0.09 on the suction side.

  14. Turbine blade cooling

    DOEpatents

    Staub, Fred Wolf; Willett, Fred Thomas

    1999-07-20

    A turbine rotor blade comprises a shank portion, a tip portion and an airfoil. The airfoil has a pressure side wall and a suction side wall that are interconnected by a plurality of partition sidewalls, defining an internal cooling passageway within the airfoil. The internal cooling passageway includes at least one radial outflow passageway to direct a cooling medium flow from the shank portion towards the tip portion and at least one radial inflow passageway to direct a cooling medium flow from the tip portion towards the shank portion. A number of mixing ribs are disposed on the partition sidewalls within the radial outflow passageways so as to enhance the thermal mixing of the cooling medium flow, thereby producing improved heat transfer over a broad range of the Buoyancy number.

  15. Turbine blade cooling

    DOEpatents

    Staub, Fred Wolf; Willett, Fred Thomas

    2000-01-01

    A turbine rotor blade comprises a shank portion, a tip portion and an airfoil. The airfoil has a pressure side wall and a suction side wall that are interconnected by a plurality of partition sidewalls, defining an internal cooling passageway within the airfoil. The internal cooling passageway includes at least one radial outflow passageway to direct a cooling medium flow from the shank portion towards the tip portion and at least one radial inflow passageway to direct a cooling medium flow from the tip portion towards the shank portion. A number of mixing ribs are disposed on the partition sidewalls within the radial outflow passageways so as to enhance the thermal mixing of the cooling medium flow, thereby producing improved heat transfer over a broad range of the Buoyancy number.

  16. Turbine blade cooling

    SciTech Connect

    Staub, F.W.; Willett, F.T.

    1999-07-20

    A turbine rotor blade comprises a shank portion, a tip portion and an airfoil. The airfoil has a pressure side wall and a suction side wall that are interconnected by a plurality of partition sidewalls, defining an internal cooling passageway within the airfoil. The internal cooling passageway includes at least one radial outflow passageway to direct a cooling medium flow from the shank portion towards the tip portion and at least one radial inflow passageway to direct a cooling medium flow from the tip portion towards the shank portion. A number of mixing ribs are disposed on the partition sidewalls within the radial outflow passageways so as to enhance the thermal mixing of the cooling medium flow, thereby producing improved heat transfer over a broad range of the Buoyancy number. 13 figs.

  17. TURBINE COOLING FLOW AND THE RESULTING DECREASE IN TURBINE EFFICIENCY

    NASA Technical Reports Server (NTRS)

    Gauntner, J. W.

    1994-01-01

    This algorithm has been developed for calculating both the quantity of compressor bleed flow required to cool a turbine and the resulting decrease in efficiency due to cooling air injected into the gas stream. Because of the trend toward higher turbine inlet temperatures, it is important to accurately predict the required cooling flow. This program is intended for use with axial flow, air-breathing jet propulsion engines with a variety of airfoil cooling configurations. The algorithm results have compared extremely well with figures given by major engine manufacturers for given bulk metal temperatures and cooling configurations. The program calculates the required cooling flow and corresponding decrease in stage efficiency for each row of airfoils throughout the turbine. These values are combined with the thermodynamic efficiency of the uncooled turbine to predict the total bleed airflow required and the altered turbine efficiency. There are ten airfoil cooling configurations and the algorithm allows a different option for each row of cooled airfoils. Materials technology is incorporated and requires the date of the first year of service for the turbine stator vane and rotor blade. The user must specify pressure, temperatures, and gas flows into the turbine. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 3080 series computer with a central memory requirement of approximately 61K of 8 bit bytes. This program was developed in 1980.

  18. Burner rig study of variables involved in hole plugging of air cooled turbine engine vanes

    NASA Technical Reports Server (NTRS)

    Deadmore, D. L.; Lowell, C. E.

    1983-01-01

    The effects of combustion gas composition, flame temperatures, and cooling air mass flow on the plugging of film cooling holes by a Ca-Fe-P-containing deposit were investigated. The testing was performed on film-cooled vanes exposed to the combustion gases of an atmospheric Mach 0.3 burner rig. The extent of plugging was determined by measurement of the open hole area at the conclusion of the tests as well as continuous monitoring of some of the tests using stop-action photography. In general, as the P content increased, plugging rates also increased. The plugging was reduced by increasing flame temperature and cooling air mass flow rates. At times up to approximately 2 hours little plugging was observed. This apparent incubation period was followed by rapid plugging, reaching in several hours a maximum closure whose value depended on the conditions of the test.

  19. Cooling of Gas Turbines I - Effects of Addition of Fins to Blade Tips and Rotor, Admission of Cooling Air Through Part of Nozzles, and Change in Thermal Conductivity of Turbine Components

    NASA Technical Reports Server (NTRS)

    Brown, Byron

    1947-01-01

    An analysis was developed for calculating the radial temperature distribution in a gas turbine with only the temperatures of the gas and the cooling air and the surface heat-transfer coefficient known. This analysis was applied to determine the temperatures of a complete wheel of a conventional single-stage impulse exhaust-gas turbine. The temperatures were first calculated for the case of the turbine operating at design conditions of speed, gas flow, etc. and with only the customary cooling arising from exposure of the outer blade flange and one face of the rotor to the air. Calculations were next made for the case of fins applied to the outer blade flange and the rotor. Finally the effects of using part of the nozzles (from 0 to 40 percent) for supplying cooling air and the effects of varying the metal thermal conductivity from 12 to 260 Btu per hour per foot per degree Farenheit on the wheel temperatures were determined. The gas temperatures at the nozzle box used in the calculations ranged from 1600F to 2000F. The results showed that if more than a few hundred degrees of cooling of turbine blades are required other means than indirect cooling with fins on the rotor and outer blade flange would be necessary. The amount of cooling indicated for the type of finning used could produce some improvement in efficiency and a large increase in durability of the wheel. The results also showed that if a large difference is to exist between the effective temperature of the exhaust gas and that of the blade material, as must be the case with present turbine materials and the high exhaust-gas temperatures desired (2000F and above), two alternatives are suggested: (a) If metal with a thermal conductivity comparable with copper is used, then the blade temperature can be reduced by strong cooling at both the blade tip and root. The center of the blade will be less than 2000F hotter than the ends; (b) With low conductivity materials some method of direct cooling other than

  20. Development of gas-pressure bonding process for air-cooled turbine blades

    NASA Technical Reports Server (NTRS)

    Meiners, K. E.

    1972-01-01

    An investigation was conducted on the application of gas-pressure bonding to the joining of components for convectively cooled turbine blades and vanes. A processing procedure was established for joining the fins of Udimet 700 and TD NiCr sheet metal airfoil shells to cast B1900 struts without the use of internal support tooling. Alternative methods employing support tooling were investigated. Testing procedures were developed and employed to determine shear strengths and internal burst pressures of flat and cylindrical bonded finned shell configurations at room temperature and 1750 F. Strength values were determined parallel and transverse to the cooling fin direction. The effect of thermal cycles from 1750 F to room temperature on strength was also investigated.

  1. Turbine airfoil film cooling

    NASA Astrophysics Data System (ADS)

    Hylton, Larry D.

    1986-10-01

    Emphasis is placed on developing more accurate analytical models for predicting turbine airfoil external heat transfer rates. Performance goals of new engines require highly refined, accurate design tools to meet durability requirements. In order to obtain improvements in analytical capabilities, programs are required which focus on enhancing analytical techniques through verification of new models by comparison with relevant experimental data. The objectives of the current program are to develop an analytical approach, based on boundary layer theory, for predicting the effects of airfoil film cooling on downstream heat transfer rates and to verify the resulting analytical method by comparison of predictions with hot cascade data obtained under this program.

  2. Measured and calculated wall temperatures on air-cooled turbine vanes with boundary layer transition

    NASA Astrophysics Data System (ADS)

    Liebert, C. H.; Gaugler, R. E.; Gladden, H. J.

    1982-11-01

    Convection cooled turbine vane metal wall temperatures experimentally obtained in a hot cascade for one vane design were compared with wall temperatures calculated with TACT1 and STAN5 computer codes which incorporated various models for predicting laminar-to-turbulent boundary layer transition. Favorable comparisons on both vane surface were obtained at high Reynolds number with only one of these transition models. When other models were used, temperature differences between calculated and experimental data obtained at the high Reynolds number were as much as 14 percent in the separation bubble region of the pressure surface. On the suction surface and at lower Reynolds number, predictions and data unsatisfactorily differed by as much as 22 percent. Temperature differences of this magnitude can represent orders of magnitude error in blade life prediction.

  3. Heat pipe turbine vane cooling

    SciTech Connect

    Langston, L.; Faghri, A.

    1995-10-01

    The applicability of using heat pipe principles to cool gas turbine vanes is addressed in this beginning program. This innovative concept involves fitting out the vane interior as a heat pipe and extending the vane into an adjacent heat sink, thus transferring the vane incident heat transfer through the heat pipe to heat sink. This design provides an extremely high heat transfer rate and an uniform temperature along the vane due to the internal change of phase of the heat pipe working fluid. Furthermore, this technology can also eliminate hot spots at the vane leading and trailing edges and increase the vane life by preventing thermal fatigue cracking. There is also the possibility of requiring no bleed air from the compressor, and therefore eliminating engine performance losses resulting from the diversion of compressor discharge air. Significant improvement in gas turbine performance can be achieved by using heat pipe technology in place of conventional air cooled vanes. A detailed numerical analysis of a heat pipe vane will be made and an experimental model will be designed in the first year of this new program.

  4. Heat pipe turbine vane cooling

    SciTech Connect

    Langston, L.; Faghri, A.

    1995-12-31

    The applicability of using heat pipe principles to cool gas turbine vanes is addressed in this beginning program. This innovative concept involves fitting out the vane interior as a heat pipe and extending the vane into an adjacent heat sink, thus transferring the vane incident heat transfer through the heat pipe to heat sink. This design provides an extremely high heat transfer rate and a uniform temperature along the vane due to the internal change of phase of the heat pipe working fluid. Furthermore, this technology can also eliminate hot spots at the vane leading and trailing edges and increase the vane life by preventing thermal fatigue cracking. There is also the possibility of requiring no bleed air from the compressor, and therefore eliminating engine performance losses resulting from the diversion of compressor discharge air. Significant improvement in gas turbine performance can be achieved by using heat pipe technology in place of conventional air cooled vanes. A detailed numerical analysis of a heat pipe vane will be made and an experimental model will be designed in the first year of this new program.

  5. Cooled blades of gas turbines /Thermal design and profiling/

    NASA Astrophysics Data System (ADS)

    Kopelev, S. Z.

    The efficiency of the air-cooling of gas turbine blades is analyzed, and various approaches to the design of air-cooled gas turbine blades are discussed. In particular, attention is given to the analysis of heat transfer in blades with an internal deflector, blades with radial air flow, and blades with convective-barrier cooling. Methods for calculating the temperature of blades with transverse flow of the cooling air are discussed, as are methods for calculating losses in an air-cooled turbine.

  6. Passively cooled direct drive wind turbine

    DOEpatents

    Costin, Daniel P.

    2008-03-18

    A wind turbine is provided that passively cools an electrical generator. The wind turbine includes a plurality of fins arranged peripherally around a generator house. Each of the fins being oriented at an angle greater than zero degrees to allow parallel flow of air over the fin. The fin is further tapered to allow a constant portion of the fin to extend beyond the air stream boundary layer. Turbulence initiators on the nose cone further enhance heat transfer at the fins.

  7. Distortion Behavior of a Heavy Hydro Turbine Blade Casting During Forced Air Cooling in Normalizing Treatment Process

    NASA Astrophysics Data System (ADS)

    Yu, Hai-Liang; Kang, Jin-Wu; Wang, Tian-Jiao; Ma, Ji-Yu; Hu, Yong-Yi; Huang, Tian-You; Wang, Shi-Bin; Wu, Ying; Zhang, Cheng-Chun; Dai, Yan-Tao; Li, Peng

    2012-01-01

    Distortion behavior of blade castings in heat treatment process determines their geometrical accuracy, and improper control of it may result in additional repair, shape righting, or even rejection. This article presents a novel approach for discovering the distortion behavior of heavy blade castings during heat treatment process in production. Real-time measurements of distortion and temperature field of a heavy hydro turbine blade casting weighted 17 ton during forced air cooling in normalizing treatment process were carried out by using deformation measurement instruments and an infrared thermal imaging camera. The distortion processes of the typical locations of blade and the temperature field of the blade were obtained. One corner on the blade outlet edge side exhibits variation of distortion with two peaks and a valley. The range reaches 97 mm and the final distortion value is 76 mm. The maximum temperature difference on blade surface reaches 460 °C after 80 min of cooling. Influences of thermal stress and phase transformation stress on the distortion of the blade were elucidated and discussed. The results are of great significance for the understanding and control of the distortion behavior of hydro turbine blades in heat treatment.

  8. Sequential cooling insert for turbine stator vane

    SciTech Connect

    Jones, Russel B; Krueger, Judson J; Plank, William L

    2014-11-04

    A sequential impingement cooling insert for a turbine stator vane that forms a double impingement for the pressure and suction sides of the vane or a triple impingement. The insert is formed from a sheet metal formed in a zigzag shape that forms a series of alternating impingement cooling channels with return air channels, where pressure side and suction side impingement cooling plates are secured over the zigzag shaped main piece. Another embodiment includes the insert formed from one or two blocks of material in which the impingement channels and return air channels are machined into each block.

  9. Sequential cooling insert for turbine stator vane

    SciTech Connect

    Jones, Russell B; Krueger, Judson J; Plank, William L

    2014-04-01

    A sequential impingement cooling insert for a turbine stator vane that forms a double impingement for the pressure and suction sides of the vane or a triple impingement. The insert is formed from a sheet metal formed in a zigzag shape that forms a series of alternating impingement cooling channels with return air channels, where pressure side and suction side impingement cooling plates are secured over the zigzag shaped main piece. Another embodiment includes the insert formed from one or two blocks of material in which the impingement channels and return air channels are machined into each block.

  10. Serial cooling of a combustor for a gas turbine engine

    DOEpatents

    Abreu, Mario E.; Kielczyk, Janusz J.

    2001-01-01

    A combustor for a gas turbine engine uses compressed air to cool a combustor liner and uses at least a portion of the same compressed air for combustion air. A flow diverting mechanism regulates compressed air flow entering a combustion air plenum feeding combustion air to a plurality of fuel nozzles. The flow diverting mechanism adjusts combustion air according to engine loading.

  11. Stress analysis study in cooled radial inflow turbine

    NASA Technical Reports Server (NTRS)

    Hamed, A.; Sheoran, Y.; Tabakoff, W.

    1978-01-01

    With increased turbine inlet temperatures, numerical methods of thermal and stress analysis are becoming more valuable in the design of air-cooled turbines. This paper presents a study of the stresses associated with different cooling patterns in a radial inflow turbine rotor. The finite element method is used in the stress calculations taking into consideration centrifugal, thermal and aerodynamic loading. The effects of temperature distribution and the presence of internal cooling passages are discussed.

  12. Algorithm for calculating turbine cooling flow and the resulting decrease in turbine efficiency

    NASA Technical Reports Server (NTRS)

    Gauntner, J. W.

    1980-01-01

    An algorithm is presented for calculating both the quantity of compressor bleed flow required to cool the turbine and the decrease in turbine efficiency caused by the injection of cooling air into the gas stream. The algorithm, which is intended for an axial flow, air routine in a properly written thermodynamic cycle code. Ten different cooling configurations are available for each row of cooled airfoils in the turbine. Results from the algorithm are substantiated by comparison with flows predicted by major engine manufacturers for given bulk metal temperatures and given cooling configurations. A list of definitions for the terms in the subroutine is presented.

  13. Preliminary analysis of problem of determining experimental performance of air-cooled turbine I : methods for determining heat-transfer characteristics

    NASA Technical Reports Server (NTRS)

    Ellerbrock, Herman H , Jr; Ziemer, Robert R

    1950-01-01

    In determining the experimental performance of an air-cooled turbine, the heat-transfer characteristics must be evaluated. The suggested formulas that are required to determine these characteristics are presented. The formulas have a form in which dependent parameters are expressed as unknown functions of independent parameters. Methods of experimenting to determine these functions are suggested. In some cases general heat-transfer discussions that lead to the suggested forms of the formulas are given.

  14. Comparison of Calculated and Experimental Temperatures and Coolant Pressure Losses for a Cascade of Small Air-Cooled Turbine Rotor Blades

    NASA Technical Reports Server (NTRS)

    Stepka, Francis S

    1958-01-01

    Average spanwise blade temperatures and cooling-air pressure losses through a small (1.4-in, span, 0.7-in, chord) air-cooled turbine blade were calculated and are compared with experimental nonrotating cascade data. Two methods of calculating the blade spanwise metal temperature distributions are presented. The method which considered the effect of the length-to-diameter ratio of the coolant passage on the blade-to-coolant heat-transfer coefficient and assumed constant coolant properties based on the coolant bulk temperature gave the best agreement with experimental data. The agreement obtained was within 3 percent at the midspan and tip regions of the blade. At the root region of the blade, the agreement was within 3 percent for coolant flows within the turbulent flow regime and within 10 percent for coolant flows in the laminar regime. The calculated and measured cooling-air pressure losses through the blade agreed within 5 percent. Calculated spanwise blade temperatures for assumed turboprop engine operating conditions of 2000 F turbine-inlet gas temperature and flight conditions of 300 knots at a 30,000-foot altitude agreed well with those obtained by the extrapolation of correlated experimental data of a static cascade investigation of these blades.

  15. Investigations of Air-Cooled Turbine Rotors for Turbojet Engines. 1: Experimental Disk Temperature Distribution in Modified J33 Split-Disk Rotor at Speeds up to 6000 RPM

    NASA Technical Reports Server (NTRS)

    Schramm, Wilson B.; Ziemer, Robert R.

    1952-01-01

    An experimental investigation is being conducted at the Lewis laboratory to establish general principles for the design of noncritical turbine rotor configurations. This investigation includes evaluation of cooling effectiveness, structural stability, cooling-air flow distribution characteristics, and methods of supplying cooling air to the turbine rotor blades. Prior to design of a noncritical rotor, a standard turbine rotor of a commerical turbojet engine was split in the plane of rotation and machined to provide a passage for distributing cooling air to the base of each blade. The rotor was fitted with nontwisted, hollow, aircooled blades containing nine tubes in the coolant passage. In the investigation reported herein, the modified turbine rotor operated successfully up to speeds of 6000 rpm with ratios of cooling-air to combustion-gas flow as low as 0.02. The disk temperatures observed at these conditions were below 450 0 F when cooling air at 100 F was used from the laboratory air system. The calculated disk temperatures based on the correlation method presented for rated engine conditions were well below 1000 F at a cooling-air flow ratio of 0.02, which is considered adequate for a noncritical rotor. An appreciable difference in temperature level existed between the forward and rear disks. This temperature difference probably introduced undesirable disk stress distributions as a result of the relative elongations of the two disks. This investigation was terminated at 6000 rpm so that slight changes in the engine configuration could be made to relieve this condition.

  16. Film cooling in a plane turbine cascade

    NASA Astrophysics Data System (ADS)

    Goldstein, R. J.; Eckert, E. R. G.; Ito, S.

    A mass transfer study is conducted using simulated turbine blades in order to determine the influence of surface curvature on film cooling effectiveness for ratios of injected gas/mainstream air density of approximately unity and of 2. Great care must be taken in applying the results of flat-plate film-cooling experiments with a single row of holes to the design of turbine blade coolant systems; relative trends for increasing or decreasing effectiveness depend on the momentum flux ratio as well as the sign of the radius-of-curvature of the surface. Greater blockage of the mainstream decreases the relative penetration of jets into the freestream, thereby decreasing the importance of the pressure gradient normal to the surface.

  17. Analytical investigation of chord size and cooling methods on turbine blade cooling requirements. Book 1: Sections 1 through 8 and appendixes A through I

    NASA Technical Reports Server (NTRS)

    Faulkner, F. E.

    1971-01-01

    A study was conducted to determine the effect of chord size on air cooled turbine blades. In the preliminary design phase, eight turbine blade cooling configurations in 0.75-in., 1.0-in., and 1.5-in. chord sizes were analyzed to determine the maximum turbine inlet temperature capabilities. A pin fin convection cooled configuration and a film-impingement cooled configuration were selected for a final design analysis in which the maximum turbine inlet temperature was determined as a function of the cooling air inlet temperature and the turbine inlet total pressure for each of the three chord sizes. The cooling air flow requirements were also determined for a varying cooling air inlet temperature with a constant turbine inlet temperature. It was determined that allowable turbine inlet temperature increases with increasing chord for the convection cooled and transpiration cooled designs, however, the film-convection cooled designs did not have a significant change in turbine inlet temperature with chord.

  18. Film cooling on the pressure surface of a turbine vane

    NASA Technical Reports Server (NTRS)

    Gauntner, J. W.; Gladden, H. J.

    1977-01-01

    Film-cooling-air ejection from the pressure surface of a turbine vane was investigated, and experimental data are presented. This investigation was conducted in a four-vane cascade on a J75-size turbine vane that had a double row of staggered holes in line with the primary flow and located downstream of the leading edge region. The results showed that: (1) the average effectiveness of film-convection cooling was higher than that of either film cooling or convection cooling separately; (2) the addition of small quantities of film-cooling air always increased the cooling effectiveness relative to the zero-injection case; however, (3) the injected film must exceed a certain threshold value to obtain a beneficial effect of film cooling relative to convection cooling alone.

  19. Recent developments in turbine blade internal cooling.

    PubMed

    Han, J C; Dutta, S

    2001-05-01

    This paper focuses on turbine blade internal cooling. Internal cooling is achieved by passing the coolant through several rib-enhanced serpentine passages inside the blade and extracting the heat from the outside of the blades. Both jet impingement and pin-fin-cooling are also used as a method of internal cooling. In the past number of years there has been considerable progress in turbine blade internal cooling research and this paper is limited to reviewing a few selected publications to reflect recent developments in turbine blade internal cooling. PMID:11460626

  20. Structural cooling fluid tube for supporting a turbine component and supplying cooling fluid to transition section

    SciTech Connect

    Charron, Richard; Pierce, Daniel

    2015-08-11

    A shaft cover support for a gas turbine engine is disclosed. The shaft cover support not only provides enhanced support to a shaft cover of the gas turbine engine, but also includes a cooling fluid chamber for passing fluids from a rotor air cooling supply conduit to an inner ring cooling manifold. Furthermore, the shaft cover support may include a cooling shield supply extending from the cooling fluid chamber between the radially outward inlet and the radially inward outlet on the radially extending region and in fluid communication with the cooling fluid chamber for providing cooling fluids to a transition section. The shaft cover support may also provide additional stiffness and reduce interference of the flow from the compressor. In addition, the shaft cover support accommodates a transition section extending between compressor and turbine sections of the gas turbine engine.

  1. Microtextured Surfaces for Turbine Blade Impingement Cooling

    NASA Technical Reports Server (NTRS)

    Fryer, Jack

    2014-01-01

    Gas turbine engine technology is constantly challenged to operate at higher combustor outlet temperatures. In a modern gas turbine engine, these temperatures can exceed the blade and disk material limits by 600 F or more, necessitating both internal and film cooling schemes in addition to the use of thermal barrier coatings. Internal convective cooling is inadequate in many blade locations, and both internal and film cooling approaches can lead to significant performance penalties in the engine. Micro Cooling Concepts, Inc., has developed a turbine blade cooling concept that provides enhanced internal impingement cooling effectiveness via the use of microstructured impingement surfaces. These surfaces significantly increase the cooling capability of the impinging flow, as compared to a conventional untextured surface. This approach can be combined with microchannel cooling and external film cooling to tailor the cooling capability per the external heating profile. The cooling system then can be optimized to minimize impact on engine performance.

  2. Measured effects of coolant injection on the performance of a film cooled turbine

    NASA Technical Reports Server (NTRS)

    Mcdonel, J. D.; Eiswerth, J. E.

    1977-01-01

    Tests have been conducted on a 20-inch diameter single-stage air-cooled turbine designed to evaluate the effects of film cooling air on turbine aerodynamic performance. The present paper reports the results of five test configurations, including two different cooling designs and three combinations of cooled and solid airfoils. A comparison is made of the experimental results with a previously published analytical method of evaluating coolant injection effects on turbine performance.

  3. Turbine stator vane segment having internal cooling circuits

    DOEpatents

    Jones, Raymond Joseph; Burns, James Lee; Bojappa, Parvangada Ganapathy; Jones, Schotsch Margaret

    2003-01-01

    A turbine stator vane includes outer and inner walls each having outer and inner chambers and a vane extending between the outer and inner walls. The vane includes first, second, third, fourth and fifth cavities for flowing a cooling medium. The cooling medium enters the outer chamber of the outer wall, flows through an impingement plate for impingement cooling of the outer band wall defining in part the hot gas path and through openings in the first, second and fourth cavities for flow radially inwardly, cooling the vane. The spent cooling medium flows into the inner wall and inner chamber for flow through an impingement plate radially outwardly to cool the inner wall. The spent cooling medium flows through the third cavity for egress from the turbine vane segment from the outer wall. The first, second or third cavities contain inserts having impingement openings for impingement cooling of the vane walls. The fifth cavity provides air cooling for the trailing edge.

  4. Cooling system for a bearing of a turbine rotor

    DOEpatents

    Schmidt, Mark Christopher

    2002-01-01

    In a gas turbine, a bore tube assembly radially inwardly of an aft bearing conveys cooling steam to the buckets of the turbine and returns the cooling steam to a return. To cool the bearing and thermally insulate the bearing from the cooling steam paths, a radiation shield is spaced from the bore tube assembly by a dead air gap. Additionally, an air passageway is provided between the radiation shield and the inner surface of an aft shaft forming part of the rotor. Air is supplied from an inlet for flow along the passage and radially outwardly through bores in the aft shaft disk to cool the bearing and insulate it from transfer of heat from the cooling steam.

  5. Heat transfer performance comparison of steam and air in gas turbine cooling channels with different rib angles

    NASA Astrophysics Data System (ADS)

    Shi, Xiaojun; Gao, Jianmin; Xu, Liang; Li, Fajin

    2013-11-01

    Using steam as working fluid to replace compressed air is a promising cooling technology for internal cooling passages of blades and vanes. The local heat transfer characteristics and the thermal performance of steam flow in wide aspect ratio channels ( W/ H = 2) with different angled ribs on two opposite walls have been experimentally investigated in this paper. The averaged Nusselt number ratios and the friction factor ratios of steam and air in four ribbed channels were also measured under the same test conditions for comparison. The Reynolds number range is 6,000-70,000. The rib angles are 90°, 60°, 45°, and 30°, respectively. The rib height to hydraulic diameter ratio is 0.047. The pitch-to-rib height ratio is 10. The results show that the Nusselt number ratios of steam are 1.19-1.32 times greater than those of air over the range of Reynolds numbers studied. For wide aspect ratio channels using steam as the coolant, the 60° angled ribs has the best heat transfer performance and is recommended for cooling design.

  6. Turbine blade cooling using Coulomb repulsion

    NASA Astrophysics Data System (ADS)

    Breidenthal, Robert; Colannino, Joseph; Dees, John; Goodson, David; Krichtafovitch, Igor; Prevo, Tracy

    2012-11-01

    Video photography and thermocouples reveal the effect of an electric field on the flow around a stationary, idealized turbine blade downstream of a combustor. The hot products of combustion naturally include positive ions. When the blade is an electrode and elevated to a positive potential, it tends to attract the free electrons and repel the positive ions. Due to their lower mass, the light electrons are rapidly swept toward the blade, while the positive ions are repelled. As they collide with the neutrals in the hot gas, the positive ions transfer their momentum so that a Coulomb body force is exerted on the hot gas. Cool, compressed air is injected out of the stationary blade near its leading edge to form a layer of film cooling. In contrast to the hot combustion products, the cool air is not ionized. At the interface between the hot gas and the cool air, the Coulomb repulsion force acts on the former but not the latter, analogous to gravity at a stratified interface. An effective Richardson number representing the ratio of potential to kinetic energy characterizes the topography of the interface. When the electric field is turned on, the repulsion of the hot gas from the idealized blade is evident in video recordings and thermocouple measurements.

  7. Cold-air annular-cascade investigation of aerodynamic performance of cooled turbine vanes. 2: Trailing-edge ejection, film cooling, and transpiration cooling

    NASA Technical Reports Server (NTRS)

    Goldman, L. J.; Mclallin, K. L.

    1975-01-01

    The aerodynamic performance of four different cooled vane configurations was experimentally determined in a full-annular cascade at a primary- to coolant-total-temperature ratio of 1.0. The vanes were tested over a range of coolant flow rates and pressure ratios. Overall vane efficiencies were obtained and compared, where possible, with the results obtained in a four-vane, annular-sector cascade. The vane efficiency and exit flow conditions as functions of radial position were also determined and compared with solid (uncooled) vane results.

  8. Floating air riding seal for a turbine

    DOEpatents

    Ebert, Todd A

    2016-08-16

    A floating air riding seal for a gas turbine engine with a rotor and a stator, an annular piston chamber with an axial moveable annular piston assembly within the annular piston chamber formed in the stator, an annular cavity formed on the annular piston assembly that faces a seal surface on the rotor, where the axial moveable annular piston includes an inlet scoop on a side opposite to the annular cavity that scoops up the swirling cooling air and directs the cooling air to the annular cavity to form an air cushion with the seal surface of the rotor.

  9. Cooling arrangement for a tapered turbine blade

    SciTech Connect

    Liang, George

    2010-07-27

    A cooling arrangement (11) for a highly tapered gas turbine blade (10). The cooling arrangement (11) includes a pair of parallel triple-pass serpentine cooling circuits (80,82) formed in an inner radial portion (50) of the blade, and a respective pair of single radial channel cooling circuits (84,86) formed in an outer radial portion (52) of the blade (10), with each single radial channel receiving the cooling fluid discharged from a respective one of the triple-pass serpentine cooling circuit. The cooling arrangement advantageously provides a higher degree of cooling to the most highly stressed radially inner portion of the blade, while providing a lower degree of cooling to the less highly stressed radially outer portion of the blade. The cooling arrangement can be implemented with known casting techniques, thereby facilitating its use on highly tapered, highly twisted Row 4 industrial gas turbine blades that could not be cooled with prior art cooling arrangements.

  10. Cooling scheme for turbine hot parts

    DOEpatents

    Hultgren, Kent Goran; Owen, Brian Charles; Dowman, Steven Wayne; Nordlund, Raymond Scott; Smith, Ricky Lee

    2000-01-01

    A closed-loop cooling scheme for cooling stationary combustion turbine components, such as vanes, ring segments and transitions, is provided. The cooling scheme comprises: (1) an annular coolant inlet chamber, situated between the cylinder and blade ring of a turbine, for housing coolant before being distributed to the turbine components; (2) an annular coolant exhaust chamber, situated between the cylinder and the blade ring and proximate the annular coolant inlet chamber, for collecting coolant exhaust from the turbine components; (3) a coolant inlet conduit for supplying the coolant to said coolant inlet chamber; (4) a coolant exhaust conduit for directing coolant from said coolant exhaust chamber; and (5) a piping arrangement for distributing the coolant to and directing coolant exhaust from the turbine components. In preferred embodiments of the invention, the cooling scheme further comprises static seals for sealing the blade ring to the cylinder and flexible joints for attaching the blade ring to the turbine components.

  11. Compatibility of gas turbine materials with steam cooling

    SciTech Connect

    Desai, V.; Tamboli, D.; Patel, Y.

    1995-10-01

    Gas turbines had been traditionally used for peak load plants and remote locations as they offer advantage of low installation costs and quick start up time. Their use as a base load generator had not been feasible owing to their poor efficiency. However, with the advent of gas turbines based combined cycle plants (CCPs), continued advances in efficiency are being made. Coupled with ultra low NO{sub x} emissions, coal compatibility and higher unit output, gas turbines are now competing with conventional power plants for base load power generation. Currently, the turbines are designed with TIT of 2300{degrees}F and metal temperatures are maintained around 1700{degrees}F by using air cooling. New higher efficiency ATS turbines will have TIT as high as 2700{degrees}F. To withstand this high temperature improved materials, coatings, and advances in cooling system and design are warranted. Development of advanced materials with better capabilities specifically for land base applications are time consuming and may not be available by ATS time frame or may prove costly for the first generation ATS gas turbines. Therefore improvement in the cooling system of hot components, which can take place in a relatively shorter time frame, is important. One way to improve cooling efficiency is to use better cooling agent. Steam as an alternate cooling agent offers attractive advantages because of its higher specific heat (almost twice that of air) and lower viscosity.

  12. Effect of cooling-hole geometry on aerodynamic performance of a film-cooled turbine vane tested with cold air in a two-dimensional cascade

    NASA Technical Reports Server (NTRS)

    Kline, J. F.; Stabe, R. G.; Moffitt, T. P.

    1978-01-01

    The effect of the orientation and cooling-hole size on turbine-vane aerodynamic losses was evaluated. The contribution of individual vane regions to the overall effect was also investigated. Test configurations were based upon a representative configuration having 45 spanwise rows of holes spaced about the entire vane profile. Nominal hole diameters of 0.0254 and 0.0356 cm and nominal hole orientations of 35 deg, 45 deg, and 55 deg from the local vane surface and 0 deg, 45 deg, and 90 deg from the main-stream flow direction were investigated. Flow conditions and aerodynamic losses were determined by vane-exit surveys of total pressure, static pressure, and flow angle.

  13. Multi-pass cooling for turbine airfoils

    DOEpatents

    Liang, George

    2011-06-28

    An airfoil for a turbine vane of a gas turbine engine. The airfoil includes an outer wall having pressure and suction sides, and a radially extending cooling cavity located between the pressure and suction sides. A plurality of partitions extend radially through the cooling cavity to define a plurality of interconnected cooling channels located at successive chordal locations through the cooling cavity. The cooling channels define a serpentine flow path extending in the chordal direction. Further, the cooling channels include a plurality of interconnected chambers and the chambers define a serpentine path extending in the radial direction within the serpentine path extending in the chordal direction.

  14. Effects of 'Cooled' Cooling Air on Pre-Swirl Nozzle Design

    NASA Technical Reports Server (NTRS)

    Scricca, J. A.; Moore, K. D.

    2006-01-01

    It is common practice to use Pre-Swirl Nozzles to facilitate getting the turbine blade cooling air onboard the rotating disk with minimum pressure loss and reduced temperature. Higher engine OPR's and expanded aircraft operating envelopes have pushed cooling air temperatures to the limits of current disk materials and are stressing the capability to cool the blade with practical levels of cooling air flow. Providing 'Cooled' Cooling Air is one approach being considered to overcome these limitations. This presentation looks at how the introduction of 'Cooled' Cooling Air impacts the design of the Pre-Swirl Nozzles, specifically in relation to the radial location of the nozzles.

  15. Cooled turbine vane with endcaps

    DOEpatents

    Cunha, Frank J.; Schiavo, Jr., Anthony L.; Nordlund, Raymond Scott; Malow, Thomas; McKinley, Barry L.

    2002-01-01

    A turbine vane assembly which includes an outer endcap having a plurality of generally straight passages and passage segments therethrough, an inner endcap having a plurality of passages and passage segments therethrough, and a vane assembly having an outer shroud, an airfoil body, and an inner shroud. The outer shroud, airfoil body and inner shroud each have a plurality of generally straight passages and passage segments therethrough as well. The outer endcap is coupled to the outer shroud so that outer endcap passages and said outer shroud passages form a fluid circuit. The inner endcap is coupled to the inner shroud so that the inner end cap passages and the inner shroud passages from a fluid circuit. Passages in the vane casting are in fluid communication with both the outer shroud passages and the inner shroud passages. Passages in the outer endcap may be coupled to a cooling system that supplies a coolant and takes away the heated exhaust.

  16. Engine tests on a cooled gas turbine stage

    NASA Astrophysics Data System (ADS)

    Graf, H. J.

    1985-09-01

    A new cooling system was designed for the 45 MW gas turbine type 8. Extensive tests were carried out in a new power station to verify the reliability of the cooled components. Wall temperatures were measured using thermocouples, thermal paints and pyrometers. Cooling air temperature, pressure and mass flow measurements allowed a detailed analysis of the first stage under operating conditions. The results and comparisons with design calculations are presented. The applicability and accuracy of the three measuring techniques are discussed.

  17. Cooling techniques for gas turbine airfoils: A survey

    NASA Astrophysics Data System (ADS)

    Metzger, D. E.

    1985-09-01

    A brief general background discussion of turbine heat transfer and cooling with compressor discharge air is given. Specific reference is made to a selection of current research areas for gas turbine engine cooling, including blade tip heat transfer, heat transfer in serpentine passages, multiple jet array impingement, heat transfer in pin fin arrays, disk heat transfer, and film cooling. An overview of various experimental methods used to acquire heat transfer data is also given, with an emphasis on newer methods used to acquire detailed local convection heat transfer information.

  18. Cooled highly twisted airfoil for a gas turbine engine

    SciTech Connect

    Kildea, R.J.

    1988-04-19

    This patent describes a cooled highly twisted airfoil for use in a gas turbine engine. The airfoil has a first cooling air cavity adjacent a leading edge of the airfoil, and a second cooling air cavity, separated from the first cavity by a wall. The second cavity provides cooling air to the first cavity by means of cooling holes provided in the wall. The improvement is characterized by: the wall comprising an integrally formed, continuous warped wall, defined as a surface of revolution about an axis, the axis determined such that the axis intersects the plane of a section close to a desired centerline of a series of impingement holes aligned in opposition to the leading edge, whereby cooling air is directed relatively precisely to the leading edge of the highly twisted airfoil through the impingement holes.

  19. Flow Integrating Section for a Gas Turbine Engine in Which Turbine Blades are Cooled by Full Compressor Flow

    SciTech Connect

    Steward, W. Gene

    1999-11-14

    Routing of full compressor flow through hollow turbine blades achieves unusually effective blade cooling and allows a significant increase in turbine inlet gas temperature and, hence, engine efficiency. The invention, ''flow integrating section'' alleviates the turbine dissipation of kinetic energy of air jets leaving the hollow blades as they enter the compressor diffuser.

  20. AIR COOLED NEUTRONIC REACTOR

    DOEpatents

    Fermi, E.; Szilard, L.

    1958-05-27

    A nuclear reactor of the air-cooled, graphite moderated type is described. The active core consists of a cubicle mass of graphite, approximately 25 feet in each dimension, having horizontal channels of square cross section extending between two of the opposite faces, a plurality of cylindrical uranium slugs disposed in end to end abutting relationship within said channels providing a space in the channels through which air may be circulated, and a cadmium control rod extending within a channel provided in the moderator. Suitable shielding is provlded around the core, as are also provided a fuel element loading and discharge means, and a means to circulate air through the coolant channels through the fuel charels to cool the reactor.

  1. Turbine cooling configuration selection and design optimization for the high-reliability gas turbine. Final report

    SciTech Connect

    Smith, M J; Suo, M

    1981-04-01

    The potential of advanced turbine convectively air-cooled concepts for application to the Department of Energy/Electric Power Research Institute (EPRI) Advanced Liquid/Gas-Fueled Engine Program was investigated. Cooling of turbine airfoils is critical technology and significant advances in cooling technology will permit higher efficiency coal-base-fuel gas turbine energy systems. Two new airfoil construction techniques, bonded and wafer, were the principal designs considered. In the bonded construction, two airfoil sections having intricate internal cooling configurations are bonded together to form a complete blade or vane. In the wafer construction, a larger number (50 or more) of wafers having intricate cooling flow passages are bonded together to form a complete blade or vane. Of these two construction techniques, the bonded airfoil is considered to be lower in risk and closer to production readiness. Bonded airfoils are being used in aircraft engines. A variety of industrial materials were evaluated for the turbine airfoils. A columnar grain nickel alloy was selected on the basis of strength and corrosion resistance. Also, cost of electricity and reliability were considered in the final concept evaluation. The bonded airfoil design yielded a 3.5% reduction in cost-of-electricity relative to a baseline Reliable Engine design. A significant conclusion of this study was that the bonded airfoil convectively air-cooled design offers potential for growth to turbine inlet temperatures above 2600/sup 0/F with reasonable development risk.

  2. Film cooling air pocket in a closed loop cooled airfoil

    DOEpatents

    Yu, Yufeng Phillip; Itzel, Gary Michael; Osgood, Sarah Jane; Bagepalli, Radhakrishna; Webbon, Waylon Willard; Burdgick, Steven Sebastian

    2002-01-01

    Turbine stator vane segments have radially inner and outer walls with vanes extending between them. The inner and outer walls are compartmentalized and have impingement plates. Steam flowing into the outer wall plenum passes through the impingement plate for impingement cooling of the outer wall upper surface. The spent impingement steam flows into cavities of the vane having inserts for impingement cooling the walls of the vane. The steam passes into the inner wall and through the impingement plate for impingement cooling of the inner wall surface and for return through return cavities having inserts for impingement cooling of the vane surfaces. To provide for air film cooing of select portions of the airfoil outer surface, at least one air pocket is defined on a wall of at least one of the cavities. Each air pocket is substantially closed with respect to the cooling medium in the cavity and cooling air pumped to the air pocket flows through outlet apertures in the wall of the airfoil to cool the same.

  3. Cold-air performance of a 12.766-centimeter-tip-diameter axial-flow cooled turbine. 1: Design and performance of a solid blade configuration

    NASA Technical Reports Server (NTRS)

    Haas, J. E.; Kofskey, M. G.

    1975-01-01

    A solid blade version of a single-stage, axial-flow turbine was investigated to determine its performance over a range of speeds from 0 to 105 percent of equivalent design speed and over a range of total to static pressure ratios from 1.62 to 5.07. The results of this investigation will be used as a baseline for comparison with those obtained from a cooled version of this turbine.

  4. Liquid cooled counter flow turbine bucket

    DOEpatents

    Dakin, James T.

    1982-09-21

    Means and a method are provided whereby liquid coolant flows radially outward through coolant passages in a liquid cooled turbine bucket under the influence of centrifugal force while in contact with countercurrently flowing coolant vapor such that liquid is entrained in the flow of vapor resulting in an increase in the wetted cooling area of the individual passages.

  5. Cooling arrangement for a gas turbine component

    DOEpatents

    Lee, Ching-Pang; Heneveld, Benjamin E

    2015-02-10

    A cooling arrangement (82) for a gas turbine engine component, the cooling arrangement (82) having a plurality of rows (92, 94, 96) of airfoils (98), wherein adjacent airfoils (98) within a row (92, 94, 96) define segments (110, 130, 140) of cooling channels (90), and wherein outlets (114, 134) of the segments (110, 130) in one row (92, 94) align aerodynamically with inlets (132, 142) of segments (130, 140) in an adjacent row (94, 96) to define continuous cooling channels (90) with non continuous walls (116, 120), each cooling channel (90) comprising a serpentine shape.

  6. Real-Time Closed Loop Modulated Turbine Cooling

    NASA Technical Reports Server (NTRS)

    Shyam, Vikram; Culley, Dennis E.; Eldridge, Jeffrey; Jones, Scott; Woike, Mark; Cuy, Michael

    2014-01-01

    It has been noted by industry that in addition to dramatic variations of temperature over a given blade surface, blade-to-blade variations also exist despite identical design. These variations result from manufacturing variations, uneven wear and deposition over the life of the part as well as limitations in the uniformity of coolant distribution in the baseline cooling design. It is proposed to combine recent advances in optical sensing, actuation, and film cooling concepts to develop a workable active, closed-loop modulated turbine cooling system to improve by 10 to 20 the turbine thermal state over the flight mission, to improve engine life and to dramatically reduce turbine cooling air usage and aircraft fuel burn. A reduction in oxides of nitrogen (NOx) can also be achieved by using the excess coolant to improve mixing in the combustor especially for rotorcraft engines. Recent patents filed by industry and universities relate to modulating endwall cooling using valves. These schemes are complex, add weight and are limited to the endwalls. The novelty of the proposed approach is twofold 1) Fluidic diverters that have no moving parts are used to modulate cooling and can operate under a wide range of conditions and environments. 2) Real-time optical sensing to map the thermal state of the turbine has never been attempted in realistic engine conditions.

  7. Gas turbine premixer with internal cooling

    DOEpatents

    York, William David; Johnson, Thomas Edward; Lacy, Benjamin Paul; Stevenson, Christian Xavier

    2012-12-18

    A system that includes a turbine fuel nozzle comprising an air-fuel premixer. The air-fuel premixed includes a swirl vane configured to swirl fuel and air in a downstream direction, wherein the swirl vane comprises an internal coolant path from a downstream end portion in an upstream direction through a substantial length of the swirl vane.

  8. Closed circuit steam cooled turbine shroud and method for steam cooling turbine shroud

    DOEpatents

    Burdgick, Steven Sebastian; Sexton, Brendan Francis; Kellock, Iain Robertson

    2002-01-01

    A turbine shroud cooling cavity is partitioned to define a plurality of cooling chambers for sequentially receiving cooling steam and impingement cooling of the radially inner wall of the shoud. An impingement baffle is provided in each cooling chamber for receiving the cooling media from a cooling media inlet in the case of the first chamber or from the immediately upstream chamber in the case of the second through fourth chambers and includes a plurality of impingement holes for effecting the impingement cooling of the shroud inner wall.

  9. Gas turbine bucket with impingement cooled platform

    DOEpatents

    Jones, Raphael Durand

    2002-01-01

    In a turbine bucket having an airfoil portion and a root portion, with a substantially planar platform at an interface between the airfoil portion and root portion, a platform cooling arrangement including at least one bore in the root portion and at least one impingement cooling tube seated in the bore, the tube extending beyond the bore with an outlet in close proximity to a targeted area on an underside of the platform.

  10. An infrared technique for evaluating turbine airfoil cooling designs

    SciTech Connect

    Sweeney, P.C.; Rhodes, J.F.

    2000-01-01

    An experimental approach is used to evaluate turbine airfoil cooling designs for advanced gas turbine engine applications by incorporating double-wall film-cooled design features into large-scale flat plate specimens. An infrared (IR) imaging system is used to make detailed, two-dimensional steady-state measurements of flat plate surface temperature with spatial resolution on the order of 0.4 mm. The technique employs a cooled zinc selenide window transparent to infrared radiation and calibrates the IR temperature readings to reference thermocouples embedded in each specimen, yielding a surface temperature measurement accuracy of {+-} 4 C. With minimal thermocouple installation required, the flat plate/IR approach is cost effective, essentially nonintrusive, and produces abundant results quickly. Design concepts can proceed from art to part to data in a manner consistent with aggressive development schedules. The infrared technique is demonstrated here by considering the effect of film hole injection angle for a staggered array of film cooling holes integrated with a highly effective internal cooling pattern. Heated free stream air and room temperature cooling air are used to produce a nominal temperature ratio of 2 over a range of blowing ratios from 0.7 to 1.5. Results were obtained at hole angles of 90 and 30 deg for two different hole spacings and are presented in terms of overall cooling effectiveness.

  11. Structural cooling fluid tube for supporting a turbine component and supplying cooling fluid

    SciTech Connect

    Charron, Richard; Pierce, Daniel

    2015-02-24

    A shaft cover support for a gas turbine engine is disclosed. The shaft cover support not only provides enhanced support to a shaft cover of the gas turbine engine, but also includes a cooling fluid chamber for passing fluids from a rotor air cooling supply conduit to an inner ring cooling manifold. As such, the shaft cover support accomplishes in a single component what was only partially accomplished in two components in conventional configurations. The shaft cover support may also provide additional stiffness and reduce interference of the flow from the compressor. In addition, the shaft cover support accommodates a transition section extending between compressor and turbine sections of the engine. The shaft cover support has a radially extending region that is offset from the inlet and outlet that enables the shaft cover support to surround the transition, thereby reducing the overall length of this section of the engine.

  12. Near-wall serpentine cooled turbine airfoil

    SciTech Connect

    Lee, Ching-Pang

    2014-10-28

    A serpentine coolant flow path is formed by inner walls in a cavity between pressure and suction side walls of a turbine airfoil, the cavity partitioned by one or more transverse partitions into a plurality of continuous serpentine cooling flow streams each having a respective coolant inlet.

  13. Composite Matrix Cooling Scheme for Small Gas Turbine Combustors

    NASA Technical Reports Server (NTRS)

    Paskin, Marc D.; Ross, Phillip T.; Mongia, Hukam C.; Acosta, Waldo A.

    1990-01-01

    The design, manufacture, and testing of a compliant metal/ceramic (CMC) wall cooling concept-implementing combustor for small gas turbine engines has been undertaken by a joint U.S. Army/NASA technology development program. CMC in principle promises greater wall cooling effectiveness than conventional designs and materials, thereby facilitating a substantial reduction in combustor cooling air requirements and furnishing greater airflow for the control of burner outlet temperature patterns as well as improving thermodynamic efficiency and reducing pollutant emissions and smoke levels. Rig test results have confirmed the projected benefits of the CMC concept at combustor outlet temperatures of the order of 2460 F, at which approximately 80 percent less cooling air than conventionally required was being employed by the CMC combustor.

  14. Cooled High-temperature Radial Turbine Program 2

    NASA Technical Reports Server (NTRS)

    Snyder, Philip H.

    1991-01-01

    The objective of this program was the design and fabrication of a air-cooled high-temperature radial turbine (HTRT) intended for experimental evaluation in a warm turbine test facility at the LeRC. The rotor and vane were designed to be tested as a scaled version (rotor diameter of 14.4 inches diameter) of a 8.021 inch diameter rotor designed to be capable of operating with a rotor inlet temperature (RIT) of 2300 F, a nominal mass flow of 4.56 lbm/sec, a work level of equal or greater than 187 Btu/lbm, and efficiency of 86 percent or greater. The rotor was also evaluated to determine it's feasibility to operate at 2500 F RIT. The rotor design conformed to the rotor blade flow path specified by NASA for compatibility with their test equipment. Fabrication was accomplished on three rotors, a bladeless rotor, a solid rotor, and an air-cooled rotor.

  15. Debris trap in a turbine cooling system

    DOEpatents

    Wilson, Ian David

    2002-01-01

    In a turbine having a rotor and a plurality of stages, each stage comprising a row of buckets mounted on the rotor for rotation therewith; and wherein the buckets of at least one of the stages are cooled by steam, the improvement comprising at least one axially extending cooling steam supply conduit communicating with an at least partially annular steam supply manifold; one or more axially extending cooling steam feed tubes connected to the manifold at a location radially outwardly of the cooling steam supply conduit, the feed tubes arranged to supply cooling steam to the buckets of at least one of the plurality of stages; the manifold extending radially beyond the feed tubes to thereby create a debris trap region for collecting debris under centrifugal loading caused by rotation of the rotor.

  16. Plugging of cooling holes in film-cooled turbine vanes

    NASA Technical Reports Server (NTRS)

    Deadmore, D. L.; Lowell, C. E.

    1977-01-01

    The plugging of vane cooling holes by impurities in a marine gas turbine was closely simulated in burner rig tests where dopants were added to the combustion products of a clean fuel (Jet-A). Hole plugging occurred when liquid phases, resulting from the dopants, were present in the combustion products. Increasing flame temperature and dopant concentration resulted in an increased rate of deposition and hole plugging.

  17. Internal cooling circuit for gas turbine bucket

    SciTech Connect

    Hyde, Susan Marie; Davis, Richard Mallory

    2005-10-25

    In a gas turbine bucket having a shank portion and an airfoil portion having leading and trailing edges and pressure and suction sides, an internal cooling circuit, the internal cooling circuit having a serpentine configuration including plural radial outflow passages and plural radial inflow passages, and wherein a coolant inlet passage communicates with a first of the radial outflow passages along the trailing edge, the first radial outflow passage having a plurality of radially extending and radially spaced elongated rib segments extending between and connecting the pressure and suction sides in a middle region of the first passage to prevent ballooning of the pressure and suction sides at the first radial outflow passage.

  18. Partially turbulated trailing edge cooling passages for gas turbine nozzles

    DOEpatents

    Thatcher, Jonathan Carl; Burdgick, Steven Sebastian

    2001-01-01

    A plurality of passages are spaced one from the other along the length of a trailing edge of a nozzle vane in a gas turbine. The passages lie in communication with a cavity in the vane for flowing cooling air from the cavity through the passages through the tip of the trailing edge into the hot gas path. Each passage is partially turbulated and includes ribs in an aft portion thereof to provide enhanced cooling effects adjacent the tip of the trailing edge. The major portions of the passages are smooth bore. By this arrangement, reduced temperature gradients across the trailing edge metal are provided. Additionally, the inlets to each of the passages have a restriction whereby a reduced magnitude of compressor bleed discharge air is utilized for trailing edge cooling purposes.

  19. Cooled silicon nitride stationary turbine vane risk reduction. Final report

    SciTech Connect

    Holowczak, John

    1999-12-31

    The purpose of this program was to reduce the technical risk factors for demonstration of air cooled silicon nitride turbine vanes. The effort involved vane prototype fabrication efforts at two U.S. based gas turbine grade silicon nitride component manufacturers. The efficacy of the cooling system was analyzed via a thermal time/temperature flow test technique previously at UTRC. By having multiple vendors work on parts fabrication, the chance of program success increased for producing these challenging components. The majority of the effort under this contract focused on developing methods for, and producing, the complex thin walled silicon nitride vanes. Components developed under this program will undergo engine environment testing within N00014-96-2-0014.

  20. Liquid-cooling technology for gas turbines review and status

    NASA Technical Reports Server (NTRS)

    Vanfossen, G. J., Jr.; Stepka, F. S.

    1978-01-01

    A review of research related to liquid cooling of gas turbines was conducted and an assessment of the state of the art was made. Various methods of liquid cooling turbines were reviewed. Examples and results with test and demonstrator turbines utilizing these methods along with the advantages and disadvantages of the various methods are discussed.

  1. Refractory inserts used to form cooling passages in cast superalloy turbine vanes

    NASA Technical Reports Server (NTRS)

    Terpay, A.

    1973-01-01

    Economical technique has been developed for manufacturing air-cooled turbine blades and vanes for gas turbine engines. Process uses tungsten inserts to form coolant passages. After casting, inserts are reduced to tungsten oxide during sublimation with oxygen at elevated temperature. Tungsten oxide is leached out of coolant passages with a molten salt solution.

  2. Cooled airfoil in a turbine engine

    SciTech Connect

    Vitt, Paul H; Kemp, David A; Lee, Ching-Pang; Marra, John J

    2015-04-21

    An airfoil in a gas turbine engine includes an outer wall and an inner wall. The outer wall includes a leading edge, a trailing edge opposed from the leading edge in a chordal direction, a pressure side, and a suction side. The inner wall is coupled to the outer wall at a single chordal location and includes portions spaced from the pressure and suction sides of the outer wall so as to form first and second gaps between the inner wall and the respective pressure and suction sides. The inner wall defines a chamber therein and includes openings that provide fluid communication between the respective gaps and the chamber. The gaps receive cooling fluid that provides cooling to the outer wall as it flows through the gaps. The cooling fluid, after traversing at least substantial portions of the gaps, passes into the chamber through the openings in the inner wall.

  3. Low pressure cooling seal system for a gas turbine engine

    DOEpatents

    Marra, John J

    2014-04-01

    A low pressure cooling system for a turbine engine for directing cooling fluids at low pressure, such as at ambient pressure, through at least one cooling fluid supply channel and into a cooling fluid mixing chamber positioned immediately downstream from a row of turbine blades extending radially outward from a rotor assembly to prevent ingestion of hot gases into internal aspects of the rotor assembly. The low pressure cooling system may also include at least one bleed channel that may extend through the rotor assembly and exhaust cooling fluids into the cooling fluid mixing chamber to seal a gap between rotational turbine blades and a downstream, stationary turbine component. Use of ambient pressure cooling fluids by the low pressure cooling system results in tremendous efficiencies by eliminating the need for pressurized cooling fluids for sealing this gap.

  4. Thermodynamic study of air-cycle and mercury-vapor-cycle systems for refrigerating cooling air for turbines or other components

    NASA Technical Reports Server (NTRS)

    Nachtigall, Alfred J; Freche, John C; Esgar, Jack B

    1956-01-01

    An analysis of air refrigeration systems indicated that air cycles are generally less satisfactory than simple heat exchangers unless high component efficiencies and high values of heat-exchanger effectiveness can be obtained. A system employing a mercury-vapor cycle appears to be feasible for refrigerating air that must enter the system at temperature levels of approximately 1500 degrees R, and this cycle is more efficient than the air cycle. Weight of the systems was not considered. The analysis of the systems is presented in a generalized dimensionless form.

  5. Investigations of Air-cooled Turbine Rotors for Turbojet Engines II : Mechanical Design, Stress Analysis, and Burst Test of Modified J33 Split-disk Rotor / Richard H. Kemp and Merland L. Moseson

    NASA Technical Reports Server (NTRS)

    Kemp, Richard H; Moseson, Merland L

    1952-01-01

    A full-scale J33 air-cooled split turbine rotor was designed and spin-pit tested to destruction. Stress analysis and spin-pit results indicated that the rotor in a J33 turbojet engine, however, showed that the rear disk of the rotor operated at temperatures substantially higher than the forward disk. An extension of the stress analysis to include the temperature difference between the two disks indicated that engine modifications are required to permit operation of the two disks at more nearly the same temperature level.

  6. Ceramic thermal-barrier coatings for cooled turbines

    NASA Technical Reports Server (NTRS)

    Liebert, C. H.; Stepka, F. S.

    1976-01-01

    Ceramic thermal-barrier coatings on hot engine parts have the potential to reduce metal temperatures, coolant requirements, cost, and complexity of the cooling configuration, and to increase life, turbine efficiency and gas temperature. Coating systems consisting of a plasma-sprayed layer of zirconia stabilized with either yttria, magnesia or calcia over a thin alloy bond coat have been developed, their potential analyzed and their durability and benefits evaluated in a turbojet engine. The coatings on air-cooled rotating blades were in good condition after completing as many as 500 two-minute cycles of engine operation between full power at a gas temperature of 1644 K and flameout, or as much as 150 hours of steady-state operation on cooled vanes and blades at gas temperatures as high as 1644 K witn 35 start and stop cycles. On the basis of durability and processing cost, the yttria-stabilized zirconia was considered the best of the three coatings investigated.

  7. Turbine engine component with cooling passages

    DOEpatents

    Arrell, Douglas J.; James, Allister W.

    2012-01-17

    A component for use in a turbine engine including a first member and a second member associated with the first member. The second member includes a plurality of connecting elements extending therefrom. The connecting elements include securing portions at ends thereof that are received in corresponding cavities formed in the first member to attach the second member to the first member. The connecting elements are constructed to space apart a first surface of the second member from a first surface of the first member such that at least one cooling passage is formed between adjacent connecting elements and the first surface of the second member and the first surface of the first member.

  8. Analysis and comparison of wall cooling schemes for advanced gas turbine applications

    NASA Technical Reports Server (NTRS)

    Colladay, R. S.

    1972-01-01

    The relative performance of (1) counterflow film cooling, (2) parallel-flow film cooling, (3) convection cooling, (4) adiabatic film cooling, (5) transpiration cooling, and (6) full-coverage film cooling was investigated for heat loading conditions expected in future gas turbine engines. Assumed in the analysis were hot-gas conditions of 2200 K (3500 F) recovery temperature, 5 to 40 atmospheres total pressure, and 0.6 gas Mach number and a cooling air supply temperature of 811 K (1000 F). The first three cooling methods involve film cooling from slots. Counterflow and parallel flow describe the direction of convection cooling air along the inside surface of the wall relative to the main gas flow direction. The importance of utilizing the heat sink available in the coolant for convection cooling prior to film injection is illustrated.

  9. Theoretical Evaluation of Methods of Cooling the Blades of Gas Turbines

    NASA Technical Reports Server (NTRS)

    Sanders, J. C.; Mendelson, Alexander

    1947-01-01

    A study was made of heat transfer in turbine blades and the effects on blade temperature of cooling the blade root and tip, changing the dimensions of the blades, raising the cycle temperatures, insulating with ceramics, and cooling by circulation of air or water through hollow blades.

  10. Metal temperatures and coolant flow in a wire cloth transpiration cooled turbine vane

    NASA Technical Reports Server (NTRS)

    Gladden, H. J.

    1975-01-01

    An experimental heat transfer investigation was conducted on an air-cooled turbine vane made from wire-wound cloth material and supported by a central strut. Vane temperature data obtained are compared with temperature data from two full-coverage film-cooled vanes made of different laminated construction. Measured porous-airfoil temperatures are compared with predicted temperatures.

  11. Liquid-cooling technology for gas turbines - Review and status

    NASA Technical Reports Server (NTRS)

    Van Fossen, G. J., Jr.; Stepka, F. S.

    1978-01-01

    After a brief review of past efforts involving the forced-convection cooling of gas turbines, the paper surveys the state of the art of the liquid cooling of gas turbines. Emphasis is placed on thermosyphon methods of cooling, including those utilizing closed, open, and closed-loop thermosyphons; other methods, including sweat, spray and stator cooling, are also discussed. The more significant research efforts, design data, correlations, and analytical methods are mentioned and voids in technology are summarized.

  12. Cooling circuit for a gas turbine bucket and tip shroud

    DOEpatents

    Willett, Fred Thomas; Itzel, Gary Michael; Stathopoulos, Dimitrios; Plemmons, Larry Wayne; Plemmons, Helen M.; Lewis, Doyle C.

    2002-01-01

    An open cooling circuit for a gas turbine bucket wherein the bucket has an airfoil portion, and a tip shroud, the cooling circuit including a plurality of radial cooling holes extending through the airfoil portion and communicating with an enlarged internal area within the tip shroud before exiting the tip shroud such that a cooling medium used to cool the airfoil portion is subsequently used to cool the tip shroud.

  13. Turbine airfoil with controlled area cooling arrangement

    SciTech Connect

    Liang, George

    2010-04-27

    A gas turbine airfoil (10) includes a serpentine cooling path (32) with a plurality of channels (34,42,44) fluidly interconnected by a plurality of turns (38,40) for cooling the airfoil wall material. A splitter component (50) is positioned within at least one of the channels to bifurcate the channel into a pressure-side channel (46) passing in between the outer wall (28) and the inner wall (30) of the pressure side (24) and a suction-side channel (48) passing in between the outer wall (28) and the inner wall (30) of the suction side (26) longitudinally downstream of an intermediate height (52). The cross-sectional area of the pressure-side channel (46) and suction-side channel (48) are thereby controlled in spite of an increasing cross-sectional area of the airfoil along its longitudinal length, ensuring a sufficiently high mach number to provide a desired degree of cooling throughout the entire length of the airfoil.

  14. Fabrication of cooled radial turbine rotor

    NASA Technical Reports Server (NTRS)

    Hammer, A. N.; Aigret, G. G.; Psichogios, T. P.; Rodgers, C.

    1986-01-01

    A design and fabrication program was conducted to evaluate a unique concept for constructing a cooled, high temperature radial turbine rotor. This concept, called split blade fabrication was developed as an alternative to internal ceramic coring. In this technique, the internal cooling cavity is created without flow dividers or any other detail by a solid (and therefore stronger) ceramic plate which can be more firmly anchored within the casting shell mold than can conventional detailed ceramic cores. Casting is conducted in the conventional manner, except that the finished product, instead of having finished internal cooling passages, is now a split blade. The internal details of the blade are created separately together with a carrier sheet. The inserts are superalloy. Both are produced by essentially the same software such that they are a net fit. The carrier assemblies are loaded into the split blade and the edges sealed by welding. The entire wheel is Hot Isostatic Pressed (HIPed), braze bonding the internal details to the inside of the blades. During this program, two wheels were successfully produced by the split blade fabrication technique.

  15. Natural Flow Air Cooled Photovoltaics

    NASA Astrophysics Data System (ADS)

    Tanagnostopoulos, Y.; Themelis, P.

    2010-01-01

    Our experimental study aims to investigate the improvement in the electrical performance of a photovoltaic installation on buildings through cooling of the photovoltaic panels with natural air flow. Our experimental study aims to investigate the improvement in the electrical performance of a photovoltaic installation on buildings through cooling of the photovoltaic panels with natural air flow. We performed experiments using a prototype based on three silicon photovoltaic modules placed in series to simulate a typical sloping building roof with photovoltaic installation. In this system the air flows through a channel on the rear side of PV panels. The potential for increasing the heat exchange from the photovoltaic panel to the circulating air by the addition of a thin metal sheet (TMS) in the middle of air channel or metal fins (FIN) along the air duct was examined. The operation of the device was studied with the air duct closed tightly to avoid air circulation (CLOSED) and the air duct open (REF), with the thin metal sheet (TMS) and with metal fins (FIN). In each case the experiments were performed under sunlight and the operating parameters of the experimental device determining the electrical and thermal performance of the system were observed and recorded during a whole day and for several days. We collected the data and form PV panels from the comparative diagrams of the experimental results regarding the temperature of solar cells, the electrical efficiency of the installation, the temperature of the back wall of the air duct and the temperature difference in the entrance and exit of the air duct. The comparative results from the measurements determine the improvement in electrical performance of the photovoltaic cells because of the reduction of their temperature, which is achieved by the naturally circulating air.

  16. Effects of Thermal Barrier Coatings on Approaches to Turbine Blade Cooling

    NASA Technical Reports Server (NTRS)

    Boyle, Robert J.

    2007-01-01

    Reliance on Thermal Barrier Coatings (TBC) to reduce the amount of air used for turbine vane cooling is beneficial both from the standpoint of reduced NOx production, and as a means of improving cycle efficiency through improved component efficiency. It is shown that reducing vane cooling from 10 to 5 percent of mainstream air can lead to NOx reductions of nearly 25 percent while maintaining the same rotor inlet temperature. An analysis is given which shows that, when a TBC is relied upon in the vane thermal design process, significantly less coolant is required using internal cooling alone compared to film cooling. This is especially true for small turbines where internal cooling without film cooling permits the surface boundary layer to remain laminar over a significant fraction of the vane surface.

  17. Laminated turbine vane design and fabrication. [utilizing film cooling as a cooling system

    NASA Technical Reports Server (NTRS)

    Hess, W. G.

    1979-01-01

    A turbine vane and associated endwalls designed for advanced gas turbine engine conditions are described. The vane design combines the methods of convection cooling and selective areas of full coverage film cooling. The film cooling technique is utilized on the leading edge, pressure side, and endwall regions. The turbine vane involves the fabrication of airfoils from a stack of laminates with cooling passages photoetched on the surface. Cold flow calibration tests, a thermal analysis, and a stress analysis were performed on the turbine vanes.

  18. Wind turbine generators having wind assisted cooling systems and cooling methods

    DOEpatents

    Bagepalli, Bharat; Barnes, Gary R.; Gadre, Aniruddha D.; Jansen, Patrick L.; Bouchard, Jr., Charles G.; Jarczynski, Emil D.; Garg, Jivtesh

    2008-09-23

    A wind generator includes: a nacelle; a hub carried by the nacelle and including at least a pair of wind turbine blades; and an electricity producing generator including a stator and a rotor carried by the nacelle. The rotor is connected to the hub and rotatable in response to wind acting on the blades to rotate the rotor relative to the stator to generate electricity. A cooling system is carried by the nacelle and includes at least one ambient air inlet port opening through a surface of the nacelle downstream of the hub and blades, and a duct for flowing air from the inlet port in a generally upstream direction toward the hub and in cooling relation to the stator.

  19. Steam cooling system for a gas turbine

    DOEpatents

    Wilson, Ian David; Barb, Kevin Joseph; Li, Ming Cheng; Hyde, Susan Marie; Mashey, Thomas Charles; Wesorick, Ronald Richard; Glynn, Christopher Charles; Hemsworth, Martin C.

    2002-01-01

    The steam cooling circuit for a gas turbine includes a bore tube assembly supplying steam to circumferentially spaced radial tubes coupled to supply elbows for transitioning the radial steam flow in an axial direction along steam supply tubes adjacent the rim of the rotor. The supply tubes supply steam to circumferentially spaced manifold segments located on the aft side of the 1-2 spacer for supplying steam to the buckets of the first and second stages. Spent return steam from these buckets flows to a plurality of circumferentially spaced return manifold segments disposed on the forward face of the 1-2 spacer. Crossover tubes couple the steam supply from the steam supply manifold segments through the 1-2 spacer to the buckets of the first stage. Crossover tubes through the 1-2 spacer also return steam from the buckets of the second stage to the return manifold segments. Axially extending return tubes convey spent cooling steam from the return manifold segments to radial tubes via return elbows.

  20. Advanced liner-cooling techniques for gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Norgren, C. T.; Riddlebaugh, S. M.

    1985-01-01

    Component research for advanced small gas turbine engines is currently underway at the NASA Lewis Research Center. As part of this program, a basic reverse-flow combustor geometry was being maintained while different advanced liner wall cooling techniques were investigated. Performance and liner cooling effectiveness of the experimental combustor configuration featuring counter-flow film-cooled panels is presented and compared with two previously reported combustors featuring: splash film-cooled liner walls; and transpiration cooled liner walls (Lamilloy).

  1. Rotational effects on turbine blade cooling

    SciTech Connect

    Govatzidakis, G.J.; Guenette, G.R.; Kerrebrock, J.L.

    1995-10-01

    An experimental investigation of the influence of rotation on the heat transfer in a smooth, rectangular passage rotating in the orthogonal mode is presented. The passage simulates one of the cooling channels found in gas turbine blades. A constant heat flux is imposed on the model with either inward or outward flow. The effects of rotation and buoyancy on the Nusselt number were quantified by systematically varying the Rotation number, Density Ratio, Reynolds number, and Buoyancy parameter. The experiment utilizes a high resolution infrared temperature measurement technique in order to measure the wall temperature distribution. The experimental results show that the rotational effects on the Nusselt number are significant and proper turbine blade design must take into account the effects of rotation, buoyancy, and flow direction. The behavior of the Nusselt number distribution depends strongly on the particular side, axial position, flow direction, and the specific range of the scaling parameters. The results show a strong coupling between buoyancy and Corollas effects throughout the passage. For outward flow, the trailing side Nusselt numbers increase with Rotation number relative to stationary values. On the leading side, the Nusselt numbers tended to decrease with rotation near the inlet and subsequently increased farther downstream in the passage. The Nusselt numbers on the side walls generally increased with rotation. For inward flow, the Nusselt numbers generally improved relative to stationary results, but increases in the Nusselt number were relatively smaller than in the case of outward flow. For outward and inward flows, increasing the density ratio generally tended to decrease Nusselt numbers on the leading and trailing sides, but the exact behavior and magnitude depended on the local axial position and specific range of Buoyancy parameters.

  2. Thermosyphon Method for Cooling the Rotor Blades of High-Temperature Steam Turbines

    NASA Astrophysics Data System (ADS)

    Bogomolov, Alexander R.; Temnikova, Elena Yu.

    2016-02-01

    The design scheme of closed two-phase thermosyphon were suggested that can provide standard thermal operation of blades of high-temperature steam turbine. The method for thermosyphon calculation is developed. The example of thermal calculation was implemented, it showed that to cool the steam turbine blades at their heating by high-temperature steam, the heat can be removed in the rear part of the blades by air with the temperature of about 440°C.

  3. Air and water cooled modulator

    DOEpatents

    Birx, Daniel L.; Arnold, Phillip A.; Ball, Don G.; Cook, Edward G.

    1995-01-01

    A compact high power magnetic compression apparatus and method for delivering high voltage pulses of short duration at a high repetition rate and high peak power output which does not require the use of environmentally unacceptable fluids such as chlorofluorocarbons either as a dielectric or as a coolant, and which discharges very little waste heat into the surrounding air. A first magnetic switch has cooling channels formed therethrough to facilitate the removal of excess heat. The first magnetic switch is mounted on a printed circuit board. A pulse transformer comprised of a plurality of discrete electrically insulated and magnetically coupled units is also mounted on said printed board and is electrically coupled to the first magnetic switch. The pulse transformer also has cooling means attached thereto for removing heat from the pulse transformer. A second magnetic switch also having cooling means for removing excess heat is electrically coupled to the pulse transformer. Thus, the present invention is able to provide high voltage pulses of short duration at a high repetition rate and high peak power output without the use of environmentally unacceptable fluids and without discharging significant waste heat into the surrounding air.

  4. Air and water cooled modulator

    DOEpatents

    Birx, D.L.; Arnold, P.A.; Ball, D.G.; Cook, E.G.

    1995-09-05

    A compact high power magnetic compression apparatus and method are disclosed for delivering high voltage pulses of short duration at a high repetition rate and high peak power output which does not require the use of environmentally unacceptable fluids such as chlorofluorocarbons either as a dielectric or as a coolant, and which discharges very little waste heat into the surrounding air. A first magnetic switch has cooling channels formed therethrough to facilitate the removal of excess heat. The first magnetic switch is mounted on a printed circuit board. A pulse transformer comprised of a plurality of discrete electrically insulated and magnetically coupled units is also mounted on said printed board and is electrically coupled to the first magnetic switch. The pulse transformer also has cooling means attached thereto for removing heat from the pulse transformer. A second magnetic switch also having cooling means for removing excess heat is electrically coupled to the pulse transformer. Thus, the present invention is able to provide high voltage pulses of short duration at a high repetition rate and high peak power output without the use of environmentally unacceptable fluids and without discharging significant waste heat into the surrounding air. 9 figs.

  5. Cooling of gas turbines IX : cooling effects from use of ceramic coatings on water-cooled turbine blades

    NASA Technical Reports Server (NTRS)

    Brown, W Byron; Livingood, John N B

    1948-01-01

    The hottest part of a turbine blade is likely to be the trailing portion. When the blades are cooled and when water is used as the coolant, the cooling passages are placed as close as possible to the trailing edge in order to cool this portion. In some cases, however, the trailing portion of the blade is so narrow, for aerodynamic reasons, that water passages cannot be located very near the trailing edge. Because ceramic coatings offer the possibility of protection for the trailing part of such narrow blades, a theoretical study has been made of the cooling effect of a ceramic coating on: (1) the blade-metal temperature when the gas temperature is unchanged, and (2) the gas temperature when the metal temperature is unchanged. Comparison is also made between the changes in the blade or gas temperatures produced by ceramic coatings and the changes produced by moving the cooling passages nearer the trailing edge. This comparison was made to provide a standard for evaluating the gains obtainable with ceramic coatings as compared to those obtainable by constructing the turbine blade in such a manner that water passages could be located very near the trailing edge.

  6. Air-cooled, hydrogen-air fuel cell

    NASA Technical Reports Server (NTRS)

    Shelekhin, Alexander B. (Inventor); Bushnell, Calvin L. (Inventor); Pien, Michael S. (Inventor)

    1999-01-01

    An air-cooled, hydrogen-air solid polymer electrolyte (SPE) fuel cell with a membrane electrode assembly operatively associated with a fluid flow plate having at least one plate cooling channel extending through the plate and at least one air distribution hole extending from a surface of the cathode flow field into the plate cooling channel.

  7. Turbine component cooling channel mesh with intersection chambers

    DOEpatents

    Lee, Ching-Pang; Marra, John J

    2014-05-06

    A mesh (35) of cooling channels (35A, 35B) with an array of cooling channel intersections (42) in a wall (21, 22) of a turbine component. A mixing chamber (42A-C) at each intersection is wider (W1, W2)) than a width (W) of each of the cooling channels connected to the mixing chamber. The mixing chamber promotes swirl, and slows the coolant for more efficient and uniform cooling. A series of cooling meshes (M1, M2) may be separated by mixing manifolds (44), which may have film cooling holes (46) and/or coolant refresher holes (48).

  8. Air cooling : an experimental method of evaluating the cooling effect of air streams on air-cooled cylinders

    NASA Technical Reports Server (NTRS)

    Alcock, J F

    1927-01-01

    In this report is described an experimental method which the writer has evolved for dealing with air-cooled engines, and some of the data obtained by its means. Methods of temperature measurement and cooling are provided.

  9. Analysis of Turbine Blade Relative Cooling Flow Factor Used in the Subroutine Coolit Based on Film Cooling Correlations

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.

    2015-01-01

    Heat transfer correlations of data on flat plates are used to explore the parameters in the Coolit program used for calculating the quantity of cooling air for controlling turbine blade temperature. Correlations for both convection and film cooling are explored for their relevance to predicting blade temperature as a function of a total cooling flow which is split between external film and internal convection flows. Similar trends to those in Coolit are predicted as a function of the percent of the total cooling flow that is in the film. The exceptions are that no film or 100 percent convection is predicted to not be able to control blade temperature, while leaving less than 25 percent of the cooling flow in the convection path results in nearing a limit on convection cooling as predicted by a thermal effectiveness parameter not presently used in Coolit.

  10. Cooled variable nozzle radial turbine for rotor craft applications

    NASA Technical Reports Server (NTRS)

    Rogo, C.

    1981-01-01

    An advanced, small 2.27 kb/sec (5 lbs/sec), high temperature, variable area radial turbine was studied for a rotor craft application. Variable capacity cycles including single-shaft and free-turbine engine configurations were analyzed to define an optimum engine design configuration. Parametric optimizations were made on cooled and uncooled rotor configurations. A detailed structural and heat transfer analysis was conducted to provide a 4000-hour life HP turbine with material properties of the 1988 time frame. A pivoted vane and a moveable sidewall geometry were analyzed. Cooling and variable geometry penalties were included in the cycle analysis. A variable geometry free-turbine engine configuration with a design 1477K (2200 F) inlet temperature and a compressor pressure ratio of 16:1 was selected. An uncooled HP radial turbine rotor with a moveable sidewall nozzle showed the highest performance potential for a time weighted duty cycle.

  11. Air cooled absorption chillers for solar cooling applications

    NASA Astrophysics Data System (ADS)

    Biermann, W. J.; Reimann, R. C.

    1982-03-01

    The chemical composition of a 'best' absorption refrigerant system is identified, and those properties of the system necessary to design hot water operated, air cooled chilling equipment are determined. Air cooled chillers from single family residential sizes into the commercial rooftop size range are designed and operated.

  12. Experimental simulation of impingement cooling in midchord region of turbine blade

    NASA Astrophysics Data System (ADS)

    Li, Liguo; Jiang, Jun; Chang, Haiping; Zhang, Donglia

    1989-10-01

    Simulation experiments have been completed to research the characteristics of impingement cooling in the midchord region of a turbine blade for a given geometric parameter of jet array, initial crossflow, and pressure ratios of the film-cooling exhaust. Comparative experiments have been made on curvilinear and flat plate impinged surfaces, and impinged surfaces with and without a chordwise fin. Moreover, the thermal patterns from a liquid crystal show that the jet hole arrangement and distance between cooled surface and jet plate affect the heat transfer distributions for jet array impingement. Finally, the effects of the cooling air axially injected into guide tube on radial jet flow have also been determined.

  13. Gas turbine combustion chamber with air scoops

    SciTech Connect

    Mumford, S.E.; Smed, J.P.

    1989-12-19

    This patent describes a gas turbine combustion chamber. It comprises: means for admission of fuel to the upstream end thereof and discharge of hot gases from the downstream end thereof, and a combustion chamber wall, having an outer surface, with apertures therethrough, and air scoops provided through the apertures to direct air into the combustion chamber.

  14. Effect of thermal barrier coatings on the performance of steam and water-cooled gas turbine/steam turbine combined cycle system

    NASA Technical Reports Server (NTRS)

    Nainiger, J. J.

    1978-01-01

    An analytical study was made of the performance of air, steam, and water-cooled gas-turbine/steam turbine combined-cycle systems with and without thermal-barrier coatings. For steam cooling, thermal barrier coatings permit an increase in the turbine inlet temperature from 1205 C (2200 F), resulting in an efficiency improvement of 1.9 percentage points. The maximum specific power improvement with thermal barriers is 32.4 percent, when the turbine inlet temperature is increased from 1425 C (2600 F) to 1675 C (3050 F) and the airfoil temperature is kept the same. For water cooling, the maximum efficiency improvement is 2.2 percentage points at a turbine inlet temperature of 1683 C (3062 F) and the maximum specific power improvement is 36.6 percent by increasing the turbine inlet temperature from 1425 C (2600 F) to 1730 C (3150 F) and keeping the airfoil temperatures the same. These improvements are greater than that obtained with combined cycles using air cooling at a turbine inlet temperature of 1205 C (2200 F). The large temperature differences across the thermal barriers at these high temperatures, however, indicate that thermal stresses may present obstacles to the use of coatings at high turbine inlet temperatures.

  15. Cooling systems for ultra-high temperature turbines.

    PubMed

    Yoshida, T

    2001-05-01

    This paper describes an introduction of research and development activities on steam cooling in gas turbines at elevated temperature of 1500 C and 1700 C level, partially including those on water cooling. Descriptions of a new cooling system that employs heat pipes are also made. From the view point of heat transfer, its promising applicability is shown with experimental data and engine performance numerical evaluation. PMID:11460628

  16. Thermotechnical performance of an air-cooled tuyere with air cooling channels in series

    NASA Astrophysics Data System (ADS)

    Shen, Yuansheng; Zhou, Yuanyuan; Zhu, Tao; Duan, Guangbin

    2016-03-01

    To reduce the cooling air consumption for an air-cooled tuyere, an air-cooled tuyere with air cooling channels in series is developed based on several hypotheses, i.e., a transparent medium in the blast furnace, among others, and the related mathematical models are introduced and developed. Referring to the data from a BF site, the thermotechnical computation for the air-cooled tuyere was performed, and the results show that when the temperature of the inlet cooling air increases, the temperatures for the outlet cooling air, the outer surface of the tuyere, the walls of the air cooling channels and the center channel as well as the heat going into the center channel increase, but the heat absorbed by the cooling air flowing through the air cooling channels decreases. When the cooling air flow rate under the standard state increases, the physical parameters mentioned above change in an opposite directions. Compared to a water-cooled tuyere, the energy savings for an air-cooled tuyere are more than 0.23 kg/min standard coal.

  17. Performance of Air-cooled Engine Cylinders Using Blower Cooling

    NASA Technical Reports Server (NTRS)

    Schey, Oscar W; Ellerbrock, Herman H , Jr

    1936-01-01

    An investigation was made to obtain information on the minimum quantity of air and power required to cool conventional air cooled cylinders at various operating conditions when using a blower. The results of these tests show that the minimum power required for satisfactory cooling with an overall blower efficiency of 100 percent varied from 2 to 6 percent of the engine power depending on the operating conditions. The shape of the jacket had a large effect on the cylinder temperatures. Increasing the air speed over the front of the cylinder by keeping the greater part of the circumference of the cylinder covered by the jacket reduced the temperatures over the entire cylinder.

  18. Numerical evaluation of single central jet for turbine disk cooling

    NASA Astrophysics Data System (ADS)

    Subbaraman, M. R.; Hadid, A. H.; McConnaughey, P. K.

    The cooling arrangement of the Space Shuttle Main Engine High Pressure Oxidizer Turbopump (HPOTP) incorporates two jet rings, each of which produces 19 high-velocity coolant jets. At some operating conditions, the frequency of excitation associated with the 19 jets coincides with the natural frequency of the turbine blades, contributing to fatigue cracking of blade shanks. In this paper, an alternate turbine disk cooling arrangement, applicable to disk faces of zero hub radius, is evaluated, which consists of a single coolant jet impinging at the center of the turbine disk. Results of the CFD analysis show that replacing the jet ring with a single central coolant jet in the HPOTP leads to an acceptable thermal environment at the disk rim. Based on the predictions of flow and temperature fields for operating conditions, the single central jet cooling system was recommended for implementation into the development program of the Technology Test Bed Engine at NASA Marshall Space Flight Center.

  19. Aerodynamic investigation of an air-cooled axial-flow turbine. Part 2: Rotor blade tip-clearance effects on overall turbine performance and internal gas flow conditions: Experimental results and prediction methods

    NASA Technical Reports Server (NTRS)

    Yamamoto, A.; Takahara, K.; Nouse, H.; Mimura, F.; Inoue, S.; Usui, H.

    1977-01-01

    Total turbine blade performance was investigated while changing the blade tip clearance in three ways. The internal flow at the moving blade outlet point was measured. Experimental results were compared with various theoretical methods. Increased blade clearance leads to decreased turbine efficiency.

  20. Air extraction in gas turbines burning coal-derived gas

    SciTech Connect

    Yang, Tah-teh; Agrawal, A.K.; Kapat, J.S.

    1993-11-01

    In the first phase of this contracted research, a comprehensive investigation was performed. Principally, the effort was directed to identify the technical barriers which might exist in integrating the air-blown coal gasification process with a hot gas cleanup scheme and the state-of-the-art, US made, heavy-frame gas turbine. The guiding rule of the integration is to keep the compressor and the expander unchanged if possible. Because of the low-heat content of coal gas and of the need to accommodate air extraction, the combustor and perhaps, the flow region between the compressor exit and the expander inlet might need to be modified. In selecting a compressed air extraction scheme, one must consider how the scheme affects the air supply to the hot section of the turbine and the total pressure loss in the flow region. Air extraction must preserve effective cooling of the hot components, such as the transition pieces. It must also ensure proper air/fuel mixing in the combustor, hence the combustor exit pattern factor. The overall thermal efficiency of the power plant can be increased by minimizing the total pressure loss in the diffusers associated with the air extraction. Therefore, a study of airflow in the pre- and dump-diffusers with and without air extraction would provide information crucial to attaining high-thermal efficiency and to preventing hot spots. The research group at Clemson University suggested using a Griffith diffuser for the prediffuser and extracting air from the diffuser inlet. The present research establishes that the analytically identified problems in the impingement cooling flow are factual. This phase of the contracted research substantiates experimentally the advantage of using the Griffith diffuser with air extraction at the diffuser inlet.

  1. Air riding seal for a turbine

    DOEpatents

    Mills, Jacob A; Brown, Wesley D; Sexton, Thomas D; Jones, Russell B

    2016-07-19

    An air riding seal between a rotor and a stator in a turbine of a gas turbine engine, where an annular piston is movable in an axial direction within a housing that extends from the stator, and a bellows is secured to the annular piston to form a flexible air passageway from a compressed air inlet through the annular piston and into a cushion cavity that forms an air riding seal between the annular piston and the rotor sealing surface. In another embodiment, a flexible seal secured to and extending from the annular piston forms a sealing surface between the annular piston chamber and the annular piston to provide a seal and allow for axial movement.

  2. Gas turbine bucket cooling circuit and related process

    DOEpatents

    Lewis, Doyle C.; Barb, Kevin Joseph

    2002-01-01

    A turbine bucket includes an airfoil portion having leading and trailing edges; at least one radially extending cooling passage within the airfoil portion, the airfoil portion joined to a platform at a radially inner end of the airfoil portion; a dovetail mounting portion enclosing a cooling medium supply passage; and, a crossover passage in fluid communication with the cooling medium supply passage and with at least one radially extending cooling passage, the crossover passage having a portion extending along and substantially parallel to an underside surface of the platform.

  3. Trap seal for open circuit liquid cooled turbines

    DOEpatents

    Grondahl, Clayton M.; Germain, Malcolm R.

    1980-01-01

    An improved trap seal for open circuit liquid cooled turbines is disclosed. The trap seal of the present invention includes an annular recess formed in the supply conduit of cooling channels formed in the airfoil of the turbine buckets. A cylindrical insert is located in the annular recesses and has a plurality of axial grooves formed along the outer periphery thereof and a central recess formed in one end thereof. The axial grooves and central recess formed in the cylindrical insert cooperate with the annular recess to define a plurality of S-shaped trap seals which permit the passage of liquid coolant but prohibit passage of gaseous coolant.

  4. Effects of turbine cooling assumptions on performance and sizing of high-speed civil transport

    NASA Technical Reports Server (NTRS)

    Senick, Paul F.

    1992-01-01

    The analytical study presented examines the effects of varying turbine cooling assumptions on the performance of a high speed civil transport propulsion system as well as the sizing sensitivity of this aircraft to these performance variations. The propulsion concept employed in this study was a two spool, variable cycle engine with a sea level thrust of 55,000 lbf. The aircraft used for this study was a 250 passenger vehicle with a cruise Mach number of 2.4 and 5000 nautical mile range. The differences in turbine cooling assumptions were represented by varying the amount of high pressure compressor bleed air used to cool the turbines. It was found that as this cooling amount increased, engine size and weight increased, but specific fuel consumption (SFC) decreased at takeoff and climb only. Because most time is spent at cruise, the SFC advantage of the higher bleed engines seen during subsonic flight was minimized and the lower bleed, lighter engines led to the lowest takeoff gross weight vehicles. Finally, the change in aircraft takeoff gross weight versus turbine cooling level is presented.

  5. 14 CFR 23.1111 - Turbine engine bleed air system.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Turbine engine bleed air system. 23.1111 Section 23.1111 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION... Induction System § 23.1111 Turbine engine bleed air system. For turbine engine bleed air systems,...

  6. 14 CFR 23.1111 - Turbine engine bleed air system.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Turbine engine bleed air system. 23.1111 Section 23.1111 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION... Induction System § 23.1111 Turbine engine bleed air system. For turbine engine bleed air systems,...

  7. 14 CFR 23.1111 - Turbine engine bleed air system.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Turbine engine bleed air system. 23.1111 Section 23.1111 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION... Induction System § 23.1111 Turbine engine bleed air system. For turbine engine bleed air systems,...

  8. 14 CFR 23.1111 - Turbine engine bleed air system.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Turbine engine bleed air system. 23.1111 Section 23.1111 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION... Induction System § 23.1111 Turbine engine bleed air system. For turbine engine bleed air systems,...

  9. 14 CFR 23.1111 - Turbine engine bleed air system.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Turbine engine bleed air system. 23.1111 Section 23.1111 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION... Induction System § 23.1111 Turbine engine bleed air system. For turbine engine bleed air systems,...

  10. An experimental investigation of a gas turbine disk cooling system

    NASA Astrophysics Data System (ADS)

    Kobayashi, N.; Matsumato, M.; Shizuya, M.

    1983-03-01

    The results of an experimental study of the cooling of a model disk similar to an engine disk are compared with the results obtained by three-dimensional finite difference computation, and it is reconfirmed that the determination of cooling air temperature is one of the most important data for predicting the disk temperature. The minimum cooling air flow rate necessary to prevent ingress of external hot gas is determined by the fluctuation of cooling air temperature inside the wheel space with the external axial hot gas flow for values of the rotational Reynolds number of 0-6.5 million. The effect of rotational speed on the minimum cooling air flow rate is found to be negligible, and it is shown that the determination of the ingress of hot gas using the pressure difference criterion underestimates the minimum cooling air flow rate.

  11. Cooling considerations for design of a radial inflow turbine

    NASA Technical Reports Server (NTRS)

    Hamed, A.; Sheoran, Y.; Tabakoff, W.

    1977-01-01

    A numerical study to determine the temperature distribution in the rotor of a radial inflow turbine is presented. Internal cooling passages are modeled in the present formulation in order to carry out solid and coolant temperature computations simultaneously resulting in a considerable computer time savings. The stresses due to centrifugal and thermal loadings are determined in an actual rotor and the effect of cooling design on its mechanical integrity is discussed.

  12. Modernization of experimental air turbine VT-400

    NASA Astrophysics Data System (ADS)

    Klimko, Marek; Okresa, Daniel

    2016-06-01

    This article briefly describes a modernization of the experimental device - air turbine (VT-400), which is a part of the research activities at the Department of Power System Engineering, University of West Bohemia (KKE-UWB). This device serves for the research of steam turbine blades within the framework of a long-term cooperation between the Department and Doosan Skoda Power (DSPW). Due to the age of the device, some necessary changes had to be performed and some obsolete components had to be replaced or new ones added. A part of this article is also a comparison with the previous state and an evaluation of the contribution after the reconstruction.

  13. Cold-air annular-cascade investigation of aerodynamic performance of core-engine-cooled turbine vanes. 1: Solid-vane performance and facility description

    NASA Technical Reports Server (NTRS)

    Goldman, L. J.; Mclallin, K. L.

    1975-01-01

    The aerodynamic performance of a solid (uncooled) version of a core engine cooled stator vane was experimentally determined in a full-annular cascade, where three-dimensional effects could be obtained. The solid vane, which serves as a basis for comparison with subsequent cooled tests, was tested over a range of aftermixed critical velocity ratios of 0.57 to 0.90. Overall vane aftermixed efficiencies were obtained over this critical velocity ratio range and compared with results from a two-dimensional cascade. The variation in vane efficiency and aftermixed flow conditions with circumferential and radial position were obtained and compared with design values. Vane surface static-pressure distributions were also measured and compared with theoretical results.

  14. Temperature distribution study in a cooled radial inflow turbine rotor

    NASA Technical Reports Server (NTRS)

    Hamed, A.; Baskharone, E.; Tabakoff, W.

    1976-01-01

    A numerical study to determine the temperature distribution in the rotor of a radial inflow turbine is presented. The study is based on the use of the finite element method in the three dimensional heat conduction problem. Different cooling techniques with various coolant to primary mass flow ratios are investigated. The resulting temperature distribution in the rotor are presented for comparison.

  15. Mid-section of a can-annular gas turbine engine with a cooling system for the transition

    SciTech Connect

    Wiebe, David J.; Rodriguez, Jose L.

    2015-12-08

    A cooling system is provided for a transition (420) of a gas turbine engine (410). The cooling system includes a cowling (460) configured to receive an air flow (111) from an outlet of a compressor section of the gas turbine engine (410). The cowling (460) is positioned adjacent to a region of the transition (420) to cool the transition region upon circulation of the air flow within the cowling (460). The cooling system further includes a manifold (121) to directly couple the air flow (111) from the compressor section outlet to an inlet (462) of the cowling (460). The cowling (460) is configured to circulate the air flow (111) within an interior space (426) of the cowling (460) that extends radially outward from an inner diameter (423) of the cowling to an outer diameter (424) of the cowling at an outer surface.

  16. Aerodynamic performance of a Wells air turbine

    NASA Astrophysics Data System (ADS)

    Raghunathan, S.; Tan, C. P.

    1983-06-01

    Experiments were performed in a unidirectional flow rig to assess the performance of the Wells self-rectifying air turbine. Results indicated that the efficiency of the turbine was very sensitive to the Reynolds number based on blade chord. Increase in Reynolds number by a factor of three resulted in an increase in peak efficiency from 37 to 60 percent. Increases in the solidity of the blade produced increases in pressure drop and power output but decreases in efficiency. The hub-to-tip ratio had only a weak influence on the turbine performance but is critical for starting conditions. It is concluded that a hub-to-tip ratio of 0.6 and a solidity of 0.6 are the most favorable values, taking into consideration both the starting and running performances.

  17. Power dissipation of air turbine VT - 400

    NASA Astrophysics Data System (ADS)

    Noga, Tomas; Žitek, Pavel

    2016-06-01

    This article provides an overview of ongoing systematic research of a turbine stage efficiency on a model air turbine VT 400. It contains an analysis of existing mathematical relations for a rotor friction dissipation calculation, on which basis a practical procedure of a calculation of those dissipations is recommended. Friction dissipations in the turbine rotor were divided into three main tasks: disc friction dissipations, shaft friction dissipations and dissipations in bearings. A contribution of performed work lies in the fact, that there is a dependence of rotor friction losses on its speed and a stage reaction has been revealed. This knowledge is completely essential for a further research, and will lead to more precise results of experiments. For the future, we plan to adjust the measuring track by adding a moment collar. We also assume an experimental verification of calculated friction losses.

  18. Thermochemically recuperated and steam cooled gas turbine system

    DOEpatents

    Viscovich, Paul W.; Bannister, Ronald L.

    1995-01-01

    A gas turbine system in which the expanded gas from the turbine section is used to generate the steam in a heat recovery steam generator and to heat a mixture of gaseous hydrocarbon fuel and the steam in a reformer. The reformer converts the hydrocarbon gas to hydrogen and carbon monoxide for combustion in a combustor. A portion of the steam from the heat recovery steam generator is used to cool components, such as the stationary vanes, in the turbine section, thereby superheating the steam. The superheated steam is mixed into the hydrocarbon gas upstream of the reformer, thereby eliminating the need to raise the temperature of the expanded gas discharged from the turbine section in order to achieve effective conversion of the hydrocarbon gas.

  19. Thermochemically recuperated and steam cooled gas turbine system

    DOEpatents

    Viscovich, P.W.; Bannister, R.L.

    1995-07-11

    A gas turbine system is described in which the expanded gas from the turbine section is used to generate the steam in a heat recovery steam generator and to heat a mixture of gaseous hydrocarbon fuel and the steam in a reformer. The reformer converts the hydrocarbon gas to hydrogen and carbon monoxide for combustion in a combustor. A portion of the steam from the heat recovery steam generator is used to cool components, such as the stationary vanes, in the turbine section, thereby superheating the steam. The superheated steam is mixed into the hydrocarbon gas upstream of the reformer, thereby eliminating the need to raise the temperature of the expanded gas discharged from the turbine section in order to achieve effective conversion of the hydrocarbon gas. 4 figs.

  20. Cooling supply system for stage 3 bucket of a gas turbine

    DOEpatents

    Eldrid, Sacheverel Quentin; Burns, James Lee; Palmer, Gene David; Leone, Sal Albert; Drlik, Gary Joseph; Gibler, Edward Eugene

    2002-01-01

    In a land based gas turbine including a compressor, a combustor and turbine section including at least three stages, an improvement comprising an inlet into a third stage nozzle from the compressor for feeding cooling air from the compressor to the third stage nozzle; at least one passageway running substantially radially through each airfoil of the third stage nozzle and an associated diaphragm, into an annular space between the rotor and the diaphragm; and passageways communicating between the annular space and individual buckets of the third stage.

  1. Hot streaks and phantom cooling in a turbine rotor passage. II - Combined effects and analytical modelling

    NASA Technical Reports Server (NTRS)

    Roback, Richard J.; Dring, Robert P.

    1992-01-01

    Experimental documentation and analytical correlations demonstrating the effects of hot streak accumulation and phantom cooling on turbine rotor airfoil surface temperature are presented. Results are shown which quantify the impact of a nonuniform temperature profile at the entrance of a turbine due to combustor-generated hot and cold streaks, and cooling air discharged from the trailing edge of the upstream stator. Experimental results are shown for a range of controlling variables to identify where streak accumulation and phantom cooling were most likely to be strongest. These variables include streak-to-free stream density ratio, streak injection location, and coolant-to-free stream density and velocity ratios. Experimental results are shown for the combined effects of hot streak and stator coolant on the adiabatic recovery temperature of the rotor.

  2. Heat transfer in cooled guide vanes. [of radial inflow turbine

    NASA Technical Reports Server (NTRS)

    Tabakoff, W.; Kotwal, R.; Hamed, A.

    1977-01-01

    A numerical study to determine the temperature distribution in the guide vanes of a radial inflow turbine is presented. A computer program has been developed to calculate the temperature distribution when the vanes are cooled internally using a combination of impingement and film cooling techniques. The study is based on the use of the finite difference method in a two dimensional heat conduction problem. The results are then compared to determine the best cooling configuration for a certain coolant to primary mass flow ratio.

  3. Contribution of heat transfer to turbine blades and vanes for high temperature industrial gas turbines. Part 1: Film cooling.

    PubMed

    Takeishi, K; Aoki, S

    2001-05-01

    This paper deals with the contribution of heat transfer to increase the turbine inlet temperature of industrial gas turbines in order to attain efficient and environmentally benign engines. High efficiency film cooling, in the form of shaped film cooling and full coverage film cooling, is one of the most important cooling technologies. Corresponding heat transfer tests to optimize the film cooling effectiveness are shown and discussed in this first part of the contribution. PMID:11460641

  4. Cooling system for a gas turbine

    DOEpatents

    Wilson, Ian David; Salamah, Samir Armando; Bylina, Noel Jacob

    2003-01-01

    A plurality of arcuate circumferentially spaced supply and return manifold segments are arranged on the rim of a rotor for respectively receiving and distributing cooling steam through exit ports for distribution to first and second-stage buckets and receiving spent cooling steam from the first and second-stage buckets through inlet ports for transmission to axially extending return passages. Each of the supply and return manifold segments has a retention system for precluding substantial axial, radial and circumferential displacement relative to the rotor. The segments also include guide vanes for minimizing pressure losses in the supply and return of the cooling steam. The segments lie substantially equal distances from the centerline of the rotor and crossover tubes extend through each of the segments for communicating steam between the axially adjacent buckets of the first and second stages, respectively.

  5. Design Concepts for Cooled Ceramic Matrix Composite Turbine Vanes

    NASA Technical Reports Server (NTRS)

    Boyle, Robert

    2014-01-01

    This project demonstrated that higher temperature capabilities of ceramic matrix composites (CMCs) can be used to reduce emissions and improve fuel consumption in gas turbine engines. The work involved closely coupling aerothermal and structural analyses for the first-stage vane of a high-pressure turbine (HPT). These vanes are actively cooled, typically using film cooling. Ceramic materials have structural and thermal properties different from conventional metals used for the first-stage HPT vane. This project identified vane configurations that satisfy CMC structural strength and life constraints while maintaining vane aerodynamic efficiency and reducing vane cooling to improve engine performance and reduce emissions. The project examined modifications to vane internal configurations to achieve the desired objectives. Thermal and pressure stresses are equally important, and both were analyzed using an ANSYS® structural analysis. Three-dimensional fluid and heat transfer analyses were used to determine vane aerodynamic performance and heat load distributions.

  6. Ducting arrangement for cooling a gas turbine structure

    DOEpatents

    Lee, Ching-Pang; Morrison, Jay A.

    2015-07-21

    A ducting arrangement (10) for a can annular gas turbine engine, including: a duct (12, 14) disposed between a combustor (16) and a first row of turbine blades and defining a hot gas path (30) therein, the duct (12, 14) having raised geometric features (54) incorporated into an outer surface (80); and a flow sleeve (72) defining a cooling flow path (84) between an inner surface (78) of the flow sleeve (72) and the duct outer surface (80). After a cooling fluid (86) traverses a relatively upstream raised geometric feature (90), the inner surface (78) of the flow sleeve (72) is effective to direct the cooling fluid (86) toward a landing (94) separating the relatively upstream raised geometric feature (90) from a relatively downstream raised geometric feature (94).

  7. Wavy flow cooling concept for turbine airfoils

    DOEpatents

    Liang, George

    2010-08-31

    An airfoil including an outer wall and a cooling cavity formed therein. The cooling cavity includes a leading edge flow channel located adjacent a leading edge of the airfoil and a trailing edge flow channel located adjacent a trailing edge of the airfoil. Each of the leading edge and trailing edge flow channels define respective first and second flow axes located between pressure and suction sides of the airfoil. A plurality of rib members are located within each of the flow channels, spaced along the flow axes, and alternately extending from opposing sides of the flow channels to define undulating flow paths through the flow channels.

  8. The effect of rotation on heat transfer in the radial cooling channels of turbine blades

    NASA Astrophysics Data System (ADS)

    Iskakov, K. M.; Trushin, V. A.

    1985-02-01

    The effect of rotation on heat transfer in the channels of moving turbine blades in a loop cooling system is investigated experimentally. The working channels consisted of round tubes with sharp edges and the tubes were fixed to a support. Calculation of the parameters required for correlating the experimental data was conducted according to local air temperature at the entry of the channel. Analysis of the measured and calculated heat transfer parameters showed that the average error in determining heat transfer was 13 percent. The error in calculating the bulk flow rate of air was 8 percent. Formulas for calculating the centrifugal and centripetal air flows are derived.

  9. Stirling Air Conditioner for Compact Cooling

    SciTech Connect

    2010-09-01

    BEETIT Project: Infinia is developing a compact air conditioner that uses an unconventional high efficient Stirling cycle system (vs. conventional vapor compression systems) to produce cool air that is energy efficient and does not rely on polluting refrigerants. The Stirling cycle system is a type of air conditioning system that uses a motor with a piston to remove heat to the outside atmosphere using a gas refrigerant. To date, Stirling systems have been expensive and have not had the right kind of heat exchanger to help cool air efficiently. Infinia is using chip cooling technology from the computer industry to make improvements to the heat exchanger and improve system performance. Infinia’s air conditioner uses helium gas as refrigerant, an environmentally benign gas that does not react with other chemicals and does not burn. Infinia’s improvements to the Stirling cycle system will enable the cost-effective mass production of high-efficiency air conditioners that use no polluting refrigerants.

  10. Design Concepts for Cooled Ceramic Composite Turbine Vane

    NASA Technical Reports Server (NTRS)

    Boyle, Robert J.; Parikh, Ankur H.; Nagpal, VInod K.

    2015-01-01

    The objective of this work was to develop design concepts for a cooled ceramic vane to be used in the first stage of the High Pressure Turbine(HPT). To insure that the design concepts were relevant to the gas turbine industry needs, Honeywell International Inc. was subcontracted to provide technical guidance for this work. The work performed under this contract can be divided into three broad categories. The first was an analysis of the cycle benefits arising from the higher temperature capability of Ceramic Matrix Composite(CMC) compared with conventional metallic vane materials. The second category was a series of structural analyses for variations in the internal configuration of first stage vane for the High Pressure Turbine(HPT) of a CF6 class commercial airline engine. The third category was analysis for a radial cooled turbine vanes for use in turboshaft engine applications. The size, shape and internal configuration of the turboshaft engine vanes were selected to investigate a cooling concept appropriate to small CMC vanes.

  11. Achieving better cooling of turbine blades using numerical simulation methods

    NASA Astrophysics Data System (ADS)

    Inozemtsev, A. A.; Tikhonov, A. S.; Sendyurev, C. I.; Samokhvalov, N. Yu.

    2013-02-01

    A new design of the first-stage nozzle vane for the turbine of a prospective gas-turbine engine is considered. The blade's thermal state is numerically simulated in conjugate statement using the ANSYS CFX 13.0 software package. Critical locations in the blade design are determined from the distribution of heat fluxes, and measures aimed at achieving more efficient cooling are analyzed. Essentially lower (by 50-100°C) maximal temperature of metal has been achieved owing to the results of the performed work.

  12. Compound cooling flow turbulator for turbine component

    DOEpatents

    Lee, Ching-Pang; Jiang, Nan; Marra, John J; Rudolph, Ronald J

    2014-11-25

    Multi-scale turbulation features, including first turbulators (46, 48) on a cooling surface (44), and smaller turbulators (52, 54, 58, 62) on the first turbulators. The first turbulators may be formed between larger turbulators (50). The first turbulators may be alternating ridges (46) and valleys (48). The smaller turbulators may be concave surface features such as dimples (62) and grooves (54), and/or convex surface features such as bumps (58) and smaller ridges (52). An embodiment with convex turbulators (52, 58) in the valleys (48) and concave turbulators (54, 62) on the ridges (46) increases the cooling surface area, reduces boundary layer separation, avoids coolant shadowing and stagnation, and reduces component mass.

  13. Pulsed film cooling on a turbine blade leading edge

    NASA Astrophysics Data System (ADS)

    Rutledge, James L.

    2009-12-01

    Unsteadiness in gas turbine film cooling jets may arise due to inherent unsteadiness of the flow through an engine or may be induced as a means of flow control. The traditional technique used to evaluate the performance of a steady film cooling scheme is demonstrated to be insufficient for use with unsteady film cooling and is modified to account for the cross coupling of the time dependent adiabatic effectiveness and heat transfer coefficient. The addition of a single term to the traditional steady form of the net heat flux reduction equation with time averaged quantities accounts for the unsteady effects. An experimental technique to account for the influence of the new term was devised and used to measure the influence of a pulsating jet on the net heat flux in the leading edge region of a turbine blade. High spatial resolution data was acquired in the near-hole region using infrared thermography coupled with experimental techniques that allowed application of the appropriate thermal boundary conditions immediately adjacent to the film cooling hole. The turbine blade leading edge was simulated by a half cylinder in cross flow with a blunt afterbody. The film cooling geometry consisted of a coolant hole located 21.5° from the leading edge, angled 20° to the surface and 90° from the streamwise direction. Investigated parameters include pulsation frequency, duty cycle, and waveform shape. Separate experiments were conducted in a water channel to provide visualization of the unsteady coolant propagation behavior. Further insight into the flow physics was obtained through computational simulations of the experimental apparatus. The computational results afforded time resolved flow field and net heat flux reduction data unobtainable with the experimental techniques. A technique to predict the performance of an unsteady film cooling scheme through knowledge of only the steady film cooling behavior was developed and demonstrated to be effective.

  14. Gas turbine row #1 steam cooled vane

    DOEpatents

    Cunha, Frank J.

    2000-01-01

    A design for a vane segment having a closed-loop steam cooling system is provided. The vane segment comprises an outer shroud, an inner shroud and an airfoil, each component having a target surface on the inside surface of its walls. A plurality of rectangular waffle structures are provided on the target surface to enhance heat transfer between each component and cooling steam. Channel systems are provided in the shrouds to improve the flow of steam through the shrouds. Insert legs located in cavities in the airfoil are also provided. Each insert leg comprises outer channels located on a perimeter of the leg, each outer channel having an outer wall and impingement holes on the outer wall for producing impingement jets of cooling steam to contact the airfoil's target surface. Each insert leg further comprises a plurality of substantially rectangular-shaped ribs located on the outer wall and a plurality of openings located between outer channels of the leg to minimize cross flow degradation.

  15. Turbine airfoil with an internal cooling system having vortex forming turbulators

    DOEpatents

    Lee, Ching-Pang

    2014-12-30

    A turbine airfoil usable in a turbine engine and having at least one cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels having a plurality of turbulators protruding from an inner surface and positioned generally nonorthogonal and nonparallel to a longitudinal axis of the airfoil cooling channel. The configuration of turbulators may create a higher internal convective cooling potential for the blade cooling passage, thereby generating a high rate of internal convective heat transfer and attendant improvement in overall cooling performance. This translates into a reduction in cooling fluid demand and better turbine performance.

  16. Cooling circuit for a gas turbine bucket and tip shroud

    SciTech Connect

    Willett, Fred Thomas

    2004-07-13

    An open cooling circuit for a gas turbine airfoil and associated tip shroud includes a first group of cooling holes internal to the airfoil and extending in a radially outward direction generally along a leading edge of the airfoil; a second group of cooling holes internal to the airfoil and extending in a radially outward direction generally along a trailing edge of the airfoil. A common plenum is formed in the tip shroud in direct communication with the first and second group of cooling holes, but a second plenum may be provided for the second group of radial holes. A plurality of exhaust holes extends from the plenum(s), through the tip shroud and opening along a peripheral edge of the tip shroud.

  17. Effects of rotation on impingement cooling of turbine blades

    NASA Technical Reports Server (NTRS)

    Kreatsoulas, J. C.; Kerrebrock, J. L.; Epstein, A. H.; Rogo, C.

    1985-01-01

    The effects of rotation on impingement cooling of turbine blades were studied experimentally as a specialized facility at M.I.T. A foil heated resistively was cooled by a jet flow on one side and temperature monitored on the other. Rotating the blade limits the heat transfer path to conduction through the support structure and radiation. IR radiometry furnishes the temperature distributions on the chamber wall, permitting the internal heat transfer coefficient to be measured. The heat transfer efficiency has been found to fall as much as 30 percent as rotational speed increases. The conditions observed confirm the significance of rotational effects, particularly with regard to potential early blade failure.

  18. Experimental Evaluation of a Cooled Radial-inflow Turbine

    NASA Technical Reports Server (NTRS)

    Tirres, Lizet; Dicicco, L. Danielle; Nowlin, Brent C.

    1993-01-01

    Two 14.4 inch tip diameter rotors were installed and tested in the Small Engines Component Turbine Facility (SECTF) at NASA Lewis Research Center. The rotors, a solid and a cooled version of a radial-inflow turbine, were tested with a 15 vane stat or over a set of rotational speeds ranging from 80 to 120 percent design speed (17,500 to 21,500 rpm). The total-to-total stage pressure ratios ranged from 2.5 to 5.5. The data obtained at the equivalent conditions using the solid version of the rotor are presented with the cooled rotor data. A Reynolds number of 381,000 was maintained for both rotors, whose stages had a design mass flow of 4.0 Ibm/sec, a design work level of 59.61 Btu/lbm, and a design efficiency of 87 percent. The results include mass flow data, turbine torque, turbine exit flow angles, stage efficiency, and rotor inlet and exit surveys.

  19. Experimental evaluation of a cooled radial-inflow turbine

    NASA Technical Reports Server (NTRS)

    Tirres, Lizet; Dicicco, L. D.; Nowlin, Brent C.

    1993-01-01

    Two 14.4 inch tip diameter rotors were installed and tested in the Small Engines Component Turbine Facility (SECTF) at NASA Lewis Research Center. The rotors, a solid and a cooled version of a radial-inflow turbine, were tested with a 15 vane stat or over a set of rotational speeds ranging from 80 to 120 percent design speed (17,500 to 21,500 rpm). The total-to-total stage pressure ratios ranged from 2.5 to 5.5. The data obtained at the equivalent conditions using the solid version of the rotor are presented with the cooled rotor data. A Reynolds number of 381,000 was maintained for both rotors, whose stages had a design mass flow of 4.0 lbm/sec, a design work level of 59.61 Btu/lbm, and a design efficiency of 87 percent. The results include mass flow data, turbine torque, turbine exit flow angles, stage efficiency, and rotor inlet and exit surveys.

  20. Advanced turbine cooling, heat transfer, and aerodynamic studies

    SciTech Connect

    Je-Chin Han; Schobeiri, M.T.

    1995-10-01

    The contractual work is in three parts: Part I - Effect of rotation on enhanced cooling passage heat transfer, Part II - Effect on Thermal Barrier Coating (TBC) spallation on surface heat transfer, and Part III - Effect of surface roughness and trailing edge ejection on turbine efficiency under unsteady flow conditions. Each section of this paper has been divided into three parts to individually accommodate each part. Part III is further divided into Parts IIIa and IIIb.

  1. An evaluation of thermal energy storage options for precooling gas turbine inlet air

    SciTech Connect

    Antoniak, Z.I.; Brown, D.R.; Drost, M.K.

    1992-12-01

    Several approaches have been used to reduce the temperature of gas turbine inlet air. One of the most successful uses off-peak electric power to drive vapor-compression-cycle ice makers. The ice is stored until the next time high ambient temperature is encountered, when the ice is used in a heat exchanger to cool the gas turbine inlet air. An alternative concept would use seasonal thermal energy storage to store winter chill for inlet air cooling. The objective of this study was to compare the performance and economics of seasonal thermal energy storage in aquifers with diurnal ice thermal energy storage for gas turbine inlet air cooling. The investigation consisted of developing computer codes to model the performance of a gas turbine, energy storage system, heat exchangers, and ancillary equipment. The performance models were combined with cost models to calculate unit capital costs and levelized energy costs for each concept. The levelized energy cost was calculated for three technologies in two locations (Minneapolis, Minnesota and Birmingham, Alabama). Precooling gas turbine inlet air with cold water supplied by an aquifer thermal energy storage system provided lower cost electricity than simply increasing the size of the turbine for meteorological and geological conditions existing in the Minneapolis vicinity. A 15 to 20% cost reduction resulted for both 0.05 and 0.2 annual operating factors. In contrast, ice storage precooling was found to be between 5 and 20% more expensive than larger gas turbines for the Minneapolis location. In Birmingham, aquifer thermal energy storage precooling was preferred at the higher capacity factor and ice storage precooling was the best option at the lower capacity factor. In both cases, the levelized cost was reduced by approximately 5% when compared to larger gas turbines.

  2. Cooling circuit for and method of cooling a gas turbine bucket

    DOEpatents

    Jacala, Ariel C. P.

    2002-01-01

    A closed internal cooling circuit for a gas turbine bucket includes axial supply and return passages in the dovetail of the bucket. A first radial outward supply passage provides cooling medium to and along a passageway adjacent the leading edge and then through serpentine arranged passageways within the airfoil to a chamber adjacent the airfoil tip. A second radial passage crosses over the radial return passage for supplying cooling medium to and along a pair of passageways along the trailing edge of the airfoil section. The last passageway of the serpentine passageways and the pair of passageways communicate one with the other in the chamber for returning spent cooling medium radially inwardly along divided return passageways to the return passage. In this manner, both the leading and trailing edges are cooled using the highest pressure, lowest temperature cooling medium.

  3. Investigation on cooling effectiveness and aerodynamic loss of a turbine cascade with film cooling

    NASA Astrophysics Data System (ADS)

    Liu, Jianjun; Lin, Xiaochun; Zhang, Xiaodong; An, Baitao

    2016-02-01

    This paper describes the numerical study on film cooling effectiveness and aerodynamic loss due to coolant and main stream mixing for a turbine guide vane. The effects of blowing ratio, mainstream Mach number, surface curvature on the cooling effectiveness and mixing loss were studied and discussed. The numerical results show that the distributions of film cooling effectiveness on the suction surface and pressure surface at the same blowing ratio (BR) are different due to local surface curvature and pressure gradient. The aerodynamic loss features for film holes on the pressure surface are also different from film holes on the suction surface.

  4. Bore tube assembly for steam cooling a turbine rotor

    SciTech Connect

    DeStefano, Thomas Daniel; Wilson, Ian David

    2002-01-01

    An axial bore tube assembly for a turbine is provided to supply cooling steam to hot gas components of the turbine wheels and return the spent cooling steam. A pair of inner and outer tubes define a steam supply passage concentric about an inner return passage. The forward ends of the tubes communicate with an end cap assembly having sets of peripheral holes communicating with first and second sets of radial tubes whereby cooling steam from the concentric passage is supplied through the end cap holes to radial tubes for cooling the buckets and return steam from the buckets is provided through the second set of radial tubes through a second set of openings of the end cap into the coaxial return passage. A radial-to-axial flow transitioning device, including anti-swirling vanes is provided in the end cap. A strut ring adjacent the aft end of the bore tube assembly permits axial and radial thermal expansion of the inner tube relative to the outer tube.

  5. Pneumomediastinum secondary to use of a high speed air turbine drill during a dental extraction.

    PubMed Central

    Torres-Melero, J.; Arias-Diaz, J.; Balibrea, J. L.

    1996-01-01

    Pneumomediastinum and subcutaneous emphysema of the neck and thorax can occur exceptionally following a dental procedure. A case is described of acute subcutaneous emphysema of the lateral region of the neck and thorax associated with pneumomediastinum during a dental extraction with an air and water cooled turbine burn drill. PMID:8779147

  6. A Blast of Cool Air

    NASA Technical Reports Server (NTRS)

    2000-01-01

    Unable to solve their engineering problem with a rotor in their Orbital Vane product, DynEco Corporation turned to Kennedy Space Center for help. KSC engineers determined that the compressor rotor was causing a large concentration of stress, which led to cracking and instant rotor failure. NASA redesigned the lubrication system, which allowed the company to move forward with its compressor that has no rubbing parts. The Orbital Vane is a refrigerant compressor suitable for mobile air conditioning and refrigeration.

  7. Near wall cooling for a highly tapered turbine blade

    DOEpatents

    Liang, George

    2011-03-08

    A turbine blade having a pressure sidewall and a suction sidewall connected at chordally spaced leading and trailing edges to define a cooling cavity. Pressure and suction side inner walls extend radially within the cooling cavity and define pressure and suction side near wall chambers. A plurality of mid-chord channels extend radially from a radially intermediate location on the blade to a tip passage at the blade tip for connecting the pressure side and suction side near wall chambers in fluid communication with the tip passage. In addition, radially extending leading edge and trailing edge flow channels are located adjacent to the leading and trailing edges, respectively, and cooling fluid flows in a triple-pass serpentine path as it flows through the leading edge flow channel, the near wall chambers and the trailing edge flow channel.

  8. Liquid metal reactor air cooling baffle

    DOEpatents

    Hunsbedt, A.

    1994-08-16

    A baffle is provided between a relatively hot containment vessel and a relatively cold silo for enhancing air cooling performance. The baffle includes a perforate inner wall positionable outside the containment vessel to define an inner flow riser therebetween, and an imperforate outer wall positionable outside the inner wall to define an outer flow riser therebetween. Apertures in the inner wall allow thermal radiation to pass laterally therethrough to the outer wall, with cooling air flowing upwardly through the inner and outer risers for removing heat. 3 figs.

  9. Offshore Floating Wind Turbine-driven Deep Sea Water Pumping for Combined Electrical Power and District Cooling

    NASA Astrophysics Data System (ADS)

    Sant, T.; Buhagiar, D.; Farrugia, R. N.

    2014-06-01

    A new concept utilising floating wind turbines to exploit the low temperatures of deep sea water for space cooling in buildings is presented. The approach is based on offshore hydraulic wind turbines pumping pressurised deep sea water to a centralised plant consisting of a hydro-electric power system coupled to a large-scale sea water-cooled air conditioning (AC) unit of an urban district cooling network. In order to investigate the potential advantages of this new concept over conventional technologies, a simplified model for performance simulation of a vapour compression AC unit was applied independently to three different systems, with the AC unit operating with (1) a constant flow of sea surface water, (2) a constant flow of sea water consisting of a mixture of surface sea water and deep sea water delivered by a single offshore hydraulic wind turbine and (3) an intermittent flow of deep sea water pumped by a single offshore hydraulic wind turbine. The analysis was based on one year of wind and ambient temperature data for the Central Mediterranean that is known for its deep waters, warm climate and relatively low wind speeds. The study confirmed that while the present concept is less efficient than conventional turbines utilising grid-connected electrical generators, a significant portion of the losses associated with the hydraulic transmission through the pipeline are offset by the extraction of cool deep sea water which reduces the electricity consumption of urban air-conditioning units.

  10. Development of Air-cooled Engines with Blower Cooling

    NASA Technical Reports Server (NTRS)

    Lohner, Kurt

    1933-01-01

    With the aid of a heating device, the heat transfer to cylinders with conical fins of various forms is determined both for shrouded and exposed cylinders. Simultaneously the pressure drop for overcoming the resistance to the motion of air between the fins of the enclosed cylinder is measured. Thus the relations between the heat transfer and the energy required for cooling are discovered. The investigations show that the heat transfer in a conducted air flow is much greater than in a free current and that further improvement, as compared with free exposure, is possible through narrower spaces between the fins.

  11. 14 CFR 23.1045 - Cooling test procedures for turbine engine powered airplanes.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Cooling test procedures for turbine engine powered airplanes. 23.1045 Section 23.1045 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION... CATEGORY AIRPLANES Powerplant Cooling § 23.1045 Cooling test procedures for turbine engine...

  12. 14 CFR 23.1045 - Cooling test procedures for turbine engine powered airplanes.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Cooling test procedures for turbine engine powered airplanes. 23.1045 Section 23.1045 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION... CATEGORY AIRPLANES Powerplant Cooling § 23.1045 Cooling test procedures for turbine engine...

  13. 14 CFR 23.1045 - Cooling test procedures for turbine engine powered airplanes.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Cooling test procedures for turbine engine powered airplanes. 23.1045 Section 23.1045 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION... CATEGORY AIRPLANES Powerplant Cooling § 23.1045 Cooling test procedures for turbine engine...

  14. 14 CFR 23.1045 - Cooling test procedures for turbine engine powered airplanes.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Cooling test procedures for turbine engine powered airplanes. 23.1045 Section 23.1045 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION... CATEGORY AIRPLANES Powerplant Cooling § 23.1045 Cooling test procedures for turbine engine...

  15. 14 CFR 23.1045 - Cooling test procedures for turbine engine powered airplanes.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Cooling test procedures for turbine engine powered airplanes. 23.1045 Section 23.1045 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION... CATEGORY AIRPLANES Powerplant Cooling § 23.1045 Cooling test procedures for turbine engine...

  16. Advanced cooled-engine shell/spar turbine vanes and blades. Final report

    SciTech Connect

    Not Available

    1986-08-01

    The objectives of the Advanced Cooling Full-Scale Engine Demonstration Program, Phase II (EPRI Contract RP1319-5), were to develop and to demonstrate an advanced aircraft cooling technology for use in the vanes, blades, and associated hot-section components of a utility-sized combustion turbine. Use of such technology would provide reduced metal-surface temperatures for improved reliability or the potential for increasing turbine inlet temperatures for improved thermal efficiency. In the shell/spar cooling technology chosen for the vane and blade construction, cooling air flows through channels between a thin metal external sheet (shell) and a hollow-cast internal support member (spar). The shell and spar are joined by diffusion bonding. The results of first-stage shell/spar blade and vane design studies are reported, and heat transfer and stress analyses of the blade and vane designs are featured. The progress made on the development of the vane fabrication technology, up to the early termination at the end of 1984, is fully delineated. The successful development of an ultrasonic inspection technique to indicate unbonded areas between the shell and spar is reported. The results of heat transfer testing with shell/spar specimens and low-cycle fatigue testing of IN617 sheet are described. Problem areas in the determination of the low-cycle fatigue life expectancy of the designs are identified. Recommendations are given for continuing the blade and vane shell/spar advanced cooling technology development.

  17. Comparison of effectiveness of convection-, transpiration-, and film-cooling methods with air as coolant

    NASA Technical Reports Server (NTRS)

    Eckert, E R G; Livingood, N B

    1954-01-01

    Various parts of aircraft propulsion engines that are in contact with hot gases often require cooling. Transpiration and film cooling, new methods that supposedly utilize cooling air more effectively than conventional convection cooling, have already been proposed. This report presents material necessary for a comparison of the cooling requirements of these three methods. Correlations that are regarded by the authors as the most reliable today are employed in evaluating each of the cooling processes. Calculations for the special case in which the gas velocity is constant along the cooled wall (flat plate) are presented. The calculations reveal that a comparison of the three cooling processes can be made on quite a general basis. The superiority of transpiration cooling is clearly shown for both laminar and turbulent flow. This superiority is reduced when the effects of radiation are included; for gas-turbine blades, however, there is evidence indicating that radiation may be neglected.

  18. Cooled variable-area radial turbine technology program

    NASA Technical Reports Server (NTRS)

    Large, G. D.; Meyer, L. J.

    1982-01-01

    The objective of this study was a conceptual evaluation and design analyses of a cooled variable-area radial turbine capable of maintaining nearly constant high efficiency when operated at a constant speed and pressure ratio over a range of flows corresponding to 50- to 100-percent maximum engine power. The results showed that a 1589K (2400 F) turbine was feasible that would satisfy a 4000-hour duty cycle life goal. The final design feasibility is based on 1988 material technology goals. A peak aerodynamic stage total efficiency of 0.88 was predicted at 100 percent power. Two candidate stators were identified: an articulated trailing-edge and a locally movable sidewall. Both concepts must be experimentally evaluated to determine the optimum configuration. A follow-on test program is proposed for this evaluation.

  19. Heat Transfer Measurements for a Film Cooled Turbine Vane Cascade

    NASA Technical Reports Server (NTRS)

    Poinsatte, Philip E.; Heidmann, James D.; Thurman, Douglas R.

    2008-01-01

    Experimental heat transfer and pressure measurements were obtained on a large scale film cooled turbine vane cascade. The objective was to investigate heat transfer on a commercial high pressure first stage turbine vane at near engine Mach and Reynolds number conditions. Additionally blowing ratios and coolant density were also matched. Numerical computations were made with the Glenn-HT code of the same geometry and compared with the experimental results. A transient thermochromic liquid crystal technique was used to obtain steady state heat transfer data on the mid-span geometry of an instrumented vane with 12 rows of circular and shaped film cooling holes. A mixture of SF6 and Argon gases was used for film coolant to match the coolant-to-gas density ratio of a real engine. The exit Mach number and Reynolds number were 0.725 and 2.7 million respectively. Trends from the experimental heat transfer data matched well with the computational prediction, particularly for the film cooled case.

  20. Rotating diffuser for pressure recovery in a steam cooling circuit of a gas turbine

    SciTech Connect

    Eldrid, Sacheverel Q.; Salamah, Samir A.; DeStefano, Thomas Daniel

    2002-01-01

    The buckets of a gas turbine are steam-cooled via a bore tube assembly having concentric supply and spent cooling steam return passages rotating with the rotor. A diffuser is provided in the return passage to reduce the pressure drop. In a combined cycle system, the spent return cooling steam with reduced pressure drop is combined with reheat steam from a heat recovery steam generator for flow to the intermediate pressure turbine. The exhaust steam from the high pressure turbine of the combined cycle unit supplies cooling steam to the supply conduit of the gas turbine.

  1. Prediction of Film Cooling on Gas Turbine Airfoils

    NASA Technical Reports Server (NTRS)

    Garg, Vijay K.; Gaugler, Raymond E.

    1994-01-01

    A three-dimensional Navier-Stokes analysis tool has been developed in order to study the effect of film cooling on the flow and heat transfer characteristics of actual turbine airfoils. An existing code (Arnone et al., 1991) has been modified for the purpose. The code is an explicit, multigrid, cell-centered, finite volume code with an algebraic turbulence model. Eigenvalue scaled artificial dissipation and variable-coefficient implicit residual smoothing are used with a full-multigrid technique. Moreover, Mayle's transition criterion (Mayle, 1991) is used. The effects of film cooling have been incorporated into the code in the form of appropriate boundary conditions at the hole locations on the airfoil surface. Each hole exit is represented by several control volumes, thus providing an ability to study the effect of hole shape on the film-cooling characteristics. Comparison is fair with near mid-span experimental data for four and nine rows of cooling holes, five on the shower head, and two rows each on the pressure and suction surfaces. The computations, however, show a strong spanwise variation of the heat transfer coefficient on the airfoil surface, specially with shower-head cooling.

  2. Cooling of Gas Turbines. 3; Analysis of Rotor and Blade Temperatures in Liquid-Cooled Gas Turbines

    NASA Technical Reports Server (NTRS)

    Brown, W. Byron; Livingood, John N. B.

    1947-01-01

    A theoretical analysis of the radial temperature distribution through the rotor and constant cross sectional area blades near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the rotor and blade temperatures of a specific turbine using a gas flow of 55 pounds per second, a coolant flow of 6.42 pounds per second, and an average coolant temperature of 200 degrees F. The effect of using kerosene, water, and ethylene glycol was determined. The effect of varying blade length and coolant passage lengths with water as the coolant was also determined. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.

  3. Method and system for providing cooling for turbine components

    DOEpatents

    Morgan, Victor John; Lacy, Benjamin Paul

    2016-08-16

    A system for providing cooling for a turbine component that includes an outer surface exposed to combustion gases is provided. A component base includes at least one fluid supply passage coupleable to a source of cooling fluid. At least one feed passage communicates with the at least one fluid supply passage. At least one delivery channel communicates with the at least one feed passage. At least one cover layer covers the at least one feed passage and the at least one delivery channel, defining at least in part the component outer surface. At least one discharge passage extends to the outer surface. A diffuser section is defined in at least one of the at least one delivery channel and the at least one discharge passage, such that a fluid channeled through the system is diffused prior to discharge adjacent the outer surface.

  4. Oil cooling system for a gas turbine engine

    NASA Technical Reports Server (NTRS)

    Coffinberry, G. A.; Kast, H. B. (Inventor)

    1977-01-01

    A gas turbine engine fuel delivery and control system is provided with means to recirculate all fuel in excess of fuel control requirements back to aircraft fuel tank, thereby increasing the fuel pump heat sink and decreasing the pump temperature rise without the addition of valving other than that normally employed. A fuel/oil heat exchanger and associated circuitry is provided to maintain the hot engine oil in heat exchange relationship with the cool engine fuel. Where anti-icing of the fuel filter is required, means are provided to maintain the fuel temperature entering the filter at or above a minimum level to prevent freezing thereof. Fluid circuitry is provided to route hot engine oil through a plurality of heat exchangers disposed within the system to provide for selective cooling of the oil.

  5. Features of steam turbine cooling by the example of an SKR-100 turbine for supercritical steam parameters

    NASA Astrophysics Data System (ADS)

    Arkadyev, B. A.

    2015-10-01

    Basic principles of cooling of high-temperature steam turbines and constructive solutions used for development of the world's first cooled steam turbine SKR-100 (R-100-300) are described. Principal differences between the thermodynamic properties of cooling medium in the steam and gas turbines and the preference of making flow passes of cooled cylinders of steam turbines as reactive are shown. Some of its operation results and their conclusions are given. This turbine with a power of 100 MW, initial steam parameters approximately 30 MPa and 650°C, and back pressure 3 MPa was made by a Kharkov turbine plant in 1961 and ran successfully at a Kashira GRES (state district power plant) up to 1979, when it was taken out of use in a still fully operating condition. For comparison, some data on construction features and operation results of the super-high pressure cylinder of steam turbines of American Philo 6 (made by General Electric Co.) and Eddystone 1 (made by Westinghouse Co.) power generating units, which are close to the SKR-100 turbine by design initial steam parameters and the implementation time, are given. The high operational reliability and effectiveness of the cooling system that was used in the super-high pressure cylinder of the SKR-100 turbine of the power-generating unit, which were demonstrated in operation, confirms rightfulness and expediency of principles and constructive solutions laid at its development. As process steam temperatures are increased, the realization of the proposed approach to cooling of multistage turbines makes it possible to limit for large turbine parts the application of new, more expensive high-temperature materials, which are required for making steam boilers, and, in some cases, to do completely away with their utilization.

  6. Feasibility of Actively Cooled Silicon Nitride Airfoil for Turbine Applications Demonstrated

    NASA Technical Reports Server (NTRS)

    Bhatt, Ramakrishna T.

    2001-01-01

    Nickel-base superalloys currently limit gas turbine engine performance. Active cooling has extended the temperature range of service of nickel-base superalloys in current gas turbine engines, but the margin for further improvement appears modest. Therefore, significant advancements in materials technology are needed to raise turbine inlet temperatures above 2400 F to increase engine specific thrust and operating efficiency. Because of their low density and high-temperature strength and thermal conductivity, in situ toughened silicon nitride ceramics have received a great deal of attention for cooled structures. However, the high processing costs and low impact resistance of silicon nitride ceramics have proven to be major obstacles for widespread applications. Advanced rapid prototyping technology in combination with conventional gel casting and sintering can reduce high processing costs and may offer an affordable manufacturing approach. Researchers at the NASA Glenn Research Center, in cooperation with a local university and an aerospace company, are developing actively cooled and functionally graded ceramic structures. The objective of this program is to develop cost-effective manufacturing technology and experimental and analytical capabilities for environmentally stable, aerodynamically efficient, foreign-object-damage-resistant, in situ toughened silicon nitride turbine nozzle vanes, and to test these vanes under simulated engine conditions. Starting with computer aided design (CAD) files of an airfoil and a flat plate with internal cooling passages, the permanent and removable mold components for gel casting ceramic slips were made by stereolithography and Sanders machines, respectively. The gel-cast part was dried and sintered to final shape. Several in situ toughened silicon nitride generic airfoils with internal cooling passages have been fabricated. The uncoated and thermal barrier coated airfoils and flat plates were burner rig tested for 30 min without

  7. Air-cooled overhead-valve engine

    SciTech Connect

    Shirai, T.

    1987-06-16

    This patent describes an air-cooled overhead-valve internal combustion engine. The engine is composed of a crankcase with a crankshaft, a cylinder block with a cylinder head and a combustion chamber mounted in the crankcase. At least a pair of intake and exhaust valves installed in intake and exhaust ports are formed in the cylinder head. A valve drive system mounted adjacent to the cylinder block drives the intake and exhaust valves through cam-driven push rods. An intake pipe is connected at one end of the intake port and at its opposite end to an air cleaner and a carburetor. An exhaust duct is connected at one end of the exhaust port. A flywheel is joined to the crankshaft at the other end of the output side end of the crankshaft and a cooling fan mounted on the flywheel. The improvements are where the cooling fan is housed, together with the crankcase and flywheel, in a fan casing having a pair of inlet and outlet openings bored in opposite walls. The inlet opening is located at the flywheel side of the crankshaft, while the outlet opening is located at the opposite side of the crankshaft from the flywheel. The cam-driven push rods are located in the crankcase on that side of the cylinder block far remote from where the intake pipe is connected to the intake port. The cooling fan is mounted in the fan casing in such a manner that the cooling air from the cooling fan is allowed to flow in a direction substantially parallel with the axis of the crankshaft, along the surface of the cylinder block and cylinder head.

  8. Review and status of liquid-cooling technology for gas turbines

    NASA Technical Reports Server (NTRS)

    Vanfossen, G. J., Jr.; Stepka, F. S.

    1979-01-01

    A review was conducted of liquid-cooled turbine technology. Selected liquid-cooled systems and methods are presented along with an assessment of the current technology status and requirements. A comprehensive bibliography is presented.

  9. Apparatus and methods of reheating gas turbine cooling steam and high pressure steam turbine exhaust in a combined cycle power generating system

    DOEpatents

    Tomlinson, Leroy Omar; Smith, Raub Warfield

    2002-01-01

    In a combined cycle system having a multi-pressure heat recovery steam generator, a gas turbine and steam turbine, steam for cooling gas turbine components is supplied from the intermediate pressure section of the heat recovery steam generator supplemented by a portion of the steam exhausting from the HP section of the steam turbine, steam from the gas turbine cooling cycle and the exhaust from the HP section of the steam turbine are combined for flow through a reheat section of the HRSG. The reheated steam is supplied to the IP section inlet of the steam turbine. Thus, where gas turbine cooling steam temperature is lower than optimum, a net improvement in performance is achieved by flowing the cooling steam exhausting from the gas turbine and the exhaust steam from the high pressure section of the steam turbine in series through the reheater of the HRSG for applying steam at optimum temperature to the IP section of the steam turbine.

  10. Gas turbine engine and its associated air intake system

    SciTech Connect

    Ballard, J.R.; Bennett, G.H.; Lee, L.A.

    1984-01-17

    A gas turbine engine and its associated air intake system are disclosed in which the air intake system comprises a generally horizontally extending duct through which an airflow is induced by an ejector pump powered by the engine. A portion of the air passing through the duct is directed through a second duct to the air inlet of the engine. The second duct is connected to the first duct in such a manner that the air directed to the engine air inlet is derived from a vertically upper region of the first duct. The arrangement is intended to reduce the amount of airborne particulate material ingested by the gas turbine engine.

  11. Transpiration cooling using air as a coolant

    SciTech Connect

    Kikkawa, Shinzo; Senda, Mamoru; Sakagushi, Katsuji; Shibutani, Hideki )

    1993-02-01

    Transpiration cooling is one of the most effective techniques for protecting a surface exposed to a high-temperature gas stream. In the present paper, the transpiration cooling effectiveness was measured under steady state. Air as a coolant was transpired from the surface of a porous plate exposed to hot gas stream, and the transpiration rate was varied in the range of 0.001 [approximately] 0.006. The transpiration cooling effectiveness was evaluated by measuring the temperature of the upper surface of the plate. Also, a theoretical study was performed and it was clarified that the effectiveness increases with increasing transpiration rate and heat-transfer coefficient of the upper surface. Further, the effectiveness was expressed as a function of the blowing parameter only. The agreement between the experimental results and theoretical ones was satisfactory.

  12. Experimental study on the evaporative cooling of an air-cooled condenser with humidifying air

    NASA Astrophysics Data System (ADS)

    Wen, Mao-Yu; Ho, Ching-Yen; Jang, Kuang-Jang; Yeh, Cheng-Hsiung

    2014-02-01

    Using six different materials to construct a water curtain, this study aims to determine the most effective spray cooling of an air cooled heat exchanger under wet conditions. The experiments were carried out at a mass flow rate of 0.005-0.01 kg/s (spraying water), an airspeed of 0.6-2.4 m/s and a run time of 0-72 h for the material degradation tests. The experimental results indicate that the cooling efficiency, the heat rejection, and the sprinkling density increase as the amount of spraying water increases, but, the air-flow of the condenser is reduced at the same time. In addition, the cooling efficiency of the pads decreases with an increase of the inlet air velocity. In terms of experimental range, the natural wood pulp fiberscan can reach 42.7-66 % for cooling efficiency and 17.17-24.48 % for increases of heat rejection. This means that the natural wood pulp fiberscan pad most effectively enhances cooling performance, followed in terms of cooling effectiveness by the special non-woven rayon pad, the woollen blanket, biochemistry cotton and kapok, non-woven cloth of rayon cotton and kapok, and white cotton pad, respectively. However, the natural wood pulp fiberscan and special non-woven rayon display a relatively greater degradation of the cooling efficiency than the other test pads used in the material degradation tests.

  13. Facial Cooling During Cold Air Exposure.

    NASA Astrophysics Data System (ADS)

    Tikuisis, Peter; Osczevski, Randall J.

    2003-07-01

    A dynamic model of facial cooling was developed in conjunction with the release of the new wind chill temperature (WCT) index, whereby the WCT provides wind chill estimates based on steady-state considerations and the dynamic model can be used to predict the rate of facial cooling and particularly the onset of freezing. In the present study, the dynamic model is applied to various combinations of air temperature and wind speed, and predictions of the resultant steady-state cheek skin temperatures are tabulated. Superimposed on these tables are times to a cheek skin temperature of 10°C, which has been reported as painful, and times to freezing. For combinations of air temperature and wind speed that result in the same final steady-state cheek temperature or the same WCT, the initial rate of change of skin temperature is higher for those combinations having higher wind speeds. This suggests that during short exposures, high winds combined with low temperatures might be perceived as more stressful than light winds with lower temperatures that result in the same "wind chill." This paper also discloses the paradox that individuals having a low cheek thermal resistance are predicted to experience a more severe WCT, but be at less risk of cooling injury than individuals with higher thermal resistances. The advantages of cooling-time predictions using the dynamic model are discussed with the recommendation/conclusion that safe exposure limits are more meaningful and less ambiguous than the reporting of the WCT.

  14. Stress concentration effects of oblique holes in aspirated-cooled turbine engine liners

    SciTech Connect

    Cencula, J.E.; Coyne, B.J. )

    1992-02-01

    Innovative cooling concepts and new applications of these concepts are used to permit operation of turbopropulsion engines at higher temperatures and with less cooling air for greatest engine performance. These cooling concepts can cause detrimental structural effects due to stress concentrations or high thermal gradients that must be predictable to be incorporated into engine designs. This study analytically predicts the stress concentration effects of various patterns of small, closely-spaced cooling holes drilled through a thin plate and subjected to a biaxial stress field that represents a gas turbine engine application. These predictions are then verified by photoelastic analysis of the cooling hole patterns. Three hole patterns, a symmetrical diamond pattern and two unsymmetric patterns, are examined. The individual cooling holes are circular and drilled at a 30 degree inclination off the surface which produces an elliptical appearance on the surface. Graphical representations of the peak stress concentration factors for a range of stress fields are presented as a result of this study. 2 refs.

  15. Effect of endwall cooling on secondary flows in turbine stator vanes

    NASA Technical Reports Server (NTRS)

    Goldman, L. J.; Mclallin, K. L.

    1977-01-01

    The effect of endwall cooling on the secondary flow behavior and the aerodynamic performance of a core turbine stator vane was determined. The investigation was conducted in a cold-air, full-annular cascade, where three-dimensional effects were obtained. Two endwall cooling configurations were tested. In the first configuration, the cooling holes were oriented so that the coolant was injected in line with the inviscid streamline direction. In the second configuration, the coolant was injected at an angle of 15 deg to the inviscid streamline direction and oriented towards the vane pressure stator. In both cases the stator vanes were solid and uncooled so that the effect of endwall cooling was obtained directly. Total-pressure surveys were taken downstream of the stator vanes over a range of cooling flows at the design, mean-radius, critical velocity ratio of 0.778. Changes in the total-pressure contours downstream of the vanes were used to obtain the effect of endwall cooling on the secondary flows in the stator.

  16. BEETIT: Building Cooling and Air Conditioning

    SciTech Connect

    2010-09-01

    BEETIT Project: The 14 projects that comprise ARPA-E’s BEETIT Project, short for “Building Energy Efficiency Through Innovative Thermodevices,” are developing new approaches and technologies for building cooling equipment and air conditioners. These projects aim to drastically improve building energy efficiency and reduce greenhouse gas emissions such as carbon dioxide (CO2) at a cost comparable to current technologies.

  17. Turbine endwall film cooling with combustor-turbine interface gap leakage flow: Effect of incidence angle

    NASA Astrophysics Data System (ADS)

    Zhang, Yang; Yuan, Xin

    2013-04-01

    This paper is focused on the film cooling performance of combustor-turbine leakage flow at off-design condition. The influence of incidence angle on film cooling effectiveness on first-stage vane endwall with combustor-turbine interface slot is studied. A baseline slot configuration is tested in a low speed four-blade cascade comprising a large-scale model of the GE-E3Nozzle Guide Vane (NGV). The slot has a forward expansion angle of 30 deg. to the endwall surface. The Reynolds number based on the axial chord and inlet velocity of the free-stream flow is 3.5 × 105 and the testing is done in a four-blade cascade with low Mach number condition (0.1 at the inlet). The blowing ratio of the coolant through the interface gap varies from M = 0.1 to M = 0.3, while the blowing ratio varies from M = 0.7 to M = 1.3 for the endwall film cooling holes. The film-cooling effectiveness distributions are obtained using the pressure sensitive paint (PSP) technique. The results show that with an increasing blowing ratio the film-cooling effectiveness increases on the endwall. As the incidence angle varies from i = +10 deg. to i = -10 deg., at low blowing ratio, the averaged film-cooling effectiveness changes slightly near the leading edge suction side area. The case of i = +10 deg. has better film-cooling performance at the downstream part of this region where the axial chord is between 0.15 and 0.25. However, the disadvantage of positive incidence appears when the blowing ratio increases, especially at the upstream part of near suction side region where the axial chord is between 0 and 0.15. On the main passage endwall surface, as the incidence angle changes from i = +10 deg. to i = -10 deg., the averaged film-cooling effectiveness changes slightly and the negative incidence appears to be more effective for the downstream part film cooling of the endwall surface where the axial chord is between 0.6 and 0.8.

  18. HUMID AIR TURBINE CYCLE TECHNOLOGY DEVELOPMENT PROGRAM

    SciTech Connect

    Richard Tuthill

    2002-07-18

    The Humid Air Turbine (HAT) Cycle Technology Development Program focused on obtaining HAT cycle combustor technology that will be the foundation of future products. The work carried out under the auspices of the HAT Program built on the extensive low emissions stationary gas turbine work performed in the past by Pratt & Whitney (P&W). This Program is an integral part of technology base development within the Advanced Turbine Systems Program at the Department of Energy (DOE) and its experiments stretched over 5 years. The goal of the project was to fill in technological data gaps in the development of the HAT cycle and identify a combustor configuration that would efficiently burn high moisture, high-pressure gaseous fuels with low emissions. The major emphasis will be on the development of kinetic data, computer modeling, and evaluations of combustor configurations. The Program commenced during the 4th Quarter of 1996 and closed in the 4th Quarter of 2001. It teamed the National Energy Technology Laboratory (NETL) with P&W, the United Technologies Research Center (UTRC), and a subcontractor on-site at UTRC, kraftWork Systems Inc. The execution of the program started with bench-top experiments that were conducted at UTRC for extending kinetic mechanisms to HAT cycle temperature, pressure, and moisture conditions. The fundamental data generated in the bench-top experiments was incorporated into the analytical tools available at P&W to design the fuel injectors and combustors. The NETL then used the hardware to conduct combustion rig experiments to evaluate the performance of the combustion systems at elevated pressure and temperature conditions representative of the HAT cycle. The results were integrated into systems analysis done by kraftWork to verify that sufficient understanding of the technology had been achieved and that large-scale technological application and demonstration could be undertaken as follow-on activity. An optional program extended the

  19. The influence of external cooling system on the performance of supercritical steam turbine cycles

    NASA Astrophysics Data System (ADS)

    Kosman, Wojciech

    2010-09-01

    The problem presented in this paper refers to the concepts applied to the design of supercritical steam turbines. The issue under the investigation is the presence of a cooling system. Cooling systems aim to protect the main components of the turbines against overheating. However the cooling flows mix with the main flow and modify the expansion line in the steam path. This affects the expansion process in the turbine and changes the performance when compared to the uncooled turbine. The analysis described here investigates the range of the influence of the cooling system on the turbine cycle. This influence is measured mainly through the change of the power generation efficiency. The paper explains the approach towards the assessment of the cooling effects and presents results of the modeling for three supercritical steam cycles.

  20. The Effect of Wake Passing on Turbine Blade Film Cooling

    NASA Technical Reports Server (NTRS)

    Heidmann, James David

    1996-01-01

    The effect of upstream blade row wake passing on the showerhead film cooling performance of a downstream turbine blade has been investigated through a combination of experimental and computational studies. The experiments were performed in a steady-flow annular turbine cascade facility equipped with an upstream rotating row of cylindrical rods to produce a periodic wake field similar to that found in an actual turbine. Spanwise, chordwise, and temporal resolution of the blade surface temperature were achieved through the use of an array of nickel thin-film surface gauges covering one unit cell of showerhead film hole pattern. Film effectiveness and Nusselt number values were determined for a test matrix of various injectants, injectant blowing ratios, and wake Strouhal numbers. Results indicated a demonstratable reduction in film effectiveness with increasing Strouhal number, as well as the expected increase in film effectiveness with blowing ratio. An equation was developed to correlate the span-average film effectiveness data. The primary effect of wake unsteadiness was found to be correlated well by a chordwise-constant decrement of 0.094-St. Measurable spanwise film effectiveness variations were found near the showerhead region, but meaningful unsteady variations and downstream spanwise variations were not found. Nusselt numbers were less sensitive to wake and injection changes. Computations were performed using a three-dimensional turbulent Navier-Stokes code which was modified to model wake passing and film cooling. Unsteady computations were found to agree well with steady computations provided the proper time-average blowing ratio and pressure/suction surface flow split are matched. The remaining differences were isolated to be due to the enhanced mixing in the unsteady solution caused by the wake sweeping normally on the pressure surface. Steady computations were found to be in excellent agreement with experimental Nusselt numbers, but to overpredict

  1. Heat Transfer Experiments in the Internal Cooling Passages of a Cooled Radial Turbine Rotor

    NASA Technical Reports Server (NTRS)

    Johnson, B. V.; Wagner, J. H.

    1996-01-01

    An experimental study was conducted (1) to experimentally measure, assess and analyze the heat transfer within the internal cooling configuration of a radial turbine rotor blade and (2) to obtain heat transfer data to evaluate and improve computational fluid dynamics (CFD) procedures and turbulent transport models of internal coolant flows. A 1.15 times scale model of the coolant passages within the NASA LERC High Temperature Radial Turbine was designed, fabricated of Lucite and instrumented for transient beat transfer tests using thin film surface thermocouples and liquid crystals to indicate temperatures. Transient heat transfer tests were conducted for Reynolds numbers of one-fourth, one-half, and equal to the operating Reynolds number for the NASA Turbine. Tests were conducted for stationary and rotating conditions with rotation numbers in the range occurring in the NASA Turbine. Results from the experiments showed the heat transfer characteristics within the coolant passage were affected by rotation. In general, the heat transfer increased and decreased on the sides of the straight radial passages with rotation as previously reported from NASA-HOST-sponsored experiments. The heat transfer in the tri-passage axial flow region adjacent to the blade exit was relatively unaffected by rotation. However, the heat transfer on one surface, in the transitional region between the radial inflow passage and axial, constant radius passages, decreased to approximately 20 percent of the values without rotation. Comparisons with previous 3-D numerical studies indicated regions where the heat transfer characteristics agreed and disagreed with the present experiment.

  2. The Wells turbine in an oscillating air flow

    SciTech Connect

    Raghunathan, S.; Ombaka,

    1984-08-01

    An experimental study of the performance of a 0.2 m diameter Wells self rectifying air turbine with NACA 0021 blades is presented. Experiments were conducted in an oscillating flowrig. The effects of Reynolds number and Strouhal number on the performance of the turbine were investigated. Finally comparison between the results with the predictions from uni-directional flow tests are made.

  3. Flow visualization of discrete hole film cooling for gas turbine applications

    NASA Technical Reports Server (NTRS)

    Colladay, R. S.; Russell, L. M.

    1975-01-01

    Film injection from discrete holes in a three row staggered array with 5-diameter spacing is studied. The boundary layer thickness-to-hole diameter ratio and Reynolds number are typical of gas turbine film cooling applications. Two different injection locations are studied to evaluate the effect of boundary layer thickness on film penetration and mixing. Detailed streaklines showing the turbulent motion of the injected air are obtained by photographing neutrally buoyant helium filled soap bubbles which follow the flow field. The bubble streaklines passing downstream injection locations are clearly identifiable and can be traced back to their origin. Visualization of surface temperature patterns obtained from infrared photographs of a similar film cooled surface are also included.

  4. An efficient liner cooling scheme for advanced small gas turbine combustors

    NASA Technical Reports Server (NTRS)

    Paskin, Marc D.; Mongia, Hukam C.; Acosta, Waldo A.

    1993-01-01

    A joint Army/NASA program was conducted to design, fabricate, and test an advanced, small gas turbine, reverse-flow combustor utilizing a compliant metal/ceramic (CMC) wall cooling concept. The objectives of this effort were to develop a design method (basic design data base and analysis) for the CMC cooling technique and then demonstrate its application to an advanced cycle, small, reverse-flow combustor with 3000 F burner outlet temperature. The CMC concept offers significant improvements in wall cooling effectiveness resulting in a large reduction in cooling air requirements. Therefore, more air is available for control of burner outlet temperature pattern in addition to the benefits of improved efficiency, reduced emissions, and lower smoke levels. The program was divided into four tasks. Task 1 defined component materials and localized design of the composite wall structure in conjunction with development of basic design models for the analysis of flow and heat transfer through the wall. Task 2 included implementation of the selected materials and validated design models during combustor preliminary design. Detail design of the selected combustor concept and its refinement with 3D aerothermal analysis were completed in Task 3. Task 4 covered detail drawings, process development and fabrication, and a series of burner rig tests. The purpose of this paper is to provide details of the investigation into the fundamental flow and heat transfer characteristics of the CMC wall structure as well as implementation of the fundamental analysis method for full-scale combustor design.

  5. Finger Cooling During Cold Air Exposure.

    NASA Astrophysics Data System (ADS)

    Tikuisis, Peter

    2004-05-01

    This paper presents a method for predicting the onset of finger freezing. It is an extension of a tissue-cooling model originally developed to predict the onset of cheek freezing. The extension to the finger is presented as a more conservative warning of wind chill. Indeed, guidance on the risk of finger freezing is important not only to safeguard the finger, but also because it pertains more closely to susceptible facial features, such as the nose, than if only the risk of cheek freezing was provided. The importance of blood flow to the finger and the modeling of vaso-constriction are demonstrated through cooling predictions that agree reasonably well with several reported observations. Differences in the prediction between the present physiologic-based model and the engineering model used to develop the wind chill index are also discussed. New wind chill charts are presented that tabulate the mean cooling rates and corresponding onset times to freezing of the finger for various combinations of air temperature and wind speed. Results indicate that the surface of the finger cools to its freezing point in approximately one-eighth of the time predicted for the cheek. For combinations that result in the same wind chill temperature (WCT), the rate of finger cooling is faster at the higher wind speed. This asymmetry was previously disclosed through the application of the model to cheek cooling, and it reiterates the ambiguity associated with the reporting of WCT. It is further emphasized that the reporting of onset times to freezing, or safe exposure limits, is a more logical and meaningful alternative to the WCT.

  6. Experimental study of condensate subcooling with the use of a model of an air-cooled condenser

    NASA Astrophysics Data System (ADS)

    Sukhanov, V. A.; Bezukhov, A. P.; Bogov, I. A.; Dontsov, N. Y.; Volkovitsky, I. D.; Tolmachev, V. V.

    2016-01-01

    Water-supply deficit is now felt in many regions of the world. This hampers the construction of new steam-turbine and combined steam-and-gas thermal power plants. The use of dry cooling systems and, specifically, steam-turbine air-cooled condensers (ACCs) expands the choice of sites for the construction of such power plants. The significance of condensate subcooling Δ t as a parameter that negatively affects the engineering and economic performance of steam-turbine plants is thereby increased. The operation and design factors that influence the condensate subcooling in ACCs are revealed, and the research objective is, thus, formulated properly. The indicated research was conducted through physical modeling with the use of the Steam-Turbine Air-Cooled Condenser Unit specialized, multipurpose, laboratory bench. The design and the combined schematic and measurement diagram of this test bench are discussed. The experimental results are presented in the form of graphic dependences of the condensate subcooling value on cooling ratio m and relative weight content ɛ' of air in steam at the ACC inlet at different temperatures of cooling air t ca ' . The typical ranges of condensate subcooling variation (4 ≤ Δ t ≤ 6°C, 2 ≤ Δ t ≤ 4°C, and 0 ≤ Δ t ≤ 2°C) are identified based on the results of analysis of the attained Δ t levels in the ACC and numerous Δ t reduction estimates. The corresponding ranges of cooling ratio variation at different temperatures of cooling air at the ACC inlet are specified. The guidelines for choosing the adjusted ranges of cooling ratio variation with account of the results of experimental studies of the dependences of the absolute pressure of the steam-air mixture in the top header of the ACC and the heat flux density on the cooling ratio at different temperatures of cooling air at the ACC inlet are given.

  7. Oil cooling system for a gas turbine engine

    NASA Technical Reports Server (NTRS)

    Coffinberry, G. A.; Kast, H. B. (Inventor)

    1977-01-01

    A gas turbine engine fuel delivery and control system is provided with means to recirculate all fuel in excess fuel control requirements back to the aircraft fuel tank. This increases the fuel pump heat sink and decreases the pump temperature rise without the addition of valving other than normally employed. A fuel/oil heat exchanger and associated circuitry is provided to maintain the hot engine oil in heat exchange relationship with the cool engine fuel. Where anti-icing of the fuel filter is required, means are provided to maintain the fuel temperature entering the filter at or above a minimum level to prevent freezing thereof. In one embodiment, a divider valve is provided to take all excess fuel from either upstream or downstream of the fuel filter and route it back to the tanks, the ratio of upstream to downstream extraction being a function of fuel pump discharge pressure.

  8. 40 CFR 92.108 - Intake and cooling air measurements.

    Code of Federal Regulations, 2013 CFR

    2013-07-01

    ... 40 Protection of Environment 21 2013-07-01 2013-07-01 false Intake and cooling air measurements. 92.108 Section 92.108 Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) CONTROL OF AIR POLLUTION FROM LOCOMOTIVES AND LOCOMOTIVE ENGINES Test Procedures § 92.108 Intake and cooling air measurements....

  9. Selection of the air heat exchanger operating in a gas turbine air bottoming cycle

    NASA Astrophysics Data System (ADS)

    Chmielniak, Tadeusz; Czaja, Daniel; Lepszy, Sebastian

    2013-12-01

    A gas turbine air bottoming cycle consists of a gas turbine unit and the air turbine part. The air part includes a compressor, air expander and air heat exchanger. The air heat exchanger couples the gas turbine to the air cycle. Due to the low specific heat of air and of the gas turbine exhaust gases, the air heat exchanger features a considerable size. The bigger the air heat exchanger, the higher its effectiveness, which results in the improvement of the efficiency of the gas turbine air bottoming cycle. On the other hand, a device with large dimensions weighs more, which may limit its use in specific locations, such as oil platforms. The thermodynamic calculations of the air heat exchanger and a preliminary selection of the device are presented. The installation used in the calculation process is a plate heat exchanger, which is characterized by a smaller size and lower values of the pressure drop compared to the shell and tube heat exchanger. Structurally, this type of the heat exchanger is quite similar to the gas turbine regenerator. The method on which the calculation procedure may be based for real installations is also presented, which have to satisfy the economic criteria of financial profitability and cost-effectiveness apart from the thermodynamic criteria.

  10. Fouling of Air Cooled Condensers On the Air Side

    NASA Astrophysics Data System (ADS)

    Marie, Hazel; Matune, Nicholas

    2013-11-01

    As the electrical power demand increases and water resources become more limited, fouling on the air side of Air Cooled Condensers (ACC) is a growing concern. The objective of this study was to experimentally and computationally calculate the convection heat transfer coefficient for both a clean and fouled condenser. Bee pollen was selected as the experimental fouling particle, and engineering data for similar particles were used for the computational model. Both the experimental and computational results showing the negative impact fouling has a on the heat transfer will be presented.

  11. Computation of flow and heat transfer in rotating cavities with peripheral flow of cooling air.

    PubMed

    Kiliç, M

    2001-05-01

    Numerical solutions of the Navier-Stokes equations have been used to model the flow and the heat transfer that occurs in the internal cooling-air systems of gas turbines. Computations are performed to study the effect of gap ratio, Reynolds number and the mass flow rate on the flow and the heat transfer structure inside isothermal and heated rotating cavities with peripheral flow of cooling air. Computations are compared with some of the recent experimental work on flow and heat transfer in rotating-cavities. The agreement between the computed and the available experimental data is reasonably good. PMID:11460668

  12. Leading edge film cooling effects on turbine blade heat transfer

    NASA Technical Reports Server (NTRS)

    Garg, Vijay K.; Gaugler, Raymond E.

    1995-01-01

    An existing three dimensional Navier-Stokes code, modified to include film cooling considerations, has been used to study the effect of spanwise pitch of shower-head holes and coolant to mainstream mass flow ratio on the adiabatic effectiveness and heat transfer coefficient on a film-cooled turbine vane. The mainstream is akin to that under real engine conditions with stagnation temperature = 1900 K and stagnation pressure = 3 MPa. It is found that with the coolant to mainstream mass flow ratio fixed, reducing P, the spanwise pitch for shower-head holes, from 7.5 d to 3.0 d, where d is the hole diameter, increases the average effectiveness considerably over the blade surface. However, when P/d= 7.5, increasing the coolant mass flow increases the effectiveness on the pressure surface but reduces it on the suction surface due to coolant jet lift-off. For P/d = 4.5 or 3.0, such an anomaly does not occur within the range of coolant to mainstream mass flow ratios analyzed. In all cases, adiabatic effectiveness and heat transfer coefficient are highly three-dimensional.

  13. Effects of film injection angle on turbine vane cooling

    NASA Technical Reports Server (NTRS)

    Gauntner, J. W.

    1977-01-01

    Film ejection from discrete holes in the suction surface of a turbine vane was studied for hole axes (1) slanted 30 deg to the surface in the streamwise direction and (2) slanted 30 deg to the surface and 45 deg from the streamwise direction toward the hub. The holes were near the throat area in a five-row staggered array with 8-diameter spacing. Mass flux ratios were as high as 1.2. The data were obtained in an annular sector cascade at conditions where both the ratio of the boundary layer momentum thickness-to-hole diameter and the momentum thickness Reynolds number were typical of an advanced turbofan engine at both takeoff and cruise. Wall temperatures were measured downstream of each of the rows of holes. Results of this study are expressed as a comparison of cooling effectiveness between the in-line angle injection and the compound-angle injection as a function of mass flux ratio. These heat transfer results are also compared with the results of a referenced flow visualization study. Also included is a closed-form analytical solution for temperature within the film cooled wall.

  14. Gas turbine combustor transition

    DOEpatents

    Coslow, Billy Joe; Whidden, Graydon Lane

    1999-01-01

    A method of converting a steam cooled transition to an air cooled transition in a gas turbine having a compressor in fluid communication with a combustor, a turbine section in fluid communication with the combustor, the transition disposed in a combustor shell and having a cooling circuit connecting a steam outlet and a steam inlet and wherein hot gas flows from the combustor through the transition and to the turbine section, includes forming an air outlet in the transition in fluid communication with the cooling circuit and providing for an air inlet in the transition in fluid communication with the cooling circuit.

  15. Gas turbine combustor transition

    DOEpatents

    Coslow, B.J.; Whidden, G.L.

    1999-05-25

    A method is described for converting a steam cooled transition to an air cooled transition in a gas turbine having a compressor in fluid communication with a combustor, a turbine section in fluid communication with the combustor, the transition disposed in a combustor shell and having a cooling circuit connecting a steam outlet and a steam inlet and wherein hot gas flows from the combustor through the transition and to the turbine section, includes forming an air outlet in the transition in fluid communication with the cooling circuit and providing for an air inlet in the transition in fluid communication with the cooling circuit. 7 figs.

  16. Ceramic thermal-barrier coatings for cooled turbines

    NASA Technical Reports Server (NTRS)

    Liebert, C. H.; Stepka, F. S.

    1976-01-01

    Coating systems consisting of a plasma sprayed layer of zirconia stabilized with either yttria, magnesia or calcia over a thin alloy bond coat have been developed, their potential was analyzed and their durability and benefits evaluated in a turbojet engine. The coatings on air cooled rotating blades were in good condition after completing as many as 500 two-minute cycles of engine operation between full power at a gas temperature of 1644 K and flameout, or as much as 150 hours of steady state operation on cooled vanes and blades at gas temperatures as high as 1644 K with 35 start and stop cycles. On the basis of durability and processing cost, the yttria stabilized zirconia was considered the best of the three coatings investigated.

  17. The problem of cooling an air-cooled cylinder on an aircraft engine

    NASA Technical Reports Server (NTRS)

    Brevoort, M J; Joyner, U T

    1941-01-01

    An analysis of the cooling problem has been to show by what means the cooling of an air-cooled aircraft engine may be improved. Each means of improving cooling is analyzed on the basis of effectiveness in cooling with respect to power for cooling. The altitude problem is analyzed for both supercharged and unsupercharged engines. The case of ground cooling is also discussed. The heat-transfer process from the hot gases to the cylinder wall is discussed on the basis of the fundamentals of heat transfer and thermodynamics. Adiabatic air-temperature rise at a stagnation point in compressible flow is shown to depend only on the velocity of flow.

  18. Simulation of the Effects of Cooling Techniques on Turbine Blade Heat Transfer

    NASA Astrophysics Data System (ADS)

    Shaw, Vince; Fatuzzo, Marco

    Increases in the performance demands of turbo machinery has stimulated the development many new technologies over the last half century. With applications that spread beyond marine, aviation, and power generation, improvements in gas turbine technologies provide a vast impact. High temperatures within the combustion chamber of the gas turbine engine are known to cause an increase in thermal efficiency and power produced by the engine. However, since operating temperatures of these engines reach above 1000 K within the turbine section, the need for advances in material science and cooling techniques to produce functioning engines under these high thermal and dynamic stresses is crucial. As with all research and development, costs related to the production of prototypes can be reduced through the use of computational simulations. By making use of Ansys Simulation Software, the effects of turbine cooling techniques were analyzed. Simulation of the Effects of Cooling Techniques on Turbine Blade Heat Transfer.

  19. Startup of air-cooled condensers and dry cooling towers at low temperatures of the cooling air

    NASA Astrophysics Data System (ADS)

    Milman, O. O.; Ptakhin, A. V.; Kondratev, A. V.; Shifrin, B. A.; Yankov, G. G.

    2016-05-01

    The problems of startup and performance of air-cooled condensers (ACC) and dry cooling towers (DCT) at low cooling air temperatures are considered. Effects of the startup of the ACC at sub-zero temperatures are described. Different options of the ACC heating up are analyzed, and examples of existing technologies are presented (electric heating, heating up with hot air or steam, and internal and external heating). The use of additional heat exchanging sections, steam tracers, in the DCT design is described. The need for high power in cases of electric heating and heating up with hot air is noted. An experimental stand for research and testing of the ACC startup at low temperatures is described. The design of the three-pass ACC unit is given, and its advantages over classical single-pass design at low temperatures are listed. The formation of ice plugs inside the heat exchanging tubes during the start-up of ACC and DCT at low cooling air temperatures is analyzed. Experimental data on the effect of the steam flow rate, steam nozzle distance from the heat-exchange surface, and their orientation in space on the metal temperature were collected, and test results are analyzed. It is noted that the surface temperature at the end of the heat up is almost independent from its initial temperature. Recommendations for the safe start-up of ACCs and DCTs are given. The heating flow necessary to sufficiently heat up heat-exchange surfaces of ACCs and DCTs for the safe startup is estimated. The technology and the process of the heat up of the ACC with the heating steam external supply are described by the example of the startup of the full-scale section of the ACC at sub-zero temperatures of the cooling air, and the advantages of the proposed start-up technology are confirmed.

  20. Air film cooling in a nonadiabatic wall conical nozzle.

    NASA Technical Reports Server (NTRS)

    Boldman, D. R.; Papell, S. S.; Ehlers, R. C.

    1972-01-01

    Experimental data for an air-film cooled conical nozzle operating with a heated-air main stream and a water-cooled wall confirm the validity of Lieu's (1964) method for correlating film cooling data in the accelerated flow of a nonadiabatic-wall nozzle. The film cooling effectiveness modified for nonadiabatic walls by Lieu can be used to correlate film cooling under the condition that the main-stream to coolant velocity ratio at the slot is about 1. Such a ratio provides the optimum cooling effectiveness.

  1. High temperature gas-cooled reactor: gas turbine application study

    SciTech Connect

    Not Available

    1980-12-01

    The high-temperature capability of the High-Temperature Gas-Cooled Reactor (HTGR) is a distinguishing characteristic which has long been recognized as significant both within the US and within foreign nuclear energy programs. This high-temperature capability of the HTGR concept leads to increased efficiency in conventional applications and, in addition, makes possible a number of unique applications in both electrical generation and industrial process heat. In particular, coupling the HTGR nuclear heat source to the Brayton (gas turbine) Cycle offers significant potential benefits to operating utilities. This HTGR-GT Application Study documents the effort to evaluate the appropriateness of the HTGR-GT as an HTGR Lead Project. The scope of this effort included evaluation of the HTGR-GT technology, evaluation of potential HTGR-GT markets, assessment of the economics of commercial HTGR-GT plants, and evaluation of the program and expenditures necessary to establish HTGR-GT technology through the completion of the Lead Project.

  2. Cooling system having reduced mass pin fins for components in a gas turbine engine

    DOEpatents

    Lee, Ching-Pang; Jiang, Nan; Marra, John J

    2014-03-11

    A cooling system having one or more pin fins with reduced mass for a gas turbine engine is disclosed. The cooling system may include one or more first surfaces defining at least a portion of the cooling system. The pin fin may extend from the surface defining the cooling system and may have a noncircular cross-section taken generally parallel to the surface and at least part of an outer surface of the cross-section forms at least a quartercircle. A downstream side of the pin fin may have a cavity to reduce mass, thereby creating a more efficient turbine airfoil.

  3. Endwall shape modification using vortex generators and fences to improve gas turbine cooling and effectiveness

    NASA Astrophysics Data System (ADS)

    Gokce, Zeki Ozgur

    The gas turbine is one of the most important parts of the air-breathing jet engine. Hence, improving its efficiency and rendering it operable under high temperatures are constant goals for the aerospace industry. Two types of flow within the gas turbine are of critical relevance: The flow around the first row of stator blades (also known as the nozzle guide vane blade - NGV) and the cooling flow inside the turbine blade cooling channel. The subject of this thesis work was to search for methods that could improve the characteristics of these two types of flows, thus enabling superior engine performance. The innovative aspect of our work was to apply an endwall shape modification previously employed by non-aerospace industries for cooling applications, to the gas turbine cooling flow which is vital to aerospace propulsion. Since the costs of investigating the possible benefits of any idea via extensive experiments could be quite high, we decided to use computational fluid dynamics (CFD) followed by experimentation as our methodology. We decided to analyze the potential benefits of using vortex generators (VGs) as well as the rectangular endwall fence. Since the pin-fins used in cooling flow are circular cylinders, and since the boundary layer flow is mainly characterized by the leading edge diameter of the NGV blade, we modeled both the pin-fins and the NGV blade as vertical circular cylinders. The baseline case consisted of the cylinder(s) being subjected to cross flow and a certain amount of freestream turbulence. The modifications we made on the endwall consisted of rectangular fences. In the case of the cooling flow, we used triangular shaped, common flow up oriented, delta winglet type vortex generators as well as rectangular endwall fences. The channel contained singular cylinders as well as staggered rows of multiple cylinders. For the NGV flow, a rectangular endwall fence and a singular cylinder were utilized. Using extensive CFD modeling and analysis, we

  4. Development of an advanced ceramic turbine wheel for an air turbine starter

    NASA Astrophysics Data System (ADS)

    Poplawsky, Carl J.; Lindberg, Laura; Robb, Scott; Roundy, James

    1992-10-01

    A ceramic turbine wheel has been designed as a retrofit for Waspaloy for a military cartridge mode air turbine starter. This results in reduced cost and weight while increasing resistance to temperature, erosion, and corrosion. Techniques used to perform ceramic turbine three-dimensional fast fracture reliability analysis were verified with spin testing of ceramic test rotors and correlated well with burst speed predictions. Reliability estimates have been made for design and proof conditions, providing guidance for selecting a ceramic supplier and for determining proof test yield. Room temperature whirlpit burst testing is planned to verify the mechanical design and reliability of the wheel.

  5. Development of an advanced ceramic turbine wheel for an air turbine starter

    SciTech Connect

    Poplawsky, C.J.; Lindberg, L.; Robb, S.; Roundy, J.

    1992-01-01

    A ceramic turbine wheel has been designed as a retrofit for Waspaloy for a military cartridge mode air turbine starter. This results in reduced cost and weight while increasing resistance to temperature, erosion, and corrosion. Techniques used to perform ceramic turbine three-dimensional fast fracture reliability analysis were verified with spin testing of ceramic test rotors and correlated well with burst speed predictions. Reliability estimates have been made for design and proof conditions, providing guidance for selecting a ceramic supplier and for determining proof test yield. Room temperature whirlpit burst testing is planned to verify the mechanical design and reliability of the wheel. 9 refs.

  6. Air-cooled condensers eliminate plant water use

    SciTech Connect

    Wurtz, W.; Peltier, R.

    2008-09-15

    River or ocean water has been the mainstay for condensing turbine exhaust steam since the first steam turbine began generating electricity. A primary challenge facing today's plant developers, especially in drought-prone regions, is incorporating processes that reduce plant water use and consumption. One solution is to shed the conventional mindset that once-through cooling is the only option and adopt dry cooling technologies that reduce plant water use from a flood to a few sips. A case study at the Astoria Energy plant, New York City is described. 14 figs.

  7. Thermal design study of an air-cooled plug-nozzle system for a supersonic cruise aircraft

    NASA Technical Reports Server (NTRS)

    Clark, J. S.; Lieberman, A.

    1972-01-01

    A heat-transfer design analysis has been made of an air-cooled plug-nozzle system for a supersonic-cruise aircraft engine. The proposed 10deg half-angle conical plug is sting supported from the turbine frame. Plug cooling is accomplished by convection and film cooling. The flight profile studied includes maximum afterburning from takeoff to Mach 2.7 and supersonic cruise at Mach 2.7 with a low afterburner setting. The calculations indicate that, for maximum afterburning, about 2 percent of the engine primary flow, removed after the second stage of the nine-stage compressor, will adequately cool the plug and sting support. Ram air may be used for cooling during supersonic-cruise operations, however. Therefore, the cycle efficiency penalty paid for air cooling the plug and sting support should be low.

  8. High-Altitude Flight Cooling Investigation of a Radial Air-Cooled Engine

    NASA Technical Reports Server (NTRS)

    Manganiello, Eugene J; Valerino, Michael F; Bell, E Barton

    1947-01-01

    An investigation of the cooling of an 18-cylinder, twin-row, radial, air-cooled engine in a high-performance pursuit airplane has been conducted for variable engine and flight conditions at altitudes ranging from 5000 to 35,000 feet in order to provide a basis for predicting high-altitude cooling performance from sea-level or low altitude experimental results. The engine cooling data obtained were analyzed by the usual NACA cooling-correlation method wherein cylinder-head and cylinder-barrel temperatures are related to the pertinent engine and cooling-air variables. A theoretical analysis was made of the effect on engine cooling of the change of density of the cooling air across the engine (the compressibility effect), which becomes of increasing importance as altitude is increased. Good agreement was obtained between the results of the theoretical analysis and the experimental data.

  9. Aero-Thermo-Structural Design Optimization of Internally Cooled Turbine Blades

    NASA Technical Reports Server (NTRS)

    Dulikravich, G. S.; Martin, T. J.; Dennis, B. H.; Lee, E.; Han, Z.-X.

    1999-01-01

    A set of robust and computationally affordable inverse shape design and automatic constrained optimization tools have been developed for the improved performance of internally cooled gas turbine blades. The design methods are applicable to the aerodynamics, heat transfer, and thermoelasticity aspects of the turbine blade. Maximum use of the existing proven disciplinary analysis codes is possible with this design approach. Preliminary computational results demonstrate possibilities to design blades with minimized total pressure loss and maximized aerodynamic loading. At the same time, these blades are capable of sustaining significantly higher inlet hot gas temperatures while requiring remarkably lower coolant mass flow rates. These results suggest that it is possible to design internally cooled turbine blades that will cost less to manufacture, will have longer life span, and will perform as good, if not better than, film cooled turbine blades.

  10. Advanced multistage turbine blade aerodynamics, performance, cooling, and heat transfer

    SciTech Connect

    Fleeter, S.; Lawless, P.B.

    1995-10-01

    The gas turbine has the potential for power production at the highest possible efficiency. The challenge is to ensure that gas turbines operate at the optimum efficiency so as to use the least fuel and produce minimum emissions. A key component to meeting this challenge is the turbine. Turbine performance, both aerodynamics and heat transfer, is one of the barrier advanced gas turbine development technologies. This is a result of the complex, highly three-dimensional and unsteady flow phenomena in the turbine. Improved turbine aerodynamic performance has been achieved with three-dimensional highly-loaded airfoil designs, accomplished utilizing Euler or Navier-Stokes Computational Fluid Dynamics (CFD) codes. These design codes consider steady flow through isolated blade rows. Thus they do not account for unsteady flow effects. However, unsteady flow effects have a significant impact on performance. Also, CFD codes predict the complete flow field. The experimental verification of these codes has traditionally been accomplished with point data - not corresponding plane field measurements. Thus, although advanced CFD predictions of the highly complex and three-dimensional turbine flow fields are available, corresponding data are not. To improve the design capability for high temperature turbines, a detailed understanding of the highly unsteady and three-dimensional flow through multi-stage turbines is necessary. Thus, unique data are required which quantify the unsteady three-dimensional flow through multi-stage turbine blade rows, including the effect of the film coolant flow. This requires experiments in appropriate research facilities in which complete flow field data, not only point measurements, are obtained and analyzed. Also, as design CFD codes do not account for unsteady flow effects, the next logical challenge and the current thrust in CFD code development is multiple-stage analyses that account for the interactions between neighboring blade rows.

  11. Cooled-turbine aerodynamic performance prediction from reduced primary to coolant total-temperature-ratio results

    NASA Technical Reports Server (NTRS)

    Goldman, L. J.

    1976-01-01

    The prediction of the cooled aerodynamic performance, for both stators and turbines, at actual primary to coolant inlet total temperature ratios from the results obtained at a reduced total temperature ratio is described. Theoretical and available experimental results were compared for convection film and transpiration cooled stator vanes and for a film cooled, single stage core turbine. For these tests the total temperature ratio varied from near 1.0 to about 2.7. The agreement between the theoretical and the experimental results was, in general, reasonable.

  12. Hydrogen-air energy storage gas-turbine system

    NASA Astrophysics Data System (ADS)

    Schastlivtsev, A. I.; Nazarova, O. V.

    2016-02-01

    A hydrogen-air energy storage gas-turbine unit is considered that can be used in both nuclear and centralized power industries. However, it is the most promising when used for power-generating plants based on renewable energy sources (RES). The basic feature of the energy storage system in question is combination of storing the energy in compressed air and hydrogen and oxygen produced by the water electrolysis. Such a process makes the energy storage more flexible, in particular, when applied to RES-based power-generating plants whose generation of power may considerably vary during the course of a day, and also reduces the specific cost of the system by decreasing the required volume of the reservoir. This will allow construction of such systems in any areas independent of the local topography in contrast to the compressed-air energy storage gas-turbine plants, which require large-sized underground reservoirs. It should be noted that, during the energy recovery, the air that arrives from the reservoir is heated by combustion of hydrogen in oxygen, which results in the gas-turbine exhaust gases practically free of substances hazardous to the health and the environment. The results of analysis of a hydrogen-air energy storage gas-turbine system are presented. Its layout and the principle of its operation are described and the basic parameters are computed. The units of the system are analyzed and their costs are assessed; the recovery factor is estimated at more than 60%. According to the obtained results, almost all main components of the hydrogen-air energy storage gas-turbine system are well known at present; therefore, no considerable R&D costs are required. A new component of the system is the H2-O2 combustion chamber; a difficulty in manufacturing it is the necessity of ensuring the combustion of hydrogen in oxygen as complete as possible and preventing formation of nitric oxides.

  13. A Computational Study for the Utilization of Jet Pulsations in Gas Turbine Film Cooling and Flow Control

    NASA Technical Reports Server (NTRS)

    Kartuzova, Olga V.

    2012-01-01

    This report is the second part of a three-part final report of research performed under an NRA cooperative Agreement contract. The first part is NASA/CR-2012-217415. The third part is NASA/CR-2012-217417. Jets have been utilized in various turbomachinery applications in order to improve gas turbines performance. Jet pulsation is a promising technique because of the reduction in the amount of air removed from compressor. In this work two areas of pulsed jets applications were computationally investigated using the commercial code Fluent (ANSYS, Inc.); the first one is film cooling of High Pressure Turbine (HPT) blades and second one is flow separation control over Low Pressure Turbine (LPT) airfoil using Vortex Generator Jets (VGJ). Using pulsed jets for film cooling purposes can help to improve the effectiveness and thus allow higher turbine inlet temperature. Effects of the film hole geometry, blowing ratio and density ratio of the jet, pulsation frequency and duty cycle of blowing on the film cooling effectiveness were investigated. As for the low-pressure turbine (LPT) stages, the boundary layer separation on the suction side of airfoils can occur due to strong adverse pressure gradients. The problem is exacerbated as airfoil loading is increased. Active flow control could provide a means for minimizing separation under conditions where it is most severe (low Reynolds number), without causing additional losses under other conditions (high Reynolds number). The effects of the jet geometry, blowing ratio, density ratio, pulsation frequency and duty cycle on the size of the separated region were examined in this work. The results from Reynolds Averaged Navier-Stokes and Large Eddy Simulation computational approaches were compared with the experimental data.

  14. Development of cooling strategy for an air cooled lithium-ion battery pack

    NASA Astrophysics Data System (ADS)

    Sun, Hongguang; Dixon, Regan

    2014-12-01

    This paper describes a cooling strategy development method for an air cooled battery pack with lithium-ion pouch cells used in a hybrid electric vehicle (HEV). The challenges associated with the temperature uniformity across the battery pack, the temperature uniformity within each individual lithium-ion pouch cell, and the cooling efficiency of the battery pack are addressed. Initially, a three-dimensional battery pack thermal model developed based on simplified electrode theory is correlated to physical test data. An analytical design of experiments (DOE) approach using Optimal Latin-hypercube technique is then developed by incorporating a DOE design model, the correlated battery pack thermal model, and a morphing model. Analytical DOE studies are performed to examine the effects of cooling strategies including geometries of the cooling duct, cooling channel, cooling plate, and corrugation on battery pack thermal behavior and to identify the design concept of an air cooled battery pack to maximize its durability and its driving range.

  15. Thermal conditions for cooled gas-turbine metal-ceramic blade

    NASA Astrophysics Data System (ADS)

    Soudarev, A. V.; Soudarev, B. V.; Molchanov, A. S.; Souryaninov, A. A.; Grishaev, V. V.

    2002-02-01

    Application of the alumo-boron-nitride heat-resistant structural ceramics allows distribution of the thermal and mechanical loads on the metal-ceramic blade elements reasonably rationally from the thermotechnical point of view. The ceramic shell, actually free of the mechanical effects, absorbs the heat from the high-temperature gas and serves as a shield for the strength core. The latter, being loaded mechanically, is cooled with air, the flow thereof is mainly the function of the heat supply from the peripheral platform and ceramic shell, additionally separated by a thin- wall metal screen from the core. Calculation of the pattern factors for the basic parts was performed at rating as applied to the nozzle vanes and rotor blades of the 2.5 MW GTE with the gas temperature at the inlet TIT=1623K. It was demonstrated that an admissible temperature level of the mechanically loaded parts could be achieved at the cooling air flows of 1.5%. Decreasing the power consumption on cooling allowed to get a high efficiency of the designed engine amounting to 42 43% (speed at rating is around 23,000 r/min). During rotation the length of the ceramic shell, installed loosely on the strength core, moves due to the action of the centrifugal forces and is pressed to the platform of the core. At the same time, a relatively lower compressive stresses of around 40 MPa are generated in the shell which ensures a feasibility of a long-term reliable operation of the turbine.

  16. Correction of Temperatures of Air-Cooled Engine Cylinders for Variation in Engine and Cooling Conditions

    NASA Technical Reports Server (NTRS)

    Schey, Oscar W; Pinkel, Benjamin; Ellerbrock, Herman H , Jr

    1939-01-01

    Factors are obtained from semiempirical equations for correcting engine-cylinder temperatures for variation in important engine and cooling conditions. The variation of engine temperatures with atmospheric temperature is treated in detail, and correction factors are obtained for various flight and test conditions, such as climb at constant indicated air speed, level flight, ground running, take-off, constant speed of cooling air, and constant mass flow of cooling air. Seven conventional air-cooled engine cylinders enclosed in jackets and cooled by a blower were tested to determine the effect of cooling-air temperature and carburetor-air temperature on cylinder temperatures. The cooling air temperature was varied from approximately 80 degrees F. to 230 degrees F. and the carburetor-air temperature from approximately 40 degrees F. to 160 degrees F. Tests were made over a large range of engine speeds, brake mean effective pressures, and pressure drops across the cylinder. The correction factors obtained experimentally are compared with those obtained from the semiempirical equations and a fair agreement is noted.

  17. Cooling Characteristics of an Experimental Tail-pipe Burner with an Annular Cooling-air Passage

    NASA Technical Reports Server (NTRS)

    Kaufman, Harold R; Koffel, William K

    1952-01-01

    The effects of tail-pipe fuel-air ratio (exhaust-gas temperatures from approximately 3060 degrees to 3825 degrees R), radial distributiion of tail-pipe fuel flow, and mass flow of combustion gas and the inside wall were determined for an experimental tail-pipe burner cooled by air flowing through and insulated cooling-air to combustion gas mass flow from 0.066 to 0.192 were also determined.

  18. Effect of Coolant Temperature and Mass Flow on Film Cooling of Turbine Blades

    NASA Technical Reports Server (NTRS)

    Garg, Vijay K.; Gaugler, Raymond E.

    1997-01-01

    A three-dimensional Navier Stokes code has been used to study the effect of coolant temperature, and coolant to mainstream mass flow ratio on the adiabatic effectiveness of a film-cooled turbine blade. The blade chosen is the VKI rotor with six rows of cooling holes including three rows on the shower head. The mainstream is akin to that under real engine conditions with stagnation temperature = 1900 K and stagnation pressure = 3 MPa. Generally, the adiabatic effectiveness is lower for a higher coolant temperature due to nonlinear effects via the compressibility of air. However, over the suction side of shower-head holes, the effectiveness is higher for a higher coolant temperature than that for a lower coolant temperature when the coolant to mainstream mass flow ratio is 5% or more. For a fixed coolant temperature, the effectiveness passes through a minima on the suction side of shower-head holes as the coolant to mainstream mass flow, ratio increases, while on the pressure side of shower-head holes, the effectiveness decreases with increase in coolant mass flow due to coolant jet lift-off. In all cases, the adiabatic effectiveness is highly three-dimensional.

  19. 14 CFR 29.1109 - Carburetor air cooling.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Carburetor air cooling. 29.1109 Section 29... AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Induction System § 29.1109 Carburetor air... to maintain the air temperature, at the carburetor inlet, at or below the maximum established...

  20. 14 CFR 29.1109 - Carburetor air cooling.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Carburetor air cooling. 29.1109 Section 29... AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Induction System § 29.1109 Carburetor air... to maintain the air temperature, at the carburetor inlet, at or below the maximum established...

  1. 14 CFR 29.1109 - Carburetor air cooling.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Carburetor air cooling. 29.1109 Section 29... AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Induction System § 29.1109 Carburetor air... to maintain the air temperature, at the carburetor inlet, at or below the maximum established...

  2. The effects of leading edge and downstream film cooling on turbine vane heat transfer

    NASA Technical Reports Server (NTRS)

    Hylton, L. D.; Nirmalan, V.; Sultanian, B. K.; Kaufman, R. M.

    1988-01-01

    The progress under contract NAS3-24619 toward the goal of establishing a relevant data base for use in improving the predictive design capabilities for external heat transfer to turbine vanes, including the effect of downstream film cooling with and without leading edge showerhead film cooling. Experimental measurements were made in a two-dimensional cascade previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils under contract NAS3-22761 and leading edge showerhead film cooled airfoils under contract NAS3-23695. The principal independent parameters (Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio) were maintained over ranges consistent with actual engine conditions and the test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. Data provide a data base for downstream film cooled turbine vanes and extends the data bases generated in the two previous studies. The vane external heat transfer obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The data obtained and presented illustrate the interaction of the variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils.

  3. The effects of leading edge and downstream film cooling on turbine vane heat transfer

    NASA Astrophysics Data System (ADS)

    Hylton, L. D.; Nirmalan, V.; Sultanian, B. K.; Kaufman, R. M.

    1988-11-01

    The progress under contract NAS3-24619 toward the goal of establishing a relevant data base for use in improving the predictive design capabilities for external heat transfer to turbine vanes, including the effect of downstream film cooling with and without leading edge showerhead film cooling. Experimental measurements were made in a two-dimensional cascade previously used to obtain vane surface heat transfer distributions on nonfilm cooled airfoils under contract NAS3-22761 and leading edge showerhead film cooled airfoils under contract NAS3-23695. The principal independent parameters (Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio) were maintained over ranges consistent with actual engine conditions and the test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. Data provide a data base for downstream film cooled turbine vanes and extends the data bases generated in the two previous studies. The vane external heat transfer obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The data obtained and presented illustrate the interaction of the variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils.

  4. Passive radiative cooling below ambient air temperature under direct sunlight.

    PubMed

    Raman, Aaswath P; Anoma, Marc Abou; Zhu, Linxiao; Rephaeli, Eden; Fan, Shanhui

    2014-11-27

    Cooling is a significant end-use of energy globally and a major driver of peak electricity demand. Air conditioning, for example, accounts for nearly fifteen per cent of the primary energy used by buildings in the United States. A passive cooling strategy that cools without any electricity input could therefore have a significant impact on global energy consumption. To achieve cooling one needs to be able to reach and maintain a temperature below that of the ambient air. At night, passive cooling below ambient air temperature has been demonstrated using a technique known as radiative cooling, in which a device exposed to the sky is used to radiate heat to outer space through a transparency window in the atmosphere between 8 and 13 micrometres. Peak cooling demand, however, occurs during the daytime. Daytime radiative cooling to a temperature below ambient of a surface under direct sunlight has not been achieved because sky access during the day results in heating of the radiative cooler by the Sun. Here, we experimentally demonstrate radiative cooling to nearly 5 degrees Celsius below the ambient air temperature under direct sunlight. Using a thermal photonic approach, we introduce an integrated photonic solar reflector and thermal emitter consisting of seven layers of HfO2 and SiO2 that reflects 97 per cent of incident sunlight while emitting strongly and selectively in the atmospheric transparency window. When exposed to direct sunlight exceeding 850 watts per square metre on a rooftop, the photonic radiative cooler cools to 4.9 degrees Celsius below ambient air temperature, and has a cooling power of 40.1 watts per square metre at ambient air temperature. These results demonstrate that a tailored, photonic approach can fundamentally enable new technological possibilities for energy efficiency. Further, the cold darkness of the Universe can be used as a renewable thermodynamic resource, even during the hottest hours of the day. PMID:25428501

  5. Air-cooled CWS warm air furnace. Final report

    SciTech Connect

    Litka, A.F.; Becker, F.E.

    1995-08-01

    Thermo Power Corporation, Tecogen Division, has developed coal water slurry (CWS) combustion technologies specifically tailored to meet the space heating needs of the residential, commercial, and industrial market sectors. This furnace was extensively tested and met all the design and operating criteria of the development program, which included combustion efficiencies in excess of 99%, response to full load from a cold start in less than 5 minutes, and steady-state thermal efficiencies as high as 85%. While this furnace design is extremely versatile, versatility came at the expense of system complexity and cost. To provide a more cost effective CWS-based option for the residential market sector, Tecogen, developed a totally air-cooled CWS-fired residential warm air heating system. To minimize system cost and to take advantage of industry manufacturing practices and experience, a commercially available oil/gas solid fuel-fired central furnace, manufactured by Yukon Energy Corporation, was used as the platform for the CWS combustor and related equipment. A prototype furnace was designed, built, and tested in the laboratory to verify system integrity and operation. This unit was then shipped to the PETC to undergo demonstration operation and serve as a showcase of the CWS technology. An in-depth Owners Manual was prepared and delivered with the furnace. This Owners Manual, which is included as Appendix A of this report, includes installation instructions, operating procedures, wiring diagrams, and equipment bulletins on the major components. It also contains coal water slurry fuel specifications and typical system operating variables, including key temperatures, pressures, and flowrates.

  6. Selection of a turbine cooling system applying multi-disciplinary design considerations.

    PubMed

    Glezer, B

    2001-05-01

    The presented paper describes a multi-disciplinary cooling selection approach applied to major gas turbine engine hot section components, including turbine nozzles, blades, discs, combustors and support structures, which maintain blade tip clearances. The paper demonstrates benefits of close interaction between participating disciplines starting from early phases of the hot section development. The approach targets advancements in engine performance and cost by optimizing the design process, often requiring compromises within individual disciplines. PMID:11460630

  7. Demonstration of Enabling Spar-Shell Cooling Technology in Gas Turbines

    SciTech Connect

    Downs, James

    2014-12-29

    In this Advanced Turbine Program-funded Phase III project, Florida Turbine Technologies, Inc. (FTT) has developed and tested, at a pre-commercial prototypescale, spar-shell turbine airfoils in a commercial gas turbine. The airfoil development is based upon FTT’s research and development to date in Phases I and II of Small Business Innovative Research (SBIR) grants. During this program, FTT has partnered with an Original Equipment Manufacturer (OEM), Siemens Energy, to produce sparshell turbine components for the first pre-commercial prototype test in an F-Class industrial gas turbine engine and has successfully completed validation testing. This project will further the commercialization of this new technology in F-frame and other highly cooled turbine airfoil applications. FTT, in cooperation with Siemens, intends to offer the spar-shell vane as a first-tier supplier for retrofit applications and new large frame industrial gas turbines. The market for the spar-shell vane for these machines is huge. According to Forecast International, 3,211 new gas turbines units (in the >50MW capacity size range) will be ordered in ten years from 2007 to 2016. FTT intends to enter the market in a low rate initial production. After one year of successful extended use, FTT will quickly ramp up production and sales, with a target to capture 1% of the market within the first year and 10% within 5 years (2020).

  8. Improving Durability of Turbine Components Through Trenched Film Cooling and Contoured Endwalls

    SciTech Connect

    Bogard, David G.; Thole, Karen A.

    2014-09-30

    The experimental and computational studies of the turbine endwall and vane models completed in this research program have provided a comprehensive understanding of turbine cooling with combined film cooling and TBC. To correctly simulate the cooling effects of TBC requires the use of matched Biot number models, a technique developed in our laboratories. This technique allows for the measurement of the overall cooling effectiveness which is a measure of the combined internal and external cooling for a turbine component. The overall cooling effectiveness provides an indication of the actual metal temperature that would occur at engine conditions, and is hence a more powerful performance indicator than the film effectiveness parameter that is commonly used for film cooling studies. Furthermore these studies include the effects of contaminant depositions which are expected to occur when gas turbines are operated with syngas fuels. Results from the endwall studies performed at Penn State University and the vane model studies performed at the University of Texas are the first direct measurements of the combined effects of film cooling and TBC. These results show that TBC has a dominating effect on the overall cooling effectiveness, which enhances the importance of the internal cooling mechanisms, and downplays the importance of the film cooling of the external surface. The TBC was found to increase overall cooling effectiveness by a factor of two to four. When combined with TBC, the primary cooling from film cooling holes was found to be due to the convective cooling within the holes, not from the film effectiveness on the surface of the TBC. Simulations of the deposition of contaminants on the endwall and vane surfaces showed that these depositions caused a large increase in surface roughness and significant degradation of film effectiveness. However, despite these negative factors, the depositions caused only a slight decrease in the overall cooling effectiveness on

  9. Unsteady High Turbulence Effects on Turbine Blade Film Cooling Heat Transfer Performance Using a Transient Liquid Crystal Technique

    NASA Technical Reports Server (NTRS)

    Han, J. C.; Ekkad, S. V.; Du, H.; Teng, S.

    2000-01-01

    Unsteady wake effect, with and without trailing edge ejection, on detailed heat transfer coefficient and film cooling effectiveness distributions is presented for a downstream film-cooled gas turbine blade. Tests were performed on a five-blade linear cascade at an exit Reynolds number of 5.3 x 10(exp 5). Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. Coolant blowing ratio was varied from 0.4 to 1.2; air and CO2 were used as coolants to simulate different density ratios. Surface heat transfer and film effectiveness distributions were obtained using a transient liquid crystal technique; coolant temperature profiles were determined with a cold wire technique. Results show that Nusselt numbers for a film cooled blade are much higher compared to a blade without film injection. Unsteady wake slightly enhances Nusselt numbers but significantly reduces film effectiveness versus no wake cases. Nusselt numbers increase only slic,htly but film cooling, effectiveness increases significantly with increasing, blowing ratio. Higher density coolant (CO2) provides higher effectiveness at higher blowing ratios (M = 1.2) whereas lower density coolant (Air) provides higher 0 effectiveness at lower blowing ratios (M = 0.8). Trailing edge ejection generally has more effect on film effectiveness than on the heat transfer, typically reducing film effectiveness and enhancing heat transfer. Similar data is also presented for a film cooled cylindrical leading edge model.

  10. Heat pipe cooled twin airfoil blade as an element for higher efficiency of longlife gas turbine

    NASA Astrophysics Data System (ADS)

    Majcen, M.; Sarunac, N.

    The present state of the art in gas turbine engines is closely tied to improvements in design techniques that have resulted, over the years, in a steady increase in operating temperatures. Higher firing temperatures are essential for development of smaller, lighter, more efficient engines. One possible way to meet aforesaid trend, a double gas turbine cycle based on heat pipe cooled twin airfoil blade is described in this paper. The basic and improved flow diagrams of the double gas turbine cycle, its performances, heat transfer analysis on, across and from twin airfoil blade and some calculated examples are presented.

  11. Alternative Liquid Fuel Effects on Cooled Silicon Nitride Marine Gas Turbine Airfoils

    SciTech Connect

    Holowczak, J.

    2002-03-01

    With prior support from the Office of Naval Research, DARPA, and U.S. Department of Energy, United Technologies is developing and engine environment testing what we believe to be the first internally cooled silicon nitride ceramic turbine vane in the United States. The vanes are being developed for the FT8, an aeroderivative stationary/marine gas turbine. The current effort resulted in further manufacturing and development and prototyping by two U.S. based gas turbine grade silicon nitride component manufacturers, preliminary development of both alumina, and YTRIA based environmental barrier coatings (EBC's) and testing or ceramic vanes with an EBC coating.

  12. The new air emission regulations for gas turbine

    SciTech Connect

    Solt, C.

    1998-07-01

    In the US, there are three new regulations now in development that will lower the limits for NO{sub x} emissions from gas turbines: (1) New National Ambient Air Quality Standards (NAAQS) for Particulate Matter, and Possibly revision to the Ozone standard (both of these new programs will target NO{sub x} emissions); (2) New regulations stemming from the Ozone Transport Assessment Group (OTAG) recommendations (again, NO{sub x} is the primary focus); (3) Revision of the New Source Performance Standard (NSPS) for gas turbines and a new rule that will impose new toxic emission requirements, (the Industrial Combustion Coordinated Rulemaking, stemming from revisions to Title III of the Clean Sir Act Amendments of 1990). The toxic rule should be of particular concern to the gas turbine industry in that it may impose the use of expensive toxic emission control techniques that may not provide any significant health benefits to the public. In addition, the European Community is currently drafting a new regulation for combustion sources that will require gas turbines to meet levels that are lower than any in Europe today. This paper will consider all 5 of these regulatory actions and will: review the proposed regulations; discuss timing for regulation development and implementation; assess the probable impact of each regulation; and provide opinions on the fate of each regulation. Both manufacturers and users of gas turbines should be aware of these proceedings and take an active role in the rule development.

  13. Heat-transfer processes in air-cooled engine cylinders

    NASA Technical Reports Server (NTRS)

    Pinkel, Benjamin

    1938-01-01

    From a consideration of heat-transfer theory, semi-empirical expressions are set up for the transfer of heat from the combustion gases to the cylinder of an air-cooled engine and from the cylinder to the cooling air. Simple equations for the average head and barrel temperatures as functions of the important engine and cooling variables are obtained from these expressions. The expressions involve a few empirical constants, which may be readily determined from engine tests. Numerical values for these constants were obtained from single-cylinder engine tests for cylinders of the Pratt & Whitney 1535 and 1340-h engines. The equations provide a means of calculating the effect of the various engine and cooling variables on the cylinder temperatures and also of correlating the results of engine cooling tests. An example is given of the application of the equations to the correlation of cooling-test data obtained in flight.

  14. 21 CFR 211.46 - Ventilation, air filtration, air heating and cooling.

    Code of Federal Regulations, 2011 CFR

    2011-04-01

    ... 21 Food and Drugs 4 2011-04-01 2011-04-01 false Ventilation, air filtration, air heating and... Buildings and Facilities § 211.46 Ventilation, air filtration, air heating and cooling. (a) Adequate ventilation shall be provided. (b) Equipment for adequate control over air pressure, micro-organisms,...

  15. 21 CFR 211.46 - Ventilation, air filtration, air heating and cooling.

    Code of Federal Regulations, 2014 CFR

    2014-04-01

    ... 21 Food and Drugs 4 2014-04-01 2014-04-01 false Ventilation, air filtration, air heating and... Buildings and Facilities § 211.46 Ventilation, air filtration, air heating and cooling. (a) Adequate ventilation shall be provided. (b) Equipment for adequate control over air pressure, micro-organisms,...

  16. 21 CFR 211.46 - Ventilation, air filtration, air heating and cooling.

    Code of Federal Regulations, 2012 CFR

    2012-04-01

    ... 21 Food and Drugs 4 2012-04-01 2012-04-01 false Ventilation, air filtration, air heating and... Buildings and Facilities § 211.46 Ventilation, air filtration, air heating and cooling. (a) Adequate ventilation shall be provided. (b) Equipment for adequate control over air pressure, micro-organisms,...

  17. 21 CFR 211.46 - Ventilation, air filtration, air heating and cooling.

    Code of Federal Regulations, 2010 CFR

    2010-04-01

    ... 21 Food and Drugs 4 2010-04-01 2010-04-01 false Ventilation, air filtration, air heating and... Buildings and Facilities § 211.46 Ventilation, air filtration, air heating and cooling. (a) Adequate ventilation shall be provided. (b) Equipment for adequate control over air pressure, micro-organisms,...

  18. Computation of the temperature distribution in cooled radial inflow turbine guide vanes

    NASA Technical Reports Server (NTRS)

    Tabakoff, W.; Hosny, W.; Hamed, A.

    1977-01-01

    A two-dimensional finite-difference numerical technique is presented to determine the temperature distribution of an internally-cooled blade of radial turbine guide vanes. A simple convection cooling is assumed inside the guide vane. Such an arrangement results in relatively small cooling effectiveness at the leading edge and at the trailing edge. Heat transfer augmentation in these critical areas may be achieved by using impingement jets and film cooling. A computer program is written in Fortran IV for IBM 370/165 computer.

  19. Side wall cooling for nozzle segments for a gas turbine

    DOEpatents

    Burdgick, Steven Sebastian

    2002-01-01

    A nozzle vane segment includes outer and inner band portions with a vane extending therebetween and defining first and second cavities separated by an impingement plate for flowing cooling medium for impingement cooling of nozzle side walls. The side wall of each nozzle segment has an undercut region. The impingement plate has an inturned flange with a plurality of openings. Cooling inserts or receptacles having an open end are received in the openings and the base and side walls of the receptacles have apertures for receiving cooling medium from the first cavity and directing the cooling medium for impingement cooling of the side wall of the nozzle segment and a portion of the nozzle wall.

  20. Numerical simulation of heat transfer performance of an air-cooled steam condenser in a thermal power plant

    NASA Astrophysics Data System (ADS)

    Gao, Xiufeng; Zhang, Chengwei; Wei, Jinjia; Yu, Bo

    2009-09-01

    Numerical simulation of the thermal-flow characteristics and heat transfer performance is made of an air-cooled steam condenser (ACSC) in a thermal power plant by considering the effects of ambient wind speed and direction, air-cooled platform height, location of the main factory building and terrain condition. A simplified physical model of the ACSC combined with the measured data as input parameters is used in the simulation. The wind speed effects on the heat transfer performance and the corresponding steam turbine back pressure for different heights of the air-cooled platform are obtained. It is found that the turbine back pressure (absolute pressure) increases with the increase of wind speed and the decrease of platform height. This is because wind can not only reduce the flowrate in the axial fans, especially at the periphery of the air-cooled platform, due to cross-flow effects, but also cause an air temperature increase at the fan inlet due to hot air recirculation, resulting in the deterioration of the heat transfer performance. The hot air recirculation is found to be the dominant factor because the main factory building is situated on the windward side of the ACSC.

  1. Effect of rotation on heat transfer and hydraulic resistance in the radial cooling channels of turbine rotor blades

    NASA Astrophysics Data System (ADS)

    Iskakov, K. M.; Trushin, O. V.; Tsaplin, M. I.; Shatalov, Yu. S.

    Results of a modeling study indicate that rotation significantly (up to 60 percent) changes local heat transfer and increases, by a factor of 5-6, hydraulic resistance in the smooth radial channels of turbine rotor blades with a low-pressure cooling system. The results of the study have been used in the design of a turbine cooling system for a turbofan engine.

  2. Experimental research of reaction blading on air turbine VT-400

    NASA Astrophysics Data System (ADS)

    Klimko, Marek; Okresa, Daniel

    2016-03-01

    The article deals with testing a reaction blading on an experimental air turbine VT-400, which is situated at the Department of Power System Engineering at University of West Bohemia. Experiments were carried out in cooperation with an industrial partner Doosan Skoda Power. The outputs of these measurements are for example: results of the stage efficiency depending on the speed ratio u/c, the course of reaction, the input and output angles and profile losses along a blade.

  3. Curved centerline air intake for a gas turbine engine

    NASA Technical Reports Server (NTRS)

    Ruehr, W. C.; Younghans, J. L.; Smith, E. B. (Inventor)

    1980-01-01

    An inlet for a gas turbine engine was disposed about a curved centerline for the purpose of accepting intake air that is flowing at an angle to engine centerline and progressively turning that intake airflow along a curved path into alignment with the engine. This curved inlet is intended for use in under the wing locations and similar regions where airflow direction is altered by aerodynamic characteristics of the airplane. By curving the inlet, aerodynamic loss and acoustic generation and emission are decreased.

  4. Hot streaks and phantom cooling in a turbine rotor passage. I - Separate effects

    NASA Technical Reports Server (NTRS)

    Roback, Richard J.; Dring, Robert P.

    1992-01-01

    Experimental documentation and analytical correlations demonstrating the effects of hot streak accumulation and phantom cooling on turbine rotor airfoil surface temperature. Test results are shown for a range of controlling variables to identify where streak accumulation and phantom cooling are most likely to be strongest. These variables include streak injection location, streak-to-free stream density ratio and coolant-to-free stream density and velocity ratios.

  5. A numerical study of the temperature field in a cooled radial turbine rotor

    NASA Technical Reports Server (NTRS)

    Hamed, A.; Baskharone, E.; Tabakoff, W.

    1977-01-01

    The three dimensional temperature distribution in the cooled rotor of a radial inflow turbine is determined numerically using the finite element method. Through this approach, the complicated geometries of the hot rotor and coolant passage surfaces are handled easily, and the temperatures are determined without loss of accuracy at these convective boundaries. Different cooling techniques with given coolant to primary flow ratios are investigated, and the corresponding rotor temperature fields are presented for comparison.

  6. Turbomachine rotor with improved cooling

    DOEpatents

    Hultgren, K.G.; McLaurin, L.D.; Bertsch, O.L.; Lowe, P.E.

    1998-05-26

    A gas turbine rotor has an essentially closed loop cooling air scheme in which cooling air drawn from the compressor discharge air that is supplied to the combustion chamber is further compressed, cooled, and then directed to the aft end of the turbine rotor. Downstream seal rings attached to the downstream face of each rotor disc direct the cooling air over the downstream disc face, thereby cooling it, and then to cooling air passages formed in the rotating blades. Upstream seal rings attached to the upstream face of each disc direct the heated cooling air away from the blade root while keeping the disc thermally isolated from the heated cooling air. From each upstream seal ring, the heated cooling air flows through passages in the upstream discs and is then combined and returned to the combustion chamber from which it was drawn. 5 figs.

  7. Turbomachine rotor with improved cooling

    DOEpatents

    Hultgren, Kent Goran; McLaurin, Leroy Dixon; Bertsch, Oran Leroy; Lowe, Perry Eugene

    1998-01-01

    A gas turbine rotor has an essentially closed loop cooling air scheme in which cooling air drawn from the compressor discharge air that is supplied to the combustion chamber is further compressed, cooled, and then directed to the aft end of the turbine rotor. Downstream seal rings attached to the downstream face of each rotor disc direct the cooling air over the downstream disc face, thereby cooling it, and then to cooling air passages formed in the rotating blades. Upstream seal rings attached to the upstream face of each disc direct the heated cooling air away from the blade root while keeping the disc thermally isolated from the heated cooling air. From each upstream seal ring, the heated cooling air flows through passages in the upstream discs and is then combined and returned to the combustion chamber from which it was drawn.

  8. Split ring floating air riding seal for a turbine

    DOEpatents

    Mills, Jacob A

    2015-11-03

    A floating air riding seal for a gas turbine engine with a rotor and a stator, an annular piston chamber with an axial moveable annular piston assembly within the annular piston chamber, an annular cavity formed on the annular piston assembly that faces a seal surface on the rotor, and a central passage connecting the annular cavity to the annular piston chamber to supply compressed air to the seal face, where the annular piston assembly is a split piston assembly to maintain a tight seal as coning of the rotor disk occurs.

  9. Cool Colored Roofs to Save Energy and Improve Air Quality

    SciTech Connect

    Akbari, Hashem; Levinson, Ronnen; Miller, William; Berdahl, Paul

    2005-08-23

    Urban areas tend to have higher air temperatures than their rural surroundings as a result of gradual surface modifications that include replacing the natural vegetation with buildings and roads. The term ''Urban Heat Island'' describes this phenomenon. The surfaces of buildings and pavements absorb solar radiation and become extremely hot, which in turn warm the surrounding air. Cities that have been ''paved over'' do not receive the benefit of the natural cooling effect of vegetation. As the air temperature rises, so does the demand for air-conditioning (a/c). This leads to higher emissions from power plants, as well as increased smog formation as a result of warmer temperatures. In the United States, we have found that this increase in air temperature is responsible for 5-10% of urban peak electric demand for a/c use, and as much as 20% of population-weighted smog concentrations in urban areas. Simple ways to cool the cities are the use of reflective surfaces (rooftops and pavements) and planting of urban vegetation. On a large scale, the evapotranspiration from vegetation and increased reflection of incoming solar radiation by reflective surfaces will cool a community a few degrees in the summer. As an example, computer simulations for Los Angeles, CA show that resurfacing about two-third of the pavements and rooftops with reflective surfaces and planting three trees per house can cool down LA by an average of 2-3K. This reduction in air temperature will reduce urban smog exposure in the LA basin by roughly the same amount as removing the basin entire onroad vehicle exhaust. Heat island mitigation is an effective air pollution control strategy, more than paying for itself in cooling energy cost savings. We estimate that the cooling energy savings in U.S. from cool surfaces and shade trees, when fully implemented, is about $5 billion per year (about $100 per air-conditioned house).

  10. Vertical air circulation in a low-speed lateral flow wind turbine with rotary blades

    NASA Astrophysics Data System (ADS)

    Cheboxarov, Vik. V.; Cheboxarov, Val. V.

    2008-01-01

    The model of a large-scale lateral flow wind turbine with rotary blades is presented and the conditions of numerical aerodynamic investigation of this turbine are described. The results of numerical experiments show that air flowing past the turbine exhibits a considerable vertical (axial) circulation, which increases the power coefficient of the turbine. In the inner space of the turbine, two stable vortices are formed through which retarded streams partly leave the turbine upon flowing past the windward side, to be replaced by faster streams from adjacent layers of air.

  11. The heat transfer of cooling fins on moving air

    NASA Technical Reports Server (NTRS)

    Doetsch, Hans

    1935-01-01

    The present report is a comparison of the experimentally defined temperature and heat output of cooling fins in the air stream with theory. The agreement is close on the basis of a mean coefficient of heat transfer with respect to the total surface. A relationship is established between the mean coefficient of heat transfer, the dimensions of the fin arrangement, and the air velocity.

  12. Synthetic optimization of air turbine for dental handpieces.

    PubMed

    Shi, Z Y; Dong, T

    2014-01-01

    A synthetic optimization of Pelton air turbine in dental handpieces concerning the power output, compressed air consumption and rotation speed in the mean time is implemented by employing a standard design procedure and variable limitation from practical dentistry. The Pareto optimal solution sets acquired by using the Normalized Normal Constraint method are mainly comprised of two piecewise continuous parts. On the Pareto frontier, the supply air stagnation pressure stalls at the lower boundary of the design space, the rotation speed is a constant value within the recommended range from literature, the blade tip clearance insensitive to while the nozzle radius increases with power output and mass flow rate of compressed air to which the residual geometric dimensions are showing an opposite trend within their respective "pieces" compared to the nozzle radius. PMID:25571069

  13. Gas turbine rotor blade film cooling with and without simulated NGV shock waves and wakes

    NASA Astrophysics Data System (ADS)

    Rigby, M. J.; Johnson, A. B.; Oldfield, M. L. G.

    1990-06-01

    Detailed heat transfer measurements have been made around a film-cooled transonic gas turbine rotor blade in the Oxford Isentropic Light Piston Tunnel. Film cooling behavior for four film cooling configurations have been analyzed for a range of blowing rates both without and with simulated nozzle guide vane (NGV) shock wave and wake passing. The superposition model of film cooling has been employed in analysis of time-mean heat transfer data, while time resolved unsteady heat transfer measurements have been analyzed to determine interaction between film-cooling and unsteady shock wave and wake passing. It is found that there is a significant change of film-cooling behavior on the suction surface when simulated NGV unsteady effects are introduced.

  14. Computer code for the calculation of the temperature distribution of cooled turbine blades

    NASA Astrophysics Data System (ADS)

    Tietz, Thomas A.; Koschel, Wolfgang W.

    A generalized computer code for the calculation of the temperature distribution in a cooled turbine blade is presented. Using an iterative procedure, this program especially allows the coupling of the aerothermodynamic values of the internal flow with the corresponding temperature distribution of the blade material. The temperature distribution of the turbine blade is calculated using a fully three-dimensional finite element computer code, so that the radial heat flux is taken into account. This code was extended to 4-node tetrahedral elements enabling an adaptive grid generation. To facilitate the mesh generation of the usually complex blade geometries, a computer program was developed, which performs the grid generation of blades having basically arbitrary shape on the basis of two-dimensional cuts. The performance of the code is demonstrated with reference to a typical cooling configuration of a modern turbine blade.

  15. Water droplet evaporation in air during compression in a gas turbine engine. Technical memo

    SciTech Connect

    Quandt, E.

    1996-04-01

    A water fog concept is being considered for evaporative cooling of the air as it is compressed in a ship gas turbine engine. The following analysis is presented to clarify the physics associated with liquid droplet evaporation in this situation, to understand the conditions affecting the cooling, and to identify any further information required to achieve such a concept. The vaporization of small liquid drops in a warm ideal gas is controlled by the outward motion of the vapor and the inward flow of heat to cause evaporation. Following the standard analysis of Spalding, as given in `Principles of Combustion` by Kuo, it is assumed that the process is `quasi steady.` This means that the conditions far removed from the drop are constant, and that there are no time varying terms in the Eulerian description of the mass and energy flows.

  16. Varying duty operation of air-cooled condenser units

    NASA Astrophysics Data System (ADS)

    Milman, O. O.; Kondratev, A. V.; Ptakhin, A. V.; Dunaev, S. N.; Kirjukhin, A. V.

    2016-05-01

    Results of experimental investigations of operation modes of air-cooled condensers (ACC) under design and varying duty conditions are presented. ACCs with varying cooling airflow rates under constant heat load and with constant cooling airflow under varying heat load are examined. Diagrams of heat transfer coefficients and condensation pressures on the heat load and cooling airflow are obtained. It is found that, if the relative heat load is in the range from 0.6 to 1.0 of the nominal value, the ACC heat transfer coefficient varies insignificantly, unlike that of the water-cooled surface condensers. The results of the determination of "zero points" are given, i.e., the attainable pressure in air-cooled condensing units (ACCU), if there is no heat load for several values of working water temperature at the input of water-jet ejectors and liquid ring vacuum pump. The results of the experimental determination of atmospheric air suction into the ACC vacuum system. The effect of additional air suctions in the steam pipe on ACCU characteristics is analyzed. The thermal mapping of ACC heat exchange surfaces from the cooling air inlet is carried out. The dependence of the inefficient heat exchange zone on the additional air suction into the ACC vacuum system is given. It is shown that, if there is no additional air suction into the ACC vacuum system, the inefficient heat exchange zone is not located at the bottom of the first pass tubes, and their portion adjacent to the bottom steam pipe works efficiently. Design procedures for the ACC varying duty of capacitors are presented, and their adequacy for the ACCU varying duty estimation is analyzed.

  17. Computer-automated multi-disciplinary analysis and design optimization of internally cooled turbine blades

    NASA Astrophysics Data System (ADS)

    Martin, Thomas Joseph

    This dissertation presents the theoretical methodology, organizational strategy, conceptual demonstration and validation of a fully automated computer program for the multi-disciplinary analysis, inverse design and optimization of convectively cooled axial gas turbine blades and vanes. Parametric computer models of the three-dimensional cooled turbine blades and vanes were developed, including the automatic generation of discretized computational grids. Several new analysis programs were written and incorporated with existing computational tools to provide computer models of the engine cycle, aero-thermodynamics, heat conduction and thermofluid physics of the internally cooled turbine blades and vanes. A generalized information transfer protocol was developed to provide the automatic mapping of geometric and boundary condition data between the parametric design tool and the numerical analysis programs. A constrained hybrid optimization algorithm controlled the overall operation of the system and guided the multi-disciplinary internal turbine cooling design process towards the objectives and constraints of engine cycle performance, aerodynamic efficiency, cooling effectiveness and turbine blade and vane durability. Several boundary element computer programs were written to solve the steady-state non-linear heat conduction equation inside the internally cooled and thermal barrier-coated turbine blades and vanes. The boundary element method (BEM) did not require grid generation inside the internally cooled turbine blades and vanes, so the parametric model was very robust. Implicit differentiations of the BEM thermal and thereto-elastic analyses were done to compute design sensitivity derivatives faster and more accurately than via explicit finite differencing. A factor of three savings of computer processing time was realized for two-dimensional thermal optimization problems, and a factor of twenty was obtained for three-dimensional thermal optimization problems

  18. A high pressure, high temperature combustor and turbine-cooling test facility

    NASA Technical Reports Server (NTRS)

    Cochran, R. P.; Norris, J. W.

    1976-01-01

    A new test facility is being constructed for developing turbine-cooling and combustor technology for future generation aircraft gas turbine engines. Prototype engine hardware will be investigated in this new facility at gas stream conditions up to 2480 K average turbine inlet temperature and 4.14 x 10 to the 6th power n sq m turbine inlet pressure. The facility will have the unique feature of fully automated control and data acquisition through the use of an integrated system of mini-computers and programmable controllers which will result in more effective use of operating time, will limit the number of operators required, and will provide built in self protection safety systems. The facility and the planning and design considerations are described.

  19. An air bearing system for small high speed gas turbines

    NASA Astrophysics Data System (ADS)

    Turner, A. B.; Davies, S. J.; Nimir, Y. L.

    1994-03-01

    This paper describes the second phase of an experimental program concerning the application of air bearings to small turbomachinery test rigs and small gas turbines. The first phase examined externally pressurized (EP) journal bearings, with a novel EP thrust bearing, for application to 'warm air' test rigs, and was entirely successful at rotational speeds in excess of 100,000 rpm. This second phase examined several designs of tilting pad-spiring journal bearings, one with a novel form of externally pressurized pad, but all using the original EP thrust bearing. The designs tested are described, including some oscillogram traces, for tests up to a maximum of 70,000 rpm; the most successful using a carbon pad-titanium beam spring arrangement. The thrust bearing which gave trouble-free operation throughout, is also described. The results of an original experiment to measure the 'runway speed' of a radial inflow turbine are also presented, which show that overspeeds of 58 percent above the design speed can result from free-power turbine coupling failure.

  20. Glenn-HT Code Validated for Complex Turbine Blade Cooling Passage

    NASA Technical Reports Server (NTRS)

    Rigby, David L.

    2003-01-01

    This work is motivated by the need to accurately predict heat transfer in turbomachinery. For efficient gas turbine operation, flow temperatures in the hot gas path exceed acceptable metal temperatures in many regions of the engine. So that the integrity of the parts can be maintained for an acceptable engine life, the parts must be cooled. Efficient cooling schemes require accurate heat transfer prediction to minimize regions that are overcooled and, even more importantly, to ensure adequate cooling in high-heat-flux regions.

  1. Unsteady, Cooled Turbine Simulation Using a PC-Linux Analysis System

    NASA Technical Reports Server (NTRS)

    List, Michael G.; Turner, Mark G.; Chen, Jen-Pimg; Remotigue, Michael G.; Veres, Joseph P.

    2004-01-01

    The fist stage of the high-pressure turbine (HPT) of the GE90 engine was simulated with a three-dimensional unsteady Navier-Sokes solver, MSU Turbo, which uses source terms to simulate the cooling flows. In addition to the solver, its pre-processor, GUMBO, and a post-processing and visualization tool, Turbomachinery Visual3 (TV3) were run in a Linux environment to carry out the simulation and analysis. The solver was run both with and without cooling. The introduction of cooling flow on the blade surfaces, case, and hub and its effects on both rotor-vane interaction as well the effects on the blades themselves were the principle motivations for this study. The studies of the cooling flow show the large amount of unsteadiness in the turbine and the corresponding hot streak migration phenomenon. This research on the GE90 turbomachinery has also led to a procedure for running unsteady, cooled turbine analysis on commodity PC's running the Linux operating system.

  2. Thermodynamic model of the expansion process in a cooled gas turbine

    NASA Astrophysics Data System (ADS)

    Romakhova, G. A.

    2015-02-01

    A method for calculating the expansion process in a gas turbine with open cooling, which is represented as a thermodynamic system with a variable mass consisting of two subsystems (a gas subsystem with a constant mass and a coolant subsystem with a variable mass) is proposed. An analytical solution of the system of equations describing the expansion process is presented. The accuracy and validity of the obtained results are demonstrated. This method allows one to analytically calculate the expansion process parameters (including the losses caused by cooling). The results of calculation of these losses for various values of the cooling system efficiency and the gas temperature at the turbine inlet are given. The exergy analysis of losses in a setup with a cooled turbine is performed. It is demonstrated that the cooling losses may amount to 5% (or more) of the fuel energy and may not be compensated by the waste-heat loop efficiency in a combined-cycle unit. The obtained results are compared with the published data on the GTE-65 unit produced by Silovye Mashiny.

  3. Development of a thermal and structural analysis procedure for cooled radial turbines

    NASA Astrophysics Data System (ADS)

    Kumar, Ganesh N.; Deanna, Russell G.

    1988-06-01

    A procedure for computing the rotor temperature and stress distributions in a cooled radial turbine is considered. Existing codes for modeling the external mainstream flow and the internal cooling flow are used to compute boundary conditions for the heat transfer and stress analyses. An inviscid, quasi three-dimensional code computes the external free stream velocity. The external velocity is then used in a boundary layer analysis to compute the external heat transfer coefficients. Coolant temperatures are computed by a viscous one-dimensional internal flow code for the momentum and energy equation. These boundary conditions are input to a three-dimensional heat conduction code for calculation of rotor temperatures. The rotor stress distribution may be determined for the given thermal, pressure and centrifugal loading. The procedure is applied to a cooled radial turbine which will be tested at the NASA Lewis Research Center. Representative results from this case are included.

  4. Development of a thermal and structural analysis procedure for cooled radial turbines

    NASA Technical Reports Server (NTRS)

    Kumar, Ganesh N.; Deanna, Russell G.

    1988-01-01

    A procedure for computing the rotor temperature and stress distributions in a cooled radial turbine are considered. Existing codes for modeling the external mainstream flow and the internal cooling flow are used to compute boundary conditions for the heat transfer and stress analysis. The inviscid, quasi three dimensional code computes the external free stream velocity. The external velocity is then used in a boundary layer analysis to compute the external heat transfer coefficients. Coolant temperatures are computed by a viscous three dimensional internal flow cade for the momentum and energy equation. These boundary conditions are input to a three dimensional heat conduction code for the calculation of rotor temperatures. The rotor stress distribution may be determined for the given thermal, pressure and centrifugal loading. The procedure is applied to a cooled radial turbine which will be tested at the NASA Lewis Research Center. Representative results are given.

  5. Development of a thermal and structural analysis procedure for cooled radial turbines

    NASA Technical Reports Server (NTRS)

    Kumar, Ganesh N.; Deanna, Russell G.

    1988-01-01

    A procedure for computing the rotor temperature and stress distributions in a cooled radial turbine is considered. Existing codes for modeling the external mainstream flow and the internal cooling flow are used to compute boundary conditions for the heat transfer and stress analyses. An inviscid, quasi three-dimensional code computes the external free stream velocity. The external velocity is then used in a boundary layer analysis to compute the external heat transfer coefficients. Coolant temperatures are computed by a viscous one-dimensional internal flow code for the momentum and energy equation. These boundary conditions are input to a three-dimensional heat conduction code for calculation of rotor temperatures. The rotor stress distribution may be determined for the given thermal, pressure and centrifugal loading. The procedure is applied to a cooled radial turbine which will be tested at the NASA Lewis Research Center. Representative results from this case are included.

  6. Methods and apparatus for cooling wind turbine generators

    DOEpatents

    Salamah, Samir A.; Gadre, Aniruddha Dattatraya; Garg, Jivtesh; Bagepalli, Bharat Sampathkumaran; Jansen, Patrick Lee; Carl, Jr., Ralph James

    2008-10-28

    A wind turbine generator includes a stator having a core and a plurality of stator windings circumferentially spaced about a generator longitudinal axis. A rotor is rotatable about the generator longitudinal axis, and the rotor includes a plurality of magnetic elements coupled to the rotor and cooperating with the stator windings. The magnetic elements are configured to generate a magnetic field and the stator windings are configured to interact with the magnetic field to generate a voltage in the stator windings. A heat pipe assembly thermally engaging one of the stator and the rotor to dissipate heat generated in the stator or rotor.

  7. Cold-air performance of compressor-drive turbine of department of energy upgraded automobile gas turbine engine. 3: Performance of redesigned turbine

    NASA Technical Reports Server (NTRS)

    Roelke, R. J.; Haas, J. E.

    1984-01-01

    The aerodynamic performance of a redesigned compressor drive turbine of the gas turbine engine is determined in air at nominal inlet conditions of 325 K and 0.8 bar absolute. The turbine is designed with a lower flow factor, higher rotor reaction and a redesigned inlet volute compared to the first turbine. Comparisons between this turbine and the originally designed turbine show about 2.3 percentage points improvement in efficiency at the same rotor tip clearance. Two versions of the same rotor are tested: (1) an as cast rotor, and (2) the same rotor with reduced surface roughness. The effect of reducing surface roughness is about one half percentage point improvement in efficiency. Tests made to determine the effect of Reynolds number on the turbine performance show no effect for the range from 100,000 to 500,000.

  8. Impact of rotor-stator interaction on turbine blade film cooling

    SciTech Connect

    Abhari, R.S.

    1996-01-01

    The goal of this study is to quantify the impact of rotor-stator interaction on surface heat transfer of film cooled turbine blades. In Section 1, a steady-state injection model of the film cooling is incorporated into a two-dimensional, thin shear layer, multiblade row CFD code. This injection model accounts for the penetration and spreading of the coolant jet, as well as the entrainment of the boundary layer fluid by the coolant. The code is validated, in the steady state, by comparing its predictions to data from a blade tested in linear cascade. In Section 2, time-resolved film cooled turbine rotor heat transfer measurements are compared with numerical predictions. Data were taken on a fully film cooled blade in a transonic, high pressure ratio, single-stage turbine in a short duration turbine test facility, which simulates full-engine non-dimensional conditions. Film cooled heat flux on the pressure surface is predicted to be as much as a factor of two higher in the time average of the unsteady calculations compared to the steady-state case. Time-resolved film cooled heat transfer comparison of data to prediction at two spanwise positions is used to validate the numerical code. The unsteady stator-rotor interaction results in the pulsation of the coolant injection flow out of the film holes with large-scale fluctuations. The combination of pulsating coolant flow and the interaction of the coolant with this unsteady external flow is shown to lower the local pressure side adiabatic film effectiveness by as much as 64% when compared to the steady-state case.

  9. Numerical solution for the temperature distribution in a cooled guide vane blade of a radial gas turbine

    NASA Technical Reports Server (NTRS)

    Hosny, W. M.; Tabakoff, W.

    1977-01-01

    A two dimensional finite difference numerical technique is presented to determine the temperature distribution of an internal cooled blade of radial turbine guide vanes. A simple convection cooling is assumed inside the guide vane blade. Such cooling has relatively small cooling effectiveness at the leading edge and at the trailing edge. Heat transfer augmentation in these critical areas may be achieved by using impingement jets and film cooling. A computer program is written in FORTRAN IV for IBM 370/165 computer.

  10. Air conditioning system with supplemental ice storing and cooling capacity

    DOEpatents

    Weng, Kuo-Lianq; Weng, Kuo-Liang

    1998-01-01

    The present air conditioning system with ice storing and cooling capacity can generate and store ice in its pipe assembly or in an ice storage tank particularly equipped for the system, depending on the type of the air conditioning system. The system is characterized in particular in that ice can be produced and stored in the air conditioning system whereby the time of supplying cooled air can be effectively extended with the merit that the operation cycle of the on and off of the compressor can be prolonged, extending the operation lifespan of the compressor in one aspect. In another aspect, ice production and storage in great amount can be performed in an off-peak period of the electrical power consumption and the stored ice can be utilized in the peak period of the power consumption so as to provide supplemental cooling capacity for the compressor of the air conditioning system whereby the shift of peak and off-peak power consumption can be effected with ease. The present air conditioning system can lower the installation expense for an ice-storing air conditioning system and can also be applied to an old conventional air conditioning system.

  11. Cooling of Gas Turbines. 2; Effectiveness of Rim Cooling of Blades

    NASA Technical Reports Server (NTRS)

    Wolfenstein, Lincoln; Meyer, Gene L.; McCarthy, John S.

    1945-01-01

    An analysis of rim cooling, which cools the blade by condition alone, was conducted. Gas temperatures ranged from 1300 degrees to 1900 degrees F and rim temperatures from 0 degrees to 1000 degrees F below gas temperatures. Results show that gas temperature increases up to 200 degrees F are permissible provided that the blades are cooled by 400 degrees to 500 degrees F below the gas temperature. Relatively small amounts of blade cooling, at constant gas temperature, give large increases in blade life. Dependence of rim cooling on heat-transfer coefficient, blade dimensions, and thermal conductivity is determined by a single parameter.

  12. An investigation of ingress for an 'air-cooled' shrouded rotating disk system with radial-clearance seals

    NASA Astrophysics Data System (ADS)

    Phadke, U. P.; Owen, J. M.

    1982-04-01

    The quest for improved performance has led to great interest in the study of disk sealing and cooling air systems of gas turbines. The disk cooling air must not only remove the heat conducted in the disk from the blades but must also prevent the ingress of hot gas into the cavity between the disk and the stator. The present investigation is concerned with the study of several different rotor-stator seals with radial clearances between cylindrical shrouds on both the rotor and the stator. The tests were conducted in the absence of an external axial flow, which occurs in an actual gas turbine. Flow visualization and pressure measurements were used to study the performance of the radial-clearance seals.

  13. An experimental study of turbine vane heat transfer with leading edge and downstream film cooling

    NASA Technical Reports Server (NTRS)

    Nirmalan, V.; Hylton, L. D.

    1989-01-01

    This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine-vane external heat transfer. Steady-state experimental measurements were made in a three-vane linear two-dimensional cascade. The principal independent parameters were maintained over ranges consistent with actual engine conditions. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. The data obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil.

  14. An experimental study of turbine vane heat transfer with leading edge and downstream film cooling

    NASA Astrophysics Data System (ADS)

    Nirmalan, V.; Hylton, L. D.

    1989-06-01

    This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine-vane external heat transfer. Steady-state experimental measurements were made in a three-vane linear two-dimensional cascade. The principal independent parameters were maintained over ranges consistent with actual engine conditions. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. The data obtained indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil.

  15. Cooling System Using Natural Circulation for Air Conditioning

    NASA Astrophysics Data System (ADS)

    Okazaki, Takashi; Seshimo, Yu

    In this paper, Cooling systems with natural circulation loop of refrigerants are reviewed. The cooling system can largely reduce energy consumption of a cooling system for the telecommunication base site. The cooling system consists of two refrigeration units; vapor compression refrigeration unit and sub-cooling unit with a natural-circulation loop. The experiments and calculations were carried out to evaluate the cycle performance of natural circulation loop with HFCs and CO2. The experimental results showed that the cooling capacity of R410A is approximately 30% larger than that of R407C at the temperature difference of 20K and the cooling capacity of CO2 was approximately 4-13% larger than that of R410A under the two-phase condition. On the other hand, the cooling capacity of CO2 was approximately 11% smaller than that of R410A under the supercritical condition. The cooling capacity took a maximum value at an amount of refrigerant and lineally increased as the temperature difference increases and the slightly increased as the height difference. The air intake temperature profile in the inlet of the heat exchangers makes the reverse circulation under the supercritical state and the driving head difference for the reverse circulation depends on the density change to temperature under the supercritical state. Also, a new fan control method to convert the reverse circulation into the normal circulation was reviewed.

  16. A numerical procedure to determine blade temperature of cooled stator blades in gas turbines: A numerical and experimental comparison

    SciTech Connect

    Carcasci, C.; Facchini, B.; Corradini, U.

    1994-12-31

    The continuing need to improve both efficiency and specific power of gas turbines requires progressively increasing turbine inlet temperatures and blade cooling efficiency. This paper presents a blade cooling simulation code for gas turbine stator blades analysis. A comparison between the predictions of simulation and the experimental data of the external nozzle temperatures of a small heavy duty gas turbine is presented. The simulation code allows for several cooling techniques on the basis of different solutions combined (as they often are) on the same blade. The cooling model considered is simplified: the limitation in the accuracy of the results is far overcome by the simplicity and versatility of the approach. A two-dimensional Navier Stokes simulation of film cooling is also compared with a correlation approach.

  17. Potential use of ceramic coating as a thermal insulation on cooled turbine hardware

    NASA Technical Reports Server (NTRS)

    Liebert, C. H.; Stepka, F. S.

    1976-01-01

    An analysis was made to determine the potential benefits of using a ceramic thermal insulation coating of calcia-stabilized zirconia on cooled engine parts. The analysis was applied to turbine vanes of a high temperature and high pressure core engine and a moderate temperature and low pressure research engine. Measurements made during engine operation showed that the coating substantially reduced vane metal wall temperatures. Evaluation of the durability of the coating on turbine vanes and blades in a furnace and engine were encouraging.

  18. Cooling system for a gas turbine using a cylindrical insert having V-shaped notch weirs

    DOEpatents

    Grondahl, Clayton M.; Germain, Malcolm R.

    1981-01-01

    An improved cooling system for a gas turbine is disclosed. A plurality of V-shaped notch weirs are utilized to meter a coolant liquid from a pool of coolant into a plurality of platform and airfoil coolant channels formed in the buckets of the turbine. The V-shaped notch weirs are formed in a separately machined cylindrical insert and serve to desensitize the flow of coolant into the individual platform and airfoil coolant channels to design tolerances and non-uniform flow distribution.

  19. Cooling/heating augmentation during turbine startup/shutdown using a seal positioned by thermal response of turbine parts and consequent relative movement thereof

    DOEpatents

    Schmidt, Mark Christopher

    2000-01-01

    In a turbine rotor, a thermal mismatch between various component parts of the rotor occurs particularly during transient operations such as shutdown and startup. A thermal medium flows past and heats or cools one part of the turbine which may have a deleterious thermal mismatch with another part. By passively controlling the flow of cooling medium past the one part in response to relative movement of thermally responsive parts of the turbine, the flow of thermal medium along the flow path can be regulated to increase or reduce the flow, thereby to regulate the temperature of the one part to maintain the thermal mismatch within predetermined limits.

  20. Advances in measuring techniques for turbine cooling test rigs

    NASA Technical Reports Server (NTRS)

    Pollack, F. G.

    1972-01-01

    Surface temperature distribution measurements for turbine vanes and blades were obtained by measuring the infrared energy emitted by the airfoil. The IR distribution can be related to temperature distribution by suitable calibration methods and the data presented in the form of isotherm maps. Both IR photographic and real time electro-optical methods are being investigated. The methods can be adapted to rotating as well as stationary targets, and both methods can utilize computer processing. Pressure measurements on rotating components are made with a rotating system incorporating 10 miniature transducers. A mercury wetted slip ring assembly was used to supply excitation power and as a signal transfer device. The system was successfully tested up to speeds of 9000 rpm and is now being adapted to measure rotating blade airflow quantities in a spin rig and a research engine.

  1. Requirements for high-temperature air-cooled central receivers

    SciTech Connect

    Wright, J.D.; Copeland, R.J.

    1983-12-01

    The design of solar thermal central receivers will be shaped by the end user's need for energy. This paper identifies the requirements for receivers supplying heat for industrial processes or electric power generation in the temperature range 540 to 1000/sup 0/C and evaluates the effects of the requirements on air-cooled central receivers. Potential IPH applications are identified as large baseload users that are located some distance from the receiver. In the electric power application, the receiver must supply heat to a pressurized gas power cycle. The difficulty in providing cost-effective thermal transport and thermal storage for air-cooled receivers is a critical problem.

  2. Small-scale AFBC hot air gas turbine power cycle

    SciTech Connect

    Ashworth, R.A.; Keener, H.M.; Hall, A.W.

    1995-12-31

    The Energy and Environmental Research Corporation (EER), the Ohio Agricultural Research and Development Center (OARDC), the Will-Burt Company (W-B) and the US Department of Energy (DOE) have successfully developed and completed pilot plant tests on a small scale atmospheric fluidized bed combustion (AFBC) system. This system can be used to generate electricity, and/or hot water, steam. Following successful pilot plant operation, commercial demonstration will take place at Cedar Lane Farms (CLF), near Wooster, Ohio. The system demonstration will be completed by the end of 1995. The project is being funded through a cooperative effort between the DOE, EER, W-B, OARDC, CLF and the Ohio Coal Development Office (OCDO). The small scale AFBC, has no internal heat transfer surfaces in the fluid bed proper. Combining the combustor with a hot air gas turbine (HAGT) for electrical power generation, can give a relatively high overall system thermal efficiency. Using a novel method of recovering waste heat from the gas turbine, a gross heat rate of 13,500 Btu/kWhr ({approximately}25% efficiency) can be achieved for a small 1.5 MW{sub e} plant. A low technology industrial recuperation type gas turbine is used that operates with an inlet blade temperature of 1,450 F and a compression ratio of 3.9:1. The AFBC-HAGT technology can be used to generate power for remote rural communities to replace diesel generators, or can be used for small industrial co-generation applications.

  3. An experimental study of film cooling in a rotating transonic turbine

    NASA Astrophysics Data System (ADS)

    Abhari, Reza S.; Epstein, A. H.

    1992-06-01

    Time-resolved measurements of heat transfer on a fully cooled transonic turbine stage have been taken in a short duration turbine test facility which simulates full engine nondimensional conditions. The time average of this data is compared to uncooled rotor data and cooled linear cascade measurements made on the same profile. The film cooling reduces the time-averaged heat transfer compared to the uncooled rotor on the blade suction surface by as much as 60 percent, but has relative little effect on the pressure surface. The suction surface rotor heat transfer is lower than that measured in the cascade. The results are similar over the central 3/4 of the span implying that the flow here is mainly two-dimensional. The film cooling is shown to be much less effective at high blowing ratios than at low ones. Time-resolved measurements reveal that the cooling, when effective, both reduce the dc level of heat transfer and changed the shape of the unsteady waveform. Unsteady blowing is shown to be a principal driver of film cooling fluctuations, and a linear model is shown to do a good job in predicting the unsteady heat transfer. The unsteadiness results in a 12 percent decrease in heat transfer on the suction surface and a 5 percent increase on the pressure surface.

  4. An experimental study of film cooling in a rotating transonic turbine

    SciTech Connect

    Abhari, R.S. ); Epstein, A.H. )

    1994-01-01

    Time-resolved measurements of heat transfer on a fully cooled transonic turbine stage have been taken in a short duration turbine test facility, which simulates full engine nondimensional conditions. The time average of this data is compared to uncooled rotor data and cooled linear cascade measurements made on the same profile. The film cooling reduces the time-averaged heat transfer compared to the uncooled rotor on the blade suction surface by as much as 60 percent, but has relatively little effect on the pressure surface. The suction surface rotor heat transfer is lower than that measured in the cascade. The results are similar over the central 3/4 of the span, implying that the flow here is mainly two dimensional. The film cooling is shown to be much less effective at high blowing ratios than at low ones. Time-resolved measurements reveal that the cooling, when effective, both reduced the dc level of heat transfer and changed the shape of the unsteady waveform. Unsteady blowing is shown to be a principal driver of film cooling fluctuations, and a linear model is shown to do a good job in predicting the unsteady heat transfer. The unsteadiness results in a 12 percent decrease in heat transfer on the suction surface and a 5 percent increase on the pressure surface.

  5. Performance improvement of a cross-flow hydro turbine by air layer effect

    NASA Astrophysics Data System (ADS)

    Choi, Y. D.; Yoon, H. Y.; Inagaki, M.; Ooike, S.; Kim, Y. J.; Lee, Y. H.

    2010-08-01

    The purpose of this study is not only to investigate the effects of air layer in the turbine chamber on the performance and internal flow of the cross-flow turbine, but also to suggest a newly developed air supply method. Field test is performed in order to measure the output power of the turbine by a new air supply method. CFD analysis on the performance and internal flow of the turbine is conducted by an unsteady state calculation using a two-phase flow model in order to embody the air layer effect on the turbine performance effectively.The result shows that air layer effect on the performance of the turbine is considerable. The air layer located in the turbine runner passage plays the role of preventing a shock loss at the runner axis and suppressing a recirculation flow in the runner. The location of air suction hole on the chamber wall is very important factor for the performance improvement. Moreover, the ratio between air from suction pipe and water from turbine inlet is also significant factor of the turbine performance.

  6. New concept of power generation analysis for chemical gas turbine in thermodynamic process with heat sources or cooling devices

    SciTech Connect

    Yokogawa, M.; Taniguchi, H.; Yang, W.J.; Nakahara, T.; Arai, N.

    1998-07-01

    Many years ago, a thermodynamic process of power generation was developed for gas turbines, supported by adiabatic expansion and compression. Recently, the possible inlet temperature of gas turbines has increased to 1,500 C, so its blades on the high temperature side have to be cooled by some fluid. If the authors use the actual adiabatic expansion, it is necessary to check the fluid-dynamic friction caused by fluid flow between the blades. In this case, the gas turbine blades have a cooling effect and frictional heat-generating effect, as well. If the authors introduce a new concept, the chemical gas turbine, which creates a reheating effect by heat sources in the expansion process, the outlet temperature of the gas turbine will be increased by this continuous reheating effect. Therefore, when they estimate the performance of a chemical gas turbine, these cooling, frictional and reheating effects have to be checked by theoretical and experimental procedures. The authors here analyze the thermodynamic process with heat sources or cooling devices to illustrate their theoretical approach to estimating these effects. In this study, the definition of the heat-exchange rate is introduced to analyze each heating or cooling process. If the authors introduce this heat-exchange rate into their analysis of the thermodynamic process, it is possible to differentiate between adiabatic, cooling and heating processes in gas turbines and other machines.

  7. [Development of new type plastics air turbine handpiece for dental use].

    PubMed

    Kusano, M

    1989-06-01

    The noise generated by the metal air turbine handpiece employed in dental practice is considerable and attended with predominant high frequency components. Therefore, investigation of the noise generation mechanism and development of a silent air turbine handpiece was only a matter of course. In addition, the metal air turbine hardpiece is comparatively heavy and its production cost is high. From this point of view as well, production of a light air turbine handpiece at low cost is also desirable. In order to overcome the objections to the metal air turbine handpiece, appropriate plastics materials were employed wherever possible. In this study, the number of revolutions, noise level, frequency analysis, start pressure and weight of newly produced plastics handpieces and metal handpieces were examined and compared. The following results were obtained: 1. The number of revolutions of single-nozzle type air turbine handpieces encased in plastics housings and fitted with metal turbine rotors was higher than that of all-metal air turbine handpieces. The noise level of the former tended to be lower. 2. The number of revolutions of multi-nozzle type air turbine handpieces encased in plastics housings and fitted with turbine rotors with plastics turbine blades was almost equal to that of similar metal handpieces, with the noise level tending to be lower. 3. In the case of handpieces fitted with turbine rotors with dynamic balance, the number of revolutions was high and the noise level was low. This indicated that dynamic balance was a factor affecting the number of revolutions and noise level. 4. Narrow band sound frequency analysis of single-nozzle type air turbine handpieces showed a sharp peak at the fundamental frequency which was the same as the number of revolutions multiplied by the number of rotor turbine blades. It is thought that the noise from air turbine handpieces was aerodynamic in origin, being generated by the periodical interruption of steady air flow by

  8. Optimization of engines for a commercial Mach 0.98 transport using advanced turbine cooling methods

    NASA Technical Reports Server (NTRS)

    Kraft, G. A.; Whitlow, J. B., Jr.

    1972-01-01

    A study was made of an advanced technology airplane using supercritical aerodynamics. Cruise Mach number was 0.98 at 40,000 feet altitude with a payload of 60,000 pounds and a range of 3000 nautical miles. Separate-flow turbofans were examined parametrically to determine the effect of sea-level-static design turbine-inlet-temperature and noise on takeoff gross weight (TOGW) assuming full-film turbine cooling. The optimum turbine inlet temperature was 2650 F. Two-stage-fan engines, with cruise fan pressure ratio of 2.25, achieved a noise goal of 103.5 EPNdB with todays noise technology while one-stage-fan engines, achieved a noise goal of 98 EPNdB. The take-off gross weight penalty to use the one-stage fan was 6.2 percent.

  9. TACT1- TRANSIENT THERMAL ANALYSIS OF A COOLED TURBINE BLADE OR VANE EQUIPPED WITH A COOLANT INSERT

    NASA Technical Reports Server (NTRS)

    Gaugler, R. E.

    1994-01-01

    As turbine-engine core operating conditions become more severe, designers must develop more effective means of cooling blades and vanes. In order to design reliable, cooled turbine blades, advanced transient thermal calculation techniques are required. The TACT1 computer program was developed to perform transient and steady-state heat-transfer and coolant-flow analyses for cooled blades, given the outside hot-gas boundary condition, the coolant inlet conditions, the geometry of the blade shell, and the cooling configuration. TACT1 can analyze turbine blades, or vanes, equipped with a central coolant-plenum insert from which coolant-air impinges on the inner surface of the blade shell. Coolant-side heat-transfer coefficients are calculated with the heat transfer mode at each station being user specified as either impingement with crossflow, forced convection channel flow, or forced convection over pin fins. A limited capability to handle film cooling is also available in the program. The TACT1 program solves for the blade temperature distribution using a transient energy equation for each node. The nodal energy balances are linearized, one-dimensional, heat-conduction equations which are applied at the wall-outer-surface node, at the junction of the cladding and the metal node, and at the wall-inner-surface node. At the mid-metal node a linear, three-dimensional, heat-conduction equation is used. Similarly, the coolant pressure distribution is determined by solving the set of transfer momentum equations for the one-dimensional flow between adjacent fluid nodes. In the coolant channel, energy and momentum equations for one-dimensional compressible flow, including friction and heat transfer, are used for the elemental channel length between two coolant nodes. The TACT1 program first obtains a steady-state solution using iterative calculations to obtain convergence of stable temperatures, pressures, coolant-flow split, and overall coolant mass balance. Transient

  10. Apparatus for impingement cooling a side wall adjacent an undercut region of a turbine nozzle segment

    DOEpatents

    Burdgick, Steven Sebastian

    2002-01-01

    A gas turbine nozzle segment has outer and inner bands and vanes therebetween. Each band includes a side wall, a cover and an impingement plate between the cover and nozzle wall defining two cavities on opposite sides of the impingement plate. Cooling steam is supplied to one cavity for flow through apertures of the impingement plate to cool the nozzle wall. The side wall of the band and inturned flange define with the nozzle wall an undercut region. Slots are formed through the inturned flange along the nozzle side wall. A plate having through-apertures extending between opposite edges thereof is disposed in each slot, the slots and plates being angled such that the cooling medium exiting the apertures in the second cavity lie close to the side wall for focusing and targeting cooling medium onto the side wall.

  11. Techniques for obtaining detailed heat transfer coefficient measurements within gas turbine blade and vane cooling passages

    NASA Astrophysics Data System (ADS)

    Clifford, R. J.; Jones, T. V.; Dunnne, S. T.

    1983-03-01

    Techniques developed jointly by Rolls-Royce Bristol and Oxford University for determining detailed heat transfer distributions inside turbine blade and vane cooling passages are reviewed. Use is made of a low temperature phase change paint to map the heat flux distributions within models of the cooling passages; the paints change from an opaque coating to a clear liquid at a well-defined melting point. In this way the surface temperature history of a model subjected to transient convective heating is recorded. The heat transfer coefficient distribution is deduced from this history using a transient conduction analysis within the model. Results are presented on detailed heat transfer coefficient distributions within a variety of cooling passages; and data obtained from a comprehensive study of a typical engine multipass cooling geometry are examined.

  12. Numerical and Experimental Study of a Cooling for Vanes in a Small Turbine Engine

    NASA Astrophysics Data System (ADS)

    Šimák, Jan; Michálek, Jan

    2016-03-01

    This paper is concerned with a cooling system for inlet guide vanes of a small turbine engine which are exposed to a high temperature gas leaving a combustion chamber. Because of small dimensions of the vanes, only a simple internal cavity and cooling holes can be realized. The idea was to utilize a film cooling technique. The proposed solution was simulated by means of a numerical method based on a coupling of CFD and heat transfer solvers. The numerical results of various scenarios (different coolant temperature, heat transfer to surroundings) showed a desired decrease of the temperature, especially on the most critical part - the trailing edge. The numerical data are compared to results obtained by experimental measurements performed in a test facility in our institute. A quarter segment model of the inlet guide vanes wheel was equipped with thermocouples in order to verify an effect of cooling. Despite some uncertainty in the results, a verifiable decrease of the vane temperature was observed.

  13. An Experimental Study of the Effect of Wake Passing on Turbine Blade Film Cooling

    NASA Technical Reports Server (NTRS)

    Heidmann, James D.; Lucci, Barbara L.; Reshotko, Eli

    1997-01-01

    The effect of wake passing on the showerhead film cooling performance of a turbine blade has been investigated experimentally. The experiments were performed in an annular turbine cascade with an upstream rotating row of cylindrical rods. Nickel thin-film gauges were used to determine local film effectiveness and Nusselt number values for various injectants, blowing ratios, and Strouhal numbers. Results indicated a reduction in film effectiveness with increasing Strouhal number, as well as the expected increase in film effectiveness with blowing ratio. An equation was developed to correlate the span-average film effectiveness data. The primary effect of wake unsteadiness was found to be correlated by a streamwise-constant decrement of 0.094.St. Steady computations were found to be in excellent agreement with experimental Nusselt numbers, but to overpredict experimental film effectiveness values. This is likely due to the inability to match actual hole exit velocity profiles and the absence of a credible turbulence model for film cooling.

  14. 14 CFR 29.1109 - Carburetor air cooling.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Carburetor air cooling. 29.1109 Section 29.1109 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT Powerplant Induction System § 29.1109 Carburetor...

  15. Verifying heat transfer analysis of high pressure cooled turbine blades and disk.

    PubMed

    Yamawaki, S

    2001-05-01

    To demonstrate cooling and heat transfer technology, a core engine test was conducted with a turbine inlet temperature 1700 degrees C. Measurement data were compared with predictions for a vane, a blade, and a disk. Measured cooling effectiveness of the blade and the vane agreed well with predictions. CFD analysis was carried out for verification of the heat transfer coefficient which was adopted from a heat conduction analysis over the disk. The CFD model including bolt heads showed better results than an axisymmetric model. PMID:11460667

  16. Heat transfer measurements with film cooling on a turbine blade profile in cascade

    NASA Astrophysics Data System (ADS)

    Horton, F. G.; Schultz, D. L.; Forest, A. E.

    1985-03-01

    Heat transfer measurements with film cooling have been made on a gas turbine rotor profile in a cascade at engine representative operating conditions. The blade temperature was varied independently to investigate the scaling of heat transfer coefficient, and a superposition model was found to correlate the data. Contrasting results are presented for films on the two surfaces, along with predictions from a two-dimensional boundary layer method.

  17. Measurement of the Coolant Channel Temperatures and Pressures of a Cooled Radial-Inflow Turbine

    NASA Technical Reports Server (NTRS)

    Dicicco, L. Danielle; Nowlin, Brent C.; Tirres, Lizet

    1994-01-01

    Instrumentation has been installed on the surface of a cooled radial-inflow turbine. Thermocouples and miniature integrated sensor pressure transducers were installed to measure steady state coolant temperatures, blade wall temperatures, and coolant pressures. These measurements will eventually be used to determine the heat transfer characteristics of the rotor. This paper will describe the procedures used to install and calibrate the instrumentation and the testing methods followed. A limited amount of data will compare the measured values to the predicted values.

  18. Turbine systems and methods for using internal leakage flow for cooling

    DOEpatents

    Hernandez, Nestor; Gazzillo, Clement; Boss, Michael J.; Parry, William; Tyler, Karen J.

    2010-02-09

    A cooling system for a turbine with a first section and a second section. The first section may include a first line for diverting a first flow with a first temperature from the first section, a second line for diverting a second flow with a second temperature less than the first temperature from the first section, and a merged line for directing a merged flow of the first flow and the second flow to the second section.

  19. A comparison of the analytical and experimental performance of the solid version of a cooled radial turbine

    NASA Technical Reports Server (NTRS)

    Tirres, Lizet

    1991-01-01

    An evaluation of the aerodynamic performance of the solid version of an Allison-designed cooled radial turbine was conducted at NASA Lewis' Warm Turbine Test Facility. The resulting pressure and temperature measurements are used to calculate vane, rotor, and overall stage performance. These performance results are then compared to the analytical results obtained by using NASA's MTSB (MERIDL-TSONIC-BLAYER) code.

  20. Analytical determination of local surface heat-transfer coefficients for cooled turbine blades from measured metal temperatures

    NASA Technical Reports Server (NTRS)

    Brown, W Byron; Esgar, Jack B

    1950-01-01

    Analytical methods are presented for the determination of local values of outside and inside heat-transfer coefficients and effective gas temperatures by use of turbine-blade-temperature measurements. The methods are derived for a number of configurations that can be applied to typical cooled-turbine-blade shapes as well as to other types of heat-transfer apparatus.

  1. Local heat transfer in internally cooled turbine airfoil leading edge regions. I - Impingement cooling without film coolant extraction. II - Impingement cooling with film coolant extraction

    NASA Astrophysics Data System (ADS)

    Bunker, R. S.; Metzger, D. E.

    The highly localized internal heat transfer characteristics of large-scale models of impingement-cooled turbine blade leading edge regions presently studied derives its cooling from a single line of equally-spaced multiple jets aimed at the leading-edge apex, and exiting the leading-edge region in the opposite or chordwise direction. Detailed two-dimensional local surface Nusselt number distributions have been obtained with temperature-indicating coatings. Results indicate generally increasing heat transfer with the 0.6 power of jet Reynolds number. In the second part of this study, in which the same cooling process is used in conjunction with the extraction of the coolant fluid, the results obtained indicate that heat transfer is primarily dependent on jet Reynolds number, with smaller influences from the flow-extraction rate.

  2. A Numerical Study of the Effect of Wake Passing on Turbine Blade Film Cooling

    NASA Technical Reports Server (NTRS)

    Heidmann, James D.

    1995-01-01

    Time-accurate and steady three-dimensional viscous turbulent numerical simulations were performed to study the effect of upstream blade wake passing unsteadiness on the performance of film cooling on a downstream axial turbine blade. The simulations modeled the blade as spanwise periodic and of infinite span. Both aerodynamic and heat transfer quantities were explored. A showerhead film cooling arrangement typical of modern gas turbine engines was employed. Showerhead cooling was studied because of its anticipated strong sensitivity to upstream flow fluctuations. The wake was modeled as a region of zero axial velocity on the upstream computational boundary which translated with each iteration. This model is compatible with a planned companion experiment in which the wakes will be produced by a rotating row of cylindrical rods upstream of an annular turbine cascade. It was determined that a steady solution with appropriate upstream swirl and stagnation pressure predicted the span-average film effectiveness quite well. The major difference is a 2 to 3 percent overprediction of span-average film effectiveness by the steady simulation on the pressure surface and in the showerhead region. Local overpredictions of up to 8 percent were observed in the showerhead region. These differences can be explained by the periodic relative lifting of the boundary layer and enhanced mixing in the unsteady simulations.

  3. Compressor discharge bleed air circuit in gas turbine plants and related method

    DOEpatents

    Anand, Ashok Kumar; Berrahou, Philip Fadhel; Jandrisevits, Michael

    2003-04-08

    A gas turbine system that includes a compressor, a turbine component and a load, wherein fuel and compressor discharge bleed air are supplied to a combustor and gaseous products of combustion are introduced into the turbine component and subsequently exhausted to atmosphere. A compressor discharge bleed air circuit removes bleed air from the compressor and supplies one portion of the bleed air to the combustor and another portion of the compressor discharge bleed air to an exhaust stack of the turbine component in a single cycle system, or to a heat recovery steam generator in a combined cycle system. In both systems, the bleed air diverted from the combustor may be expanded in an air expander to reduce pressure upstream of the exhaust stack or heat recovery steam generator.

  4. Compressor discharge bleed air circuit in gas turbine plants and related method

    DOEpatents

    Anand, Ashok Kumar; Berrahou, Philip Fadhel; Jandrisevits, Michael

    2002-01-01

    A gas turbine system that includes a compressor, a turbine component and a load, wherein fuel and compressor discharge bleed air are supplied to a combustor and gaseous products of combustion are introduced into the turbine component and subsequently exhausted to atmosphere. A compressor discharge bleed air circuit removes bleed air from the compressor and supplies one portion of the bleed air to the combustor and another portion of the compressor discharge bleed air to an exhaust stack of the turbine component in a single cycle system, or to a heat recovery steam generator in a combined cycle system. In both systems, the bleed air diverted from the combustor may be expanded in an air expander to reduce pressure upstream of the exhaust stack or heat recovery steam generator.

  5. Design Evaluation Using Finite Element Analysis of Cooled Silicon Nitride Plates for a Turbine Blade Application

    NASA Technical Reports Server (NTRS)

    Abdul-Aziz, Ali; Baaklini, George Y.; Bhatt, Ramakrishna T.

    2001-01-01

    Two- and three-dimensional finite element analyses were performed on uncoated and thermal barrier coated (TBC) silicon nitride plates with and without internal cooling by air. Steady-state heat-transfer analyses were done to optimize the size and the geometry of the cooling channels to reduce thermal stresses, and to evaluate the thermal environment experienced by the plate during burner rig testing. The limited experimental data available were used to model the thermal profile exerted by the flame on the plate. Thermal stress analyses were performed to assess the stress response due to thermal loading. Contours for the temperature and the representative stresses for the plates were generated and presented for different cooling hole sizes and shapes. Analysis indicates that the TBC experienced higher stresses, and the temperature gradient was much reduced when the plate was internally cooled by air. The advantages and disadvantages of several cooling channel layouts were evaluated.

  6. Numerical Simulation and Experimental Study of a Dental Handpiece Air Turbine

    NASA Astrophysics Data System (ADS)

    Hsu, Chih-Neng; Chiang, Hsiao-Wei D.; Chang, Ya-Yi

    2011-06-01

    Dental air turbine handpieces have been widely used in clinical dentistry for over 30 years, however, little work has been reported on their performance. In dental air turbine handpieces, the types of flow channel and turbine blade shape can have very different designs. These different designs can have major influence on the torque, rotating speed, and power performance. This research is focused on the turbine blade and the flow channel designs. Using numerical simulation and experiments, the key design parameters which influence the performance of dental hand pieces can be studied. Three types of dental air turbine designs with different turbine blades, nozzle angles, nozzle flow channels, and shroud clearances were tested and analyzed. Very good agreement was demonstrated between the numerical simulation analyses and the experiments. Using the analytical model, parametric studies were performed to identify key design parameters.

  7. Experimental Heat Transfer and Bulk Air Temperature Measurements for a Multipass Internal Cooling Model with Ribs and Bleed

    NASA Technical Reports Server (NTRS)

    Thurman, Douglas; Poinsatte, Philip

    2001-01-01

    An experimental study was made to obtain heat transfer and air temperature data for a simple three-leg serpentine test section that simulates a turbine blade internal cooling passage with trip strips and bleed holes. The objectives were to investigate the interaction of ribs and various bleed conditions on internal cooling and to gain a better understanding of bulk air temperature in an internal passage. Steady-state heat transfer measurements were obtained using a transient technique with thermochromic liquid crystals. Trip strips were attached to one wall of the test section and were located either between or near the bleed holes. The bleed holes, used for film cooling, were metered to simulate the effect of external pressure on the turbine blade. Heat transfer enhancement was found to be greater for ribs near bleed holes compared to ribs between holes, and both configurations were affected slightly by bleed rates upstream. Air temperature measurements were taken at discrete locations along one leg of the model. Average bulk air temperatures were found to remain fairly constant along one leg of the model.

  8. Experimental Heat Transfer and Bulk Air Temperature Measurements for a Multipass Internal Cooling Model with Ribs and Bleed

    NASA Technical Reports Server (NTRS)

    Thurman, Douglas; Poinsatte, Philip

    2000-01-01

    An experimental study was made to obtain heat transfer and air temperature data for a simple 3-leg serpentine test section that simulates a turbine blade internal cooling passage with trip strips and bleed holes. The objectives were to investigate the interaction of ribs and various bleed conditions on internal cooling and to gain a better understanding of bulk air temperature in an internal passage. Steady state heat transfer measurements were obtained using a transient technique with thermochromic liquid crystals. Trip strips were attached to one wall of the test section and were located either between or near the bleed holes. The bleed holes, used for film cooling, were metered to simulate the effect of external pressure on the turbine blade. Heat transfer enhancement was found to be greater for ribs near bleed holes compared to ribs between holes, and both configurations were affected slightly by bleed rates upstream. Air temperature measurements were taken at discreet locations along one leg of the model. Average bulk air temperatures were found to remain fairly constant along one leg of the model.

  9. Effect of Velocity and Temperature Distribution at the Hole Exit on Film Cooling of Turbine Blades

    NASA Technical Reports Server (NTRS)

    Garg, V. K.; Gaugler, R. E.

    1997-01-01

    An existing three-dimensional Navier-Stokes code (Arnone et al, 1991), modified Turbine Branch, to include film cooling considerations (Garg and Gaugler, 1994), has been used to study the effect of coolant velocity and temperature distribution at the hole exit on the heat transfer coefficient on three film-cooled turbine blades, namely, the C3X vane, the VKI rotor, and the ACE rotor. Results are also compared with the experimental data for all the blades. Moreover, Mayle's transition criterion (1991), Forest's model for augmentation of leading edge heat transfer due to free-stream turbulence (1977), and Crawford's model for augmentation of eddy viscosity due to film cooling (Crawford et al, 1980) are used. Use of Mayle's and Forest's models is relevant only for the ACE rotor due to the absence of showerhead cooling on this rotor. It is found that, in some cases, the effect of distribution of coolant velocity and temperature at the hole exit can be as much as 60 percent on the heat transfer coefficient at the blade suction surface, and 50 percent at the pressure surface. Also, different effects are observed on the pressure and suction surface depending upon the blade as well as upon the hole shape, conical or cylindrical.

  10. Energy Conservation in Air Cooled Condenser: A Case Study

    NASA Astrophysics Data System (ADS)

    Mallick, D. S.; Paul, S.

    2014-01-01

    Air cooled condensers were first introduced in the US power industry in the early 1970s, but only during the last few decades has the number of installations greatly increased, largely to mitigate the problem of available water supply. Air may be used as a cooling medium in condensers where, primarily, there is scarcity of water, or where the ambient remains significantly cold for major parts of the year. Air cooled condensers are designed considering the design ambient conditions of summer. During winter months, if the air flow rate over the heat transfer surfaces is kept constant, it leads to improved condenser vacuum, and consequently, improved heat rate. Alternatively, the fans may be run at lower speeds, by using variable frequency drives (VFD), so as to keep the condenser vacuum constant, resulting uniform heat rate. This paper compares the economics between the power saved by the use of VFD in the condenser fans, keeping constant heat rate throughout the year, vis-à-vis, the saving in fuel, effected when the fans are operated at constant speed throughout the year and thus achieving improved heat rate during colder ambient.

  11. Air Cooling for High Temperature Power Electronics (Presentation)

    SciTech Connect

    Waye, S.; Musselman, M.; King, C.

    2014-09-01

    Current emphasis on developing high-temperature power electronics, including wide-bandgap materials such as silicon carbide and gallium nitride, increases the opportunity for a completely air-cooled inverter at higher powers. This removes the liquid cooling system for the inverter, saving weight and volume on the liquid-to-air heat exchanger, coolant lines, pumps, and coolant, replacing them with just a fan and air supply ducting. We investigate the potential for an air-cooled heat exchanger from a component and systems-level approach to meet specific power and power density targets. A proposed baseline air-cooled heat exchanger design that does not meet those targets was optimized using a parametric computational fluid dynamics analysis, examining the effects of heat exchanger geometry and device location, fixing the device heat dissipation and maximum junction temperature. The CFD results were extrapolated to a full inverter, including casing, capacitor, bus bar, gate driver, and control board component weights and volumes. Surrogate ducting was tested to understand the pressure drop and subsequent system parasitic load. Geometries that met targets with acceptable loads on the system were down-selected for experimentation. Nine baseline configuration modules dissipated the target heat dissipation, but fell below specific power and power density targets. Six optimized configuration modules dissipated the target heat load, exceeding the specific power and power density targets. By maintaining the same 175 degrees C maximum junction temperature, an optimized heat exchanger design and higher device heat fluxes allowed a reduction in the number of modules required, increasing specific power and power density while still maintaining the inverter power.

  12. 21 CFR 211.46 - Ventilation, air filtration, air heating and cooling.

    Code of Federal Regulations, 2013 CFR

    2013-04-01

    ... 21 Food and Drugs 4 2013-04-01 2013-04-01 false Ventilation, air filtration, air heating and cooling. 211.46 Section 211.46 Food and Drugs FOOD AND DRUG ADMINISTRATION, DEPARTMENT OF HEALTH AND HUMAN SERVICES (CONTINUED) DRUGS: GENERAL CURRENT GOOD MANUFACTURING PRACTICE FOR FINISHED PHARMACEUTICALS Buildings and Facilities § 211.46...

  13. Apparatus and methods for impingement cooling of a side wall of a turbine nozzle segment

    DOEpatents

    Burdgick, Steven Sebastian

    2002-01-01

    A gas turbine nozzle segment has outer and inner bands and a vane therebetween. Each band includes a nozzle wall, a side wall, a cover and an impingement plate between the cover and the nozzle wall defining two cavities on opposite sides of the impingement plate. Cooling steam is supplied to one cavity for flow through apertures of the impingement plate to cool the nozzle wall. The side wall of the band and inturned flange define with the nozzle wall an undercut region. The impingement plate has a turned flange welded to the inturned flange. A backing plate overlies the turned flange and aligned apertures are formed through the backing plate and turned flange to direct and focus cooling flow onto the side wall of the nozzle segment.

  14. Numerical investigation of heat transfer on film-cooled turbine blades.

    PubMed

    Ginibre, P; Lefebvre, M; Liamis, N

    2001-05-01

    The accurate heat transfer prediction of film-cooled blades is a key issue for the aerothermal turbine design. For this purpose, advanced numerical methods have been developed at Snecma Moteurs. The goal of this paper is the assessment of a three-dimensional Navier-Stokes solver, based on the ONERA CANARI-COMET code, devoted to the steady aerothermal computations of film-cooled blades. The code uses a multidomain approach to discretize the blade to blade channel with overlapping structured meshes for the injection holes. The turbulence closure is done by means of either Michel mixing length model or Spalart-Allmaras one transport equation model. Computations of thin 3D slices of three film-cooled nozzle guide vane blades with multiple injections are performed. Aerothermal predictions are compared to experiments carried out by the von Karman Institute. The behavior of the turbulence models is discussed, and velocity and temperature injection profiles are investigated. PMID:11460651

  15. Flow structure and heat exchange analysis in internal cooling channel of gas turbine blade

    NASA Astrophysics Data System (ADS)

    Szwaba, Ryszard; Kaczynski, Piotr; Doerffer, Piotr; Telega, Janusz

    2016-08-01

    This paper presents the study of the flow structure and heat transfer, and also their correlations on the four walls of a radial cooling passage model of a gas turbine blade. The investigations focus on heat transfer and aerodynamic measurements in the channel, which is an accurate representation of the configuration used in aeroengines. Correlations for the heat transfer coefficient and the pressure drop used in the design of radial cooling passages are often developed from simplified models. It is important to note that real engine passages do not have perfect rectangular cross sections, but include corner fillet, ribs with fillet radii and special orientation. Therefore, this work provides detailed fluid flow and heat transfer data for a model of radial cooling geometry which possesses very realistic features.

  16. Two-dimensional cold-air cascade study of a film-cooled turbine stator blade. 4: Comparison of experimental and analytical aerodynamic results for blade with 12 rows of 0.076-centimeter-(0.030-inch-) diameter holes having streamwise ejection angles

    NASA Technical Reports Server (NTRS)

    Prust, H. W., Jr.

    1978-01-01

    Previously published experimental aerodynamic efficiency results for a film cooled turbine stator blade are compared with analytical results computed from two published analytical methods. One method was used as published; the other was modified for certain cases of coolant discharge from the blade suction surface. For coolant ejection from blade surface regions where the surface static pressures are higher than the blade exit pressure, both methods predict the experimental results quite well. However, for ejection from regions with surface static pressures lower than the blade exit pressure, both methods predict too small a change in efficiency. The modified method gives the better prediction.

  17. Experimental and numerical study of open-air active cooling

    NASA Astrophysics Data System (ADS)

    Al-Fifi, Salman Amsari

    The topic of my thesis is Experimental and Numerical Study of Open Air Active Cooling. The present research is intended to investigate experimentally and Numerically the effectiveness of cooling large open areas like stadiums, shopping malls, national gardens, amusement parks, zoos, transportation facilities and government facilities or even in buildings outdoor gardens and patios. Our cooling systems are simple cooling fans with different diameters and a mist system. This type of cooling systems has been chosen among the others to guarantee less energy consumption, which will make it the most favorable and applicable for cooling such places mentioned above. In the experiments, the main focus is to study the temperature domain as a function of different fan diameters aerodynamically similar in different heights till we come up with an empirical relationship that can determine the temperature domain for different fan diameters and for different heights of these fans. The experimental part has two stages. The first stage is devoted to investigate the maximum range of airspeed and profile for three different fan diameters and for different heights without mist, while the second stage is devoted to investigate the maximum range of temperature and profile for the three different diameter fans and for different heights with mist. The computational study is devoted to built an experimentally verified mathematical model to be used in the design and optimization of water mist cooling systems, and to compare the mathematical results to the experimental results and to get an insight of how to apply such evaporative mist cooling for different places for different conditions. In this study, numerical solution is presented based on experimental conditions, such dry bulb temperature, wet bulb temperature, relative humidity, operating pressure and fan airspeed. In the computational study, all experimental conditions are kept the same for the three fans except the fan airspeed

  18. Evaluation of Transpiration-Cooled Turbine Blades with Shells of "Poroloy" Wire Cloth

    NASA Technical Reports Server (NTRS)

    Richards, Hadley T.

    1959-01-01

    An experimental investigation was made to evaluate the durability and permeability of a group of transpiration-cooled, strut-supported turbine blades. The porous shells were formed from a woven-wire material. The blades were fabricated by a contractor for the Bureau of Aeronautics. The results of permeability tests indicated that the shell material exhibited large random variations in local permeability, which result in excessive coolant flows and very nonuniform cooling. For this reason no heat-transfer evaluations were made because any results would have been inconclusive. Four blades were investigated for structural soundness in a turbo-jet engine operating at a turbine-inlet temperature of approximately 1670 deg F and a turbine tip speed of approximately 1305 feet per second. The maximum temperature of the porous-shell material was approximately 1050 deg F. Inspection of the first two blades after 10 minutes of engine operation revealed that the tips of both of the blades had failed. For the second pair of blades, an improved tip cap was provided by the use of built-up weld extending from strut tip to shell. One of these blades was then operated for 33 hours without failure, and was found to be in good condition at the end of this time. The second blade of this second pair failed within the first 10 minutes of operation because of a poor bond between shell and strut lands.

  19. Emperor penguin body surfaces cool below air temperature.

    PubMed

    McCafferty, D J; Gilbert, C; Thierry, A-M; Currie, J; Le Maho, Y; Ancel, A

    2013-06-23

    Emperor penguins Aptenodytes forsteri are able to survive the harsh Antarctic climate because of specialized anatomical, physiological and behavioural adaptations for minimizing heat loss. Heat transfer theory predicts that metabolic heat loss in this species will mostly depend on radiative and convective cooling. To examine this, thermal imaging of emperor penguins was undertaken at the breeding colony of Pointe Géologie in Terre Adélie (66°40' S 140° 01' E), Antarctica in June 2008. During clear sky conditions, most outer surfaces of the body were colder than surrounding sub-zero air owing to radiative cooling. In these conditions, the feather surface will paradoxically gain heat by convection from surrounding air. However, owing to the low thermal conductivity of plumage any heat transfer to the skin surface will be negligible. Future thermal imaging studies are likely to yield further insights into the adaptations of this species to the Antarctic climate. PMID:23466479

  20. Selection and costing of heat exchangers. Air-cooled type

    NASA Astrophysics Data System (ADS)

    1994-12-01

    ESDU 94043 extends the information in ESDU 92013 which, when an air-cooled exchanger is found appropriate and is costed, provides the results for a datum design 40 ft (12.2 m) long with G-fins and 1 in (25 mm) diameter tube operating at a noise level of 85 dBa. It provides factors derived from an analysis of manufacturer's data to be applied to the cost results from ESDU 92013 to account for variations in those parameters and features. Additional guidance on the configuration and use of air-cooled exchangers is given. The data are incorporated in ESDUpac A9213 which is a Fortran program that implements the selection and costing method of ESDU 92013. It is provided on disc in the software volume compiled to run under DOS with a user-friendly interface that prompts on screen for input data.

  1. Emperor penguin body surfaces cool below air temperature

    PubMed Central

    McCafferty, D. J.; Gilbert, C.; Thierry, A.-M.; Currie, J.; Le Maho, Y.; Ancel, A.

    2013-01-01

    Emperor penguins Aptenodytes forsteri are able to survive the harsh Antarctic climate because of specialized anatomical, physiological and behavioural adaptations for minimizing heat loss. Heat transfer theory predicts that metabolic heat loss in this species will mostly depend on radiative and convective cooling. To examine this, thermal imaging of emperor penguins was undertaken at the breeding colony of Pointe Géologie in Terre Adélie (66°40′ S 140° 01′ E), Antarctica in June 2008. During clear sky conditions, most outer surfaces of the body were colder than surrounding sub-zero air owing to radiative cooling. In these conditions, the feather surface will paradoxically gain heat by convection from surrounding air. However, owing to the low thermal conductivity of plumage any heat transfer to the skin surface will be negligible. Future thermal imaging studies are likely to yield further insights into the adaptations of this species to the Antarctic climate. PMID:23466479

  2. Solar air conditioning with solid absorbents and earth cooling

    NASA Astrophysics Data System (ADS)

    Mayer, E.

    An experimental design is described for an efficient desiccant cooling system using natural cold sink to reduce the moisture content of the ambient air. Used in a warm, humid, tropical climate, the unit is shown to provide up to 0.77 ton of refrigeration under extreme conditions with an average daily coefficient of performance of 0.5. Solar heat is applied to regenerate the silica gel.

  3. The effect of rotation on heat transfer to transpiration-cooled turbine blades

    NASA Astrophysics Data System (ADS)

    Epifanov, V. M.; Kurakin, A. A.; Rusetskii, Iu. A.

    1986-08-01

    The effect of rotation on heat transfer to transpiration-cooled turbine blades was studied theoretically, considering a 'channel'-type blade design with permeable profiled skin. The turbined design included radial internal channels for the supply of a coolant from a manifold to different sections of the blade surface. The algorithm developed for calculating the heat-transfer coefficients was applied to four test regimes. The following subroutines were used in the calculation program (in Fortran-IV): (1) for calculating the loss of pressure of the flow of coolant in the metering orifices, (2) for calculating the parameters of the coolant in the flow core, (3) for calculating the accompanying thermal problem, and (4) for calculating the parameters of the dynamic boundary layers. With the number of blade channels not exceeding 12-15, full calculation of heat transfer at the outer and inner surfaces of the permeable skin takes 10-15 min.

  4. Experimental investigation of film cooling flow induced by shaped holes on a turbine blade.

    PubMed

    Barthet, S; Bario, F

    2001-05-01

    The present study is the second half of a piece of work carried out in collaboration with SNECMA. It investigates shaped hole film cooling, numerically and experimentally. The aim of this paper is the experimental analysis of shaped hole film cooling on a large scale turbine blade (1.4 m chord). The test section is a large scale turbine inlet guide vane cascade. The test airfoil is equipped with a row of nine 50 degrees sloped shaped holes. They are located on the suction side at 20% of the curvilinear length of the blade from the stagnation point. The inlet film cooling hole diameter is 12 mm. The jet flow is heated to 55 degrees C above the crossflow temperature. Velocity and temperature field measurements have been done to obtain mean and fluctuating values. The results are compared to those obtained by Béral on the same experimental apparatus and in the same test conditions, for a row of cylindrical holes. PMID:11460642

  5. Passive air cooling of liquid metal-cooled reactor with double vessel leak accommodation capability

    DOEpatents

    Hunsbedt, Anstein; Boardman, Charles E.

    1995-01-01

    A passive and inherent shutdown heat removal method with a backup air flow path which allows decay heat removal following a postulated double vessel leak event in a liquid metal-cooled nuclear reactor. The improved reactor design incorporates the following features: (1) isolation capability of the reactor cavity environment in the event that simultaneous leaks develop in both the reactor and containment vessels; (2) a reactor silo liner tank which insulates the concrete silo from the leaked sodium, thereby preserving the silo's structural integrity; and (3) a second, independent air cooling flow path via tubes submerged in the leaked sodium which will maintain shutdown heat removal after the normal flow path has been isolated.

  6. Passive air cooling of liquid metal-cooled reactor with double vessel leak accommodation capability

    DOEpatents

    Hunsbedt, A.; Boardman, C.E.

    1995-04-11

    A passive and inherent shutdown heat removal method with a backup air flow path which allows decay heat removal following a postulated double vessel leak event in a liquid metal-cooled nuclear reactor is disclosed. The improved reactor design incorporates the following features: (1) isolation capability of the reactor cavity environment in the event that simultaneous leaks develop in both the reactor and containment vessels; (2) a reactor silo liner tank which insulates the concrete silo from the leaked sodium, thereby preserving the silo`s structural integrity; and (3) a second, independent air cooling flow path via tubes submerged in the leaked sodium which will maintain shutdown heat removal after the normal flow path has been isolated. 5 figures.

  7. Coarse Grid Modeling of Turbine Film Cooling Flows Using Volumetric Source Terms

    NASA Technical Reports Server (NTRS)

    Heidmann, James D.; Hunter, Scott D.

    2001-01-01

    The recent trend in numerical modeling of turbine film cooling flows has been toward higher fidelity grids and more complex geometries. This trend has been enabled by the rapid increase in computing power available to researchers. However, the turbine design community requires fast turnaround time in its design computations, rendering these comprehensive simulations ineffective in the design cycle. The present study describes a methodology for implementing a volumetric source term distribution in a coarse grid calculation that can model the small-scale and three-dimensional effects present in turbine film cooling flows. This model could be implemented in turbine design codes or in multistage turbomachinery codes such as APNASA, where the computational grid size may be larger than the film hole size. Detailed computations of a single row of 35 deg round holes on a flat plate have been obtained for blowing ratios of 0.5, 0.8, and 1.0, and density ratios of 1.0 and 2.0 using a multiblock grid system to resolve the flows on both sides of the plate as well as inside the hole itself. These detailed flow fields were spatially averaged to generate a field of volumetric source terms for each conservative flow variable. Solutions were also obtained using three coarse grids having streamwise and spanwise grid spacings of 3d, 1d, and d/3. These coarse grid solutions used the integrated hole exit mass, momentum, energy, and turbulence quantities from the detailed solutions as volumetric source terms. It is shown that a uniform source term addition over a distance from the wall on the order of the hole diameter is able to predict adiabatic film effectiveness better than a near-wall source term model, while strictly enforcing correct values of integrated boundary layer quantities.

  8. Investigation of vessel exterior air cooling for a HLMC reactor

    SciTech Connect

    Sienicki, J. J.; Spencer, B. W.

    2000-01-13

    The Secure Transportable Autonomous Reactor (STAR) concept under development at Argonne National Laboratory provides a small (300 MWt) reactor module for steam supply that incorporates design features to attain proliferation resistance, heightened passive safety, and improved cost competitiveness through extreme simplification. Examples are the achievement of 100%+ natural circulation heat removal from the low power density/low pressure drop ultra-long lifetime core and utilization of lead-bismuth eutectic (LBE) coolant enabling elimination of main coolant pumps as well as the need for an intermediate heat transport circuit. It is required to provide a passive means of removing decay heat and effecting reactor cooldown in the event that the normal steam generator heat sink, including its normal shutdown heat removal mode, is postulated to be unavailable. In the present approach, denoted as the Reactor Exterior Cooling System (RECS), passive decay heat removal is provided by cooling the outside of the containment/guard vessel with air. RECS is similar to the Reactor Vessel Auxiliary Cooling System (RVACS) incorporated into the PRISM design. However, to enhance the heat removal, RECS incorporates fins on the containment vessel exterior to enhance heat transfer to air as well as removable steel venetian conductors that provide a conduction heat transfer path across the reactor vessel-containment vessel gap to enhance heat transfer between the vessels. The objective of the present work is to investigate the effectiveness of air cooling in removing heat from the vessel and limiting the coolant temperature increase following a sudden complete loss of the steam generator heat sink.

  9. TACT 1: A computer program for the transient thermal analysis of a cooled turbine blade or vane equipped with a coolant insert. 2. Programmers manual

    NASA Technical Reports Server (NTRS)

    Gaugler, R. E.

    1979-01-01

    A computer program to calculate transient and steady state temperatures, pressures, and coolant flows in a cooled axial flow turbine blade or vane with an impingement insert is described. Coolant-side heat transfer coefficients are calculated internally in the program, with the user specifying either impingement or convection heat transfer at each internal flow station. Spent impingement air flows in a chordwise direction and is discharged through the trailing edge and through film cooling holes. The ability of the program to handle film cooling is limited by the internal flow model. Input to the program includes a description of the blade geometry, coolant-supply conditions, outside thermal boundary conditions, and wheel speed. The blade wall can have two layers of different materials, such as a ceramic thermal barrier coating over a metallic substrate. Program output includes the temperature at each node, the coolant pressures and flow rates, and the coolant-side heat transfer coefficients.

  10. An air-cooled pulse tube cryocooler with 50 W cooling capacity at 77 K

    NASA Astrophysics Data System (ADS)

    Hu, Jianying; Wang, Xiaotao; Zhu, Jian; Chen, Shuai; Luo, Ercang; Li, Haibin

    2014-01-01

    A pulse tube cryocooler with 50 W cooling capacity at 77 K is developed to cool superconducting devices mounted on automobiles. The envisioned cryocooler weight is less than 40 kg, and the input electric power is less than 1 kW. To achieve these requirements, the working frequency is increased to 75 Hz, and the dual-opposed pistons use gas bearings to reduce compressor weight and volume. The heat from the main heat exchanger is rejected by forced convective air instead of water. The compressor and the cold finger are carefully matched to improve the efficiency. The details of these will be presented in this paper. After some adjustment, a no load temperature for the pulse tube cryocooler of 40 K was achieved with 1 kW input electric power in surroundings at 298 K. At 77 K, the cooling capacity is 50 W. If the main heat exchanger is cooled by water at 293 K, the cooling capacity increases to 64 W, corresponding to a relative Carnot efficiency of 18%.

  11. Bushing retention system for thermal medium cooling delivery tubes in a gas turbine rotor

    DOEpatents

    Mashey, Thomas Charles

    2002-01-01

    Bushings are provided in counterbores for wheels and spacers for supporting thermal medium cooling tubes extending axially adjacent the rim of the gas turbine rotor. The retention system includes a retaining ring disposed in a groove adjacent an end face of the bushing and which retaining ring projects radially inwardly to prevent axial movement of the bushing in one direction. The retention ring has a plurality of circumferentially spaced tabs along its inner diameter whereby the ring is supported by the lands of the tube maintaining its bushing retention function, notwithstanding operation in high centrifugal fields and rotation of the ring in the groove into other circular orientations.

  12. Component testing of a ground based gas turbine steam cooled rich-burn primary zone combustor for emissions control of nitrogeneous fuels

    NASA Technical Reports Server (NTRS)

    Schultz, D. F.

    1986-01-01

    This effort summarizes the work performed on a steam cooled, rich-burn primary zone, variable geometry combustor designed for combustion of nitrogeneous fuels such as heavy oils or synthetic crude oils. The steam cooling was employed to determine its feasibility and assess its usefulness as part of a ground based gas turbine bottoming cycle. Variable combustor geometry was employed to demonstrate its ability to control primary and secondary zone equivalence ratios and overall pressure drop. Both concepts proved to be highly successful in achieving their desired objectives. The steam cooling reduced peak liner temperatures to less than 800 K. This low temperature offers the potential of both long life and reduced use of strategic materials for liner fabrication. These degrees of variable geometry were successfully employed to control air flow distribution within the combustor. A variable blade angle axial flow air swirler was used to control primary zone air flow, while the secondary and tertiary zone air flows were controlled by rotating bands which regulated air flow to the secondary zone quench holes and the dilutions holes respectively.

  13. Method of controlling the side wall thickness of a turbine nozzle segment for improved cooling

    DOEpatents

    Burdgick, Steven Sebastian

    2002-01-01

    A gas turbine nozzle segment has outer and inner bands and a vane extending therebetween. Each band has a side wall, a cover and an impingement plate between the cover and nozzle wall defining two cavities on opposite sides of the impingement plate. Cooling steam is supplied to one cavity for flow through apertures of the impingement plate to cool the nozzle wall. The side wall of the band has an inturned flange defining with the nozzle wall an undercut region. The outer surface of the side wall is provided with a step prior to welding the cover to the side wall. A thermal barrier coating is applied in the step and, after the cover is welded to the side wall, the side wall is finally machined to a controlled thickness removing all, some or none of the coating.

  14. Computer program for generating input for analysis of impingement-cooled, axial-flow turbine blade

    NASA Astrophysics Data System (ADS)

    Rosenbaum, D.

    1980-01-01

    A computer program, TACTGRID, was developed to generate the geometrical input for the TACTI program, a program that calculates transient and steady state temperatures, pressures, and cooling flows in an impingement cooled turbine blade. Using spline curves, the TACTGRID program constructs the blade internal geometry from the previously designed external blade surface and newly selected wall and channel thicknesses. The TACTGRID program generates the TACTI calculational grid, calculates arc length between grid points required by TACTI as input, and prepares the namelist input data set used by TACTI for the blade geometry. In addition, TACTGRID produces a scaled computer plot of each blade slice, detailing the grid and calculational stations, and thus eliminates the need for intermediate drafting.

  15. Experimental data correlations for the effects of rotation on impingement cooling of turbine blades

    NASA Technical Reports Server (NTRS)

    Kreatsoulas, J. C.

    1987-01-01

    The effects of rotation on impingement cooling have been experimentally studied under simulated gas turbine operating conditions. A large scale model of a blade is spun in vacuum. External heating is simulated by resistve dissipation in the thin wall of the model, which is impingement cooled. The local internal Nu is measured by an IR radiometric technique with high spatial resolution. At low rotational speeds the data agree well with published measurements taken under stationary conditions. At higher speeds, a decrease in the average Nu by up to 20-30 percent is observed. Severe gradients are generated near the hub region. Empirical correlations, which fit the data well and render it more useful for design purposes, are presented. Measurements and correlations suggest that rotational effects are very important and can cause premature blade failure.

  16. Integrated axial and tangential serpentine cooling circuit in a turbine airfoil

    SciTech Connect

    Lee, Ching-Pang; Jiang, Nan; Marra, John J; Rudolph, Ronald J; Dalton, John P

    2015-05-05

    A continuous serpentine cooling circuit forming a progression of radial passages (44, 45, 46, 47A, 48A) between pressure and suction side walls (52, 54) in a MID region of a turbine airfoil (24). The circuit progresses first axially, then tangentially, ending in a last radial passage (48A) adjacent to the suction side (54) and not adjacent to the pressure side (52). The passages of the axial progression (44, 45, 46) may be adjacent to both the pressure and suction side walls of the airfoil. The next to last radial passage (47A) may be adjacent to the pressure side wall and not adjacent to the suction side wall. The last two radial passages (47A, 48A) may be longer along the pressure and suction side walls respectively than they are in a width direction, providing increased direct cooling surface area on the interiors of these hot walls.

  17. Pyrometer mount for a closed-circuit thermal medium cooled gas turbine

    DOEpatents

    Jones, Raymond Joseph; Kirkpatrick, Francis Lawrence; Burns, James Lee; Fulton, John Robert

    2002-01-01

    A steam-cooled second-stage nozzle segment has an outer band and an outer cover defining a plenum therebetween for receiving cooling steam for flow through the nozzles to the inner band and cover therefor and return flow through the nozzles. To measure the temperature of the buckets of the stage forwardly of the nozzle stage, a pyrometer boss is electron beam-welded in an opening through the outer band and TIG-welded to the outer cover plate. By machining a hole through the boss and seating a linearly extending tube in the boss, a line of sight between a pyrometer mounted on the turbine frame and the buckets is provided whereby the temperature of the buckets can be ascertained. The welding of the boss to the outer band and outer cover enables steam flow through the plenum without leakage, while providing a line of sight through the outer cover and outer band to measure bucket temperature.

  18. Report on Preliminary Engineering Study for Installation of an Air Cooled Steam Condenser at Brawley Geothermal Plant, Unit No. 1

    SciTech Connect

    1982-03-01

    The Brawley Geothermal Project comprises a single 10 MW nominal geothermal steam turbine-generator unit which has been constructed and operated by the Southern California Edison Company (SCE). Geothermal steam for the unit is supplied through contract by Union Oil Company which requires the return of all condensate. Irrigation District (IID) purchases the electric power generated and provides irrigation water for cooling tower make-up to the plant for the first-five years of operation, commencing mid-1980. Because of the unavailability of irrigation water from IID in the future, SCE is investigating the application and installation of air cooled heat exchangers in conjunction with the existing wet (evaporative) cooling tower with make-up based on use of 180 gpm (nominal) of the geothermal condensate which may be made available by the steam supplier.

  19. Spot cooling. Part 1: Human responses to cooling with air jets

    SciTech Connect

    Melikov, A.K.; Halkjaer, L.; Arakelian, R.S.; Fanger, P.O.

    1994-12-31

    Eight standing male subjects and a thermal manikin were studied for thermal, physiological, and subjective responses to cooling with an air jet at room temperatures of 28 C, 33 C, and 38 C and a constant relative humidity of 50%. The subjects wore a standard uniform and performed light work. A vertical jet and a horizontal jet were employed The target area of the jet, i.e., the cross section of the jet where it first met the subject, had a diameter of 0.4 m and was located 0.5 m from the outlet. Experiments were performed at average temperatures at the jet target area of 20 C, 24 C, and 28 C. Each experiment lasted 190 minutes and was performed with three average velocities at the target area: 1 and 2 m/s and the preferred velocity selected by the subjects. The impact of the relative humidity of the room air, the jet`s turbulence intensity, and the use of a helmet on the physiological and subjective responses of the eight subjects was also studied The responses of the eight subjects were compared with the responses of a group of 29 subjects. The spot cooling improved the thermal conditions of the occupants. The average general thermal sensation for the eight subjects was linearly correlated to the average mean skin temperature and the average sweat rate. An average mean skin temperature of 33 C and an average sweat rate of 33 g{center_dot}h{sup {minus}1} m{sup {minus}2} were found to correspond to a neutral thermal sensation. The local thermal sensation at the neck and at the arm exposed to the cooling jet was found to be a function of the room air temperature and the local air velocity and temperature of the jet. The turbulence intensity of the cooling jet and the humidity of the room air had no impact on the subjects` physiological and subjective responses. Large individual differences were observed in the evaluation of the environment and in the air velocity preferred by the subjects.

  20. The Fluid Dynamics of Secondary Cooling Air-Mist Jets

    NASA Astrophysics Data System (ADS)

    Hernández C., I.; Acosta G., F. A.; Castillejos E., A. H.; Minchaca M., J. I.

    2008-10-01

    For the conditions of thin-slab continuous casting, air-mist secondary cooling occurs in the transition-boiling regime, possibly as a result of an enhanced intermittent contact of high- momentum water drops with the hot metallic surface. The dynamics of the intermittent contact or wetting/dewetting process should be primarily dependent on the drop size, drop impact-velocity and -angle and water-impact flux, which results from the nozzle design and the interaction of the drops with the conveying and entrained air stream. The aim of this article was to develop a model for predicting the last three parameters based on the design and operating characteristics of air-mist nozzles and on experimentally determined drop-size distributions. To do this, the Eulerian fluid-flow field of the air in three dimensions and steady state and the Lagrangian velocities and trajectories of water drops were computed by solving the turbulent Navier Stokes equation for the air coupled to the motion equation for the water drops. In setting this model, it was particularly important to specify appropriately the air-velocity profile at the nozzle orifice, as well as, the water-flux distribution, and the velocities (magnitude and angle) and exit positions of drops with the different sizes generated, hence special attention was given to these aspects. The computed drop velocities, water-impact flux distributions, and air-mist impact-pressure fields compared well with detailed laboratory measurements carried out at ambient temperature. The results indicate that under practical nozzle-operating conditions, the impinging-droplet Weber numbers are high, over most of the water footprint, suggesting that the droplets should establish an intimate contact with the solid surface. However, the associated high mean-droplet fluxes hint that this contact may be obstructed by drop interference at the surface, which would undermine the heat-extraction effectiveness of the impinging mist. The model also points

  1. A fundamentally new approach to air-cooled heat exchangers.

    SciTech Connect

    Koplow, Jeffrey P.

    2010-01-01

    We describe breakthrough results obtained in a feasibility study of a fundamentally new architecture for air-cooled heat exchangers. A longstanding but largely unrealized opportunity in energy efficiency concerns the performance of air-cooled heat exchangers used in air conditioners, heat pumps, and refrigeration equipment. In the case of residential air conditioners, for example, the typical performance of the air cooled heat exchangers used for condensers and evaporators is at best marginal from the standpoint the of achieving maximum the possible coefficient of performance (COP). If by some means it were possible to reduce the thermal resistance of these heat exchangers to a negligible level, a typical energy savings of order 30% could be immediately realized. It has long been known that a several-fold increase in heat exchanger size, in conjunction with the use of much higher volumetric flow rates, provides a straight-forward path to this goal but is not practical from the standpoint of real world applications. The tension in the market place between the need for energy efficiency and logistical considerations such as equipment size, cost and operating noise has resulted in a compromise that is far from ideal. This is the reason that a typical residential air conditioner exhibits significant sensitivity to reductions in fan speed and/or fouling of the heat exchanger surface. The prevailing wisdom is that little can be done to improve this situation; the 'fan-plus-finned-heat-sink' heat exchanger architecture used throughout the energy sector represents an extremely mature technology for which there is little opportunity for further optimization. But the fact remains that conventional fan-plus-finned-heat-sink technology simply doesn't work that well. Their primary physical limitation to performance (i.e. low thermal resistance) is the boundary layer of motionless air that adheres to and envelops all surfaces of the heat exchanger. Within this boundary layer

  2. Performance Prediction Method of CO2 Cycle for Air Cooling

    NASA Astrophysics Data System (ADS)

    Koyama, Shigeru; Xue, Jun; Kuwahara, Ken

    From the perspective of global environmental protection and energy-saving, the research and development on high-efficiency heat pump and refrigeration systems using environment-friendly refrigerants have become one of the most important issues in the air-conditioning and refrigeration sector. In the present work, a steady-state model of the CO2 transcritical cycle for air cooling, which consists of a rotary compressor, a fin-tube gas cooler,a fin-tube evaporator and an expansion valve, has been developed. The detailed model of fin-tube heat exchanger has been constructed by means of the finite volume method, in which the local heat transfer and flow characteristics are evaluated. It should be noted that the effects of the dew condensation generated on the cooling surface are considered in the evaporator model. As a calculation example, the effects of the indoor air wet-bulb temperature on the cycle performance have been examined with this developed simulator.

  3. Single rotor turbine engine

    DOEpatents

    Platts, David A.

    2002-01-01

    There has been invented a turbine engine with a single rotor which cools the engine, functions as a radial compressor, pushes air through the engine to the ignition point, and acts as an axial turbine for powering the compressor. The invention engine is designed to use a simple scheme of conventional passage shapes to provide both a radial and axial flow pattern through the single rotor, thereby allowing the radial intake air flow to cool the turbine blades and turbine exhaust gases in an axial flow to be used for energy transfer. In an alternative embodiment, an electric generator is incorporated in the engine to specifically adapt the invention for power generation. Magnets are embedded in the exhaust face of the single rotor proximate to a ring of stationary magnetic cores with windings to provide for the generation of electricity. In this alternative embodiment, the turbine is a radial inflow turbine rather than an axial turbine as used in the first embodiment. Radial inflow passages of conventional design are interleaved with radial compressor passages to allow the intake air to cool the turbine blades.

  4. Characterization of an inline row impingement channel for turbine blade cooling applications

    NASA Astrophysics Data System (ADS)

    Ricklick, Mark A.

    Gas turbines have become an intricate part of today's society. Besides powering practically all 200,000+ passenger aircraft in use today, they are also a predominate form of power generation when coupled with a generator. The fact that they are highly efficient, and capable of large power to weight ratios, makes gas turbines an ideal solution for many power requirement issues faced today. Designers have even been able to develop small, 'micro' turbines capable of producing efficient portable power. Part of the turbine's success is the fact that their efficiency levels have continuously risen since their introduction in the early 1800's. Along with improvements in our understanding and designs of the aerodynamic components of the turbine, as well as improvements in the areas of material design and combustion control, advances in component cooling techniques have predominantly contributed to this success. This is the result of a simple thermodynamic concept; as the turbine inlet temperature is increased, the overall efficiency of the machine increases as well. Designers have exploited this fact to the extent that modern gas turbines produce rotor inlet temperatures beyond the melting point of the sophisticated materials used within them. This has only been possible through the use of sophisticated cooling techniques, particularly in the 1st stage vanes and blades. Some of the cooling techniques employed today have been internal cooling channels enhanced with various features, film and showerhead cooling, as well as internal impingement cooling scenarios. Impingement cooling has proven to be one of the most capable heat removal processes, and the combination of this cooling feature with that of channel flow, as is done in impingement channel cooling, creates a scenario that has understandably received a great deal of attention in recent years. This study has investigated several of the unpublished characteristics of these impingement channels, including the channel

  5. A Three-Dimensional Coupled Internal/External Simulation of a Film-Cooled Turbine Vane

    NASA Technical Reports Server (NTRS)

    Heidmann, James D.; Rigby, David L.; Ameri, Ali A.

    1999-01-01

    A three-dimensional Navier-Stokes simulation has been performed for a realistic film-cooled turbine vane using the LeRC-HT code. The simulation includes the flow regions inside the coolant plena and film cooling holes in addition to the external flow. The vane is the subject of an upcoming NASA Glenn Research Center experiment and has both circular cross-section and shaped film cooling holes. This complex geometry is modeled using a multi-block grid which accurately discretizes the actual vane geometry including shaped holes. The simulation matches operating conditions for the planned experiment and assumes periodicity in the spanwise direction on the scale of one pitch of the film cooling hole pattern. Two computations were performed for different isothermal wall temperatures, allowing independent determination of heat transfer coefficients and film effectiveness values. The results indicate separate localized regions of high heat transfer coefficient values, while the shaped holes provide a reduction in heat flux through both parameters. Hole exit data indicate rather simple skewed profiles for the round holes, but complex profiles for the shaped holes with mass fluxes skewed strongly toward their leading edges.

  6. Effect of velocity and temperature distribution at the hole exit on film cooling of turbine blades

    NASA Technical Reports Server (NTRS)

    Garg, Vijay K.; Gaugler, Raymond E.

    1995-01-01

    An existing three-dimensional Navier-Stokes code, modified to include film cooling considerations, has been used to study the effect of coolant velocity and temperature distribution at the hole exit on the heat transfer coefficient on three-film-cooled turbine blades, namely, the C3X vane, the VKI rotor, and the ACE rotor. Results are also compared with the experimental data for all the blades. Moreover, Mayle's transition criterion, Forest's model for augmentation of leading edge heat transfer due to freestream turbulence, and Crawford's model for augmentation of eddy viscosity due to film cooling are used. Use of Mayle's and Forest's models is relevant only for the ACE rotor due to the absence of showerhead cooling on this rotor. It is found that, in some cases, the effect of distribution of coolant velocity and temperature at the hole exit can be as much as 60% on the heat transfer coefficient at the blade suction surface, and 50% at the pressure surface. Also, different effects are observed on the pressure and suction surface depending upon the blade as well as upon the hole shape, conical or cylindrical.

  7. Methods for disassembling, replacing and assembling parts of a steam cooling system for a gas turbine

    DOEpatents

    Wilson, Ian D.; Wesorick, Ronald R.

    2002-01-01

    The steam cooling circuit for a gas turbine includes a bore tube assembly supplying steam to circumferentially spaced radial tubes coupled to supply elbows for transitioning the radial steam flow in an axial direction along steam supply tubes adjacent the rim of the rotor. The supply tubes supply steam to circumferentially spaced manifold segments located on the aft side of the 1-2 spacer for supplying steam to the buckets of the first and second stages. Spent return steam from these buckets flows to a plurality of circumferentially spaced return manifold segments disposed on the forward face of the 1-2 spacer. Crossover tubes couple the steam supply from the steam supply manifold segments through the 1-2 spacer to the buckets of the first stage. Crossover tubes through the 1-2 spacer also return steam from the buckets of the second stage to the return manifold segments. Axially extending return tubes convey spent cooling steam from the return manifold segments to radial tubes via return elbows. The bore tube assembly, radial tubes, elbows, manifold segments and crossover tubes are removable from the turbine rotor and replaceable.

  8. Experimental and Analytical Investigation of the Coolant Flow Characteristics in Cooled Turbine Airfoils

    NASA Technical Reports Server (NTRS)

    Damerow, W. P.; Murtaugh, J. P.; Burggraf, F.

    1972-01-01

    The flow characteristics of turbine airfoil cooling system components were experimentally investigated. Flow models representative of leading edge impingement, impingement with crossflow (midchord cooling), pin fins, feeder supply tube, and a composite model of a complete airfoil flow system were tested. Test conditions were set by varying pressure level to cover the Mach number and Reynolds number range of interest in advanced turbine applications. Selected geometrical variations were studied on each component model to determine these effects. Results of these tests were correlated and compared with data available in the literature. Orifice flow was correlated in terms of discharge coefficients. For the leading edge model this was found to be a weak function of hole Mach number and orifice-to-impinged wall spacing. In the impingement with crossflow tests, the discharge coefficient was found to be constant and thus independent of orifice Mach number, Reynolds number, crossflow rate, and impingement geometry. Crossflow channel pressure drop showed reasonable agreement with a simple one-dimensional momentum balance. Feeder tube orifice discharge coefficients correlated as a function of orifice Mach number and the ratio of the orifice-to-approach velocity heads. Pin fin data was correlated in terms of equivalent friction factor, which was found to be a function of Reynolds number and pin spacing but independent of pin height in the range tested.

  9. High-pressure ceramic air heater for indirectly fired gas turbine applications

    NASA Astrophysics Data System (ADS)

    Lahaye, P. G.; Briggs, G. F.; Vandervort, C. L.; Seger, J. L.

    The Externally-Fired Combined Cycle (EFCC) offers a method for operating high-efficiency gas and steam turbine combined cycles on coal. In the EFCC, an air heater replaces the gas turbine combustor so that the turbine can be indirectly fired. Ceramic materials are required for the heat exchange surfaces to accommodate the operating temperatures of modern gas turbines. The ceramic air heater or heat exchanger is the focus of this program, and the two primary objectives are (1) to demonstrate that a ceramic air heater can be reliably pressurized to a level of 225 psia (1.5 MPa); and (2) to show that the air heater can withstand exposure to the products of coal combustion at elevated temperatures. By replacing the gas turbine combustor with a ceramic air heater, the cycle can use coal or other ash-bearing fuels. Numerous programs have attempted to fuel high efficiency gas turbines directly with coal, often resulting in significant ash deposition upon turbine components and corrosion or erosion of turbine blades. This report will show that a ceramic air heater is significantly less susceptible to ash deposition or corrosion than a gas turbine when protected by rudimentary methods of gas-stream clean-up. A 25 x 10(sup 6) Btu/hr (7 MW) test facility is under construction in Kennebunk, Maine. It is anticipated that this proof of concept program will lead to commercialization of the EFCC by electric utility and industrial organizations. Applications are being pursued for power plants ranging from 10 to 100 megawatts.

  10. Rotational coherent anti-stokes Raman spectroscopy measurements in a rotating cavity with axial throughflow of cooling air: oxygen concentration measurements.

    PubMed

    Black, J D; Long, C A

    1992-07-20

    In a rotating cavity rig, which models cooling air flow in the spaces between disks of a gas turbine compressor, the buildup of oxygen concentration after the cooling gas was changed from nitrogen to air was monitored using rotational coherent anti-Stokes Raman spectroscopy (CARS). From this information an estimate of the fraction of the throughflow entering the rotating cavity was obtained. This demonstrates that rotational CARS can be applied as a nonintrusive concentration-measurement technique in a rotating engineering test rig. PMID:20725415

  11. Apparatus and methods for impingement cooling of an undercut region adjacent a side wall of a turbine nozzle segment

    DOEpatents

    Burdgick, Steven Sebastian; Itzel, Gary Michael

    2001-01-01

    A gas turbine nozzle segment has outer and inner bands. Each band includes a side wall, a cover and an impingement plate between the cover and nozzle wall defining two cavities on opposite sides of the impingement plate. Cooling steam is supplied to one cavity for flow through apertures of the impingement plate to cool the nozzle wall. The side wall of the band and inturned flange define with the nozzle wall an undercut region. The inturned flange has a plurality of apertures for directing cooling steam to cool the side wall between adjacent nozzle segments.

  12. Extension of boundary-layer heat-transfer theory to cooled turbine blades

    NASA Technical Reports Server (NTRS)

    Brown, W Byron; Donoughe, Patrick L

    1950-01-01

    An equation for average heat transfer of a surface was derived when the boundary layer changed from laminar to turbulent. Influences on the heat transfer through a laminar boundary layer of Mach number, temperature ratio (gas temperature divided by wall temperature), and exponents of gas-property temperature relations were shown to be relatively small for air with Mach numbers less than 2 and temperature ratios between 1 and 4. Good agreement was obtained with experimental results from cylinders, an airfoil, and turbine blades.

  13. Experimental study of heat transfer from shaft in cooled radial bearing of GNT-25 gas turbine

    NASA Astrophysics Data System (ADS)

    Rukhlinskiy, V. V.; Usayev, I. D.; Yermolenko, A. V.

    1984-02-01

    The heat transfer from the shaft in a cooled radial bearing design was studied experimentally in a GTN-25 gas turbine. The basic dimensions of the bearing were 315 mm inside diameter and 140 mm width. This split bearing had two oil feed orifices in the plane of separation and its housing was cooled with oil fed through an annular chamber. Heating of the shaft neck and the bearing housing under operating conditions was simulated. The experimental data have been processed according to methods of similarity and dimensional analysis, the results yielding semiempirical relations for the temperature and the thermal flux at the rubbing surface during laminar and transitional flow. Relations have also been obtained from these data for the hot spot temperature and the friction coefficient at the rubbing surface. The former characterizes the cooling system design and performance, the latter characterizes the bearing efficiency and economy. The results confirm that the effect of energy dissipation in the lubricant on the intensity of heat transfer from the shaft depends largely on the size and the shape of the shaft bearing clearance.

  14. Fabrication of gas turbine water-cooled composite nozzle and bucket hardware employing plasma spray process

    DOEpatents

    Schilke, Peter W.; Muth, Myron C.; Schilling, William F.; Rairden, III, John R.

    1983-01-01

    In the method for fabrication of water-cooled composite nozzle and bucket hardware for high temperature gas turbines, a high thermal conductivity copper alloy is applied, employing a high velocity/low pressure (HV/LP) plasma arc spraying process, to an assembly comprising a structural framework of copper alloy or a nickel-based super alloy, or combination of the two, and overlying cooling tubes. The copper alloy is plamsa sprayed to a coating thickness sufficient to completely cover the cooling tubes, and to allow for machining back of the copper alloy to create a smooth surface having a thickness of from 0.010 inch (0.254 mm) to 0.150 inch (3.18 mm) or more. The layer of copper applied by the plasma spraying has no continuous porosity, and advantageously may readily be employed to sustain a pressure differential during hot isostatic pressing (HIP) bonding of the overall structure to enhance bonding by solid state diffusion between the component parts of the structure.

  15. Air heating system

    DOEpatents

    Primeau, John J.

    1983-03-01

    A self-starting, fuel-fired, air heating system including a vapor generator, a turbine, and a condenser connected in a closed circuit such that the vapor output from the vapor generator is conducted to the turbine and then to the condenser where it is condensed for return to the vapor generator. The turbine drives an air blower which passes air over the condenser for cooling the condenser. Also, a condensate pump is driven by the turbine. The disclosure is particularly concerned with the provision of heat exchanger and circuitry for cooling the condensed fluid output from the pump prior to its return to the vapor generator.

  16. Deep mine cooling system

    SciTech Connect

    Conan, J.

    1984-11-06

    A deep mine cooling system comprising a compressor supplied with air and rotatively driven by a motor and an expansion turbine supplied with compressed air from said compressor and driving an actuating unit, wherein the compressed air, after leaving the compressor but prior to reaching the expansion turbine, passes through a steam generator whose output provides the energy required to operate an absorption refrigeration machine used to cool utility water for mining, said compressed air on leaving the steam generator going to a first heat exchanger in which it yields calories to a water circuit comprising a second heat exchanger, said second heat exchanger giving off the calories absorbed by the water in the first heat exchanger to the air fed by the second heat exchanger to a drying cell that is regenerated by said air from the second heat exchanger, said drying cell being part of a set of two cells working in alternation, the other cell in the set receiving the compressed air from the first heat exchanger, such that the compressed air is fed to said expansion turbine after leaving said drying unit, and wherein the air exhausted from said expansion turbine is sent to a third heat exchanger after which it is distributed according to the needs of the mine, said third exchanger being traversed by the water collected in the mine, cooled in said exchanger and circulated upon leaving said exchanger to meet the cool water requirements of the mine.

  17. Importance of combining convection with film cooling.

    NASA Technical Reports Server (NTRS)

    Colladay, R. S.

    1972-01-01

    The interaction of film and convection cooling and its effect on wall cooling efficiency is investigated analytically for two cooling schemes for advanced gas turbine applications. The two schemes are full coverage- and counterflow-film cooling. In full coverage film cooling, the cooling air issues from a large number of small discrete holes in the surface. Counterflow film cooling is a film-convection scheme with film injection from a slot geometry. The results indicate that it is beneficial to utilize as much of the cooling air heat sink as possible for convection cooling prior to ejecting it as a film.

  18. Importance of combining convection with film cooling

    NASA Technical Reports Server (NTRS)

    Colladay, R. S.

    1971-01-01

    The interaction of film and convection cooling and its effect on wall cooling efficiency is investigated analytically for two cooling schemes for advanced gas turbine applications. The two schemes are full coverage- and counterflow-film cooling. In full coverage film cooling, the cooling air issues from a large number of small discrete holes in the surface. Counterflow film cooling is a film-convection scheme with film injection from a slot geometry. The results indicate that it is beneficial to utilize as much of the cooling air heat sink as possible for convection cooling prior to ejecting it as a film.

  19. Heat transfer measurements to a gas turbine cooling passage with inclined ribs

    SciTech Connect

    Wang, Z.; Ireland, P.T.; Kohler, S.T.; Chew, J.W.

    1998-01-01

    The local heat transfer coefficient distribution over all four walls of a large-scale model of a gas turbine cooling passage have been measured in great detail. A new method of determining the heat transfer coefficient to the rib surface has been developed and the contribution of the rib, at 5% blockage, to the overall roughened heat transfer coefficient was found to be considerable. The vortex-dominated flow field was interpreted from the detailed form of the measured local heat transfer contours. Computational Fluid Dynamics calculations support this model of the flow and yield friction factors that agree with measured values. Advances in the heat transfer measuring technique and data analysis procedure that confirm the accuracy of the transient method are described in full.

  20. Channel flow modeling of impingement cooling of a rotating turbine blade

    NASA Technical Reports Server (NTRS)

    Koo, J. J.

    1984-01-01

    Local heat transfer distributions in impingement cooling have been measured by Kreatsoulas and Prieser for a range of conditions which model those in actual turbine blades, including the effects of rotation. These data were reported as local Nusselt numbers, but referred to coolant supply conditions. By means of a channel flow modeling of the flow in the supply and impingement passages, the same data are here presented in terms of local Nusselt number distributions such as are used in design. The results in this form are compared to the nonrotating impingement results of Chupp and to the rotating but nonimpingement results of Morris. Rotation reduces the mean Nusselt numbers from these found by Chupp by about 30 percent, and introduces important radial variations which are sensitive to rotation and to leading edge stagger angle.

  1. Emittance and absorptance of NASA ceramic thermal barrier coating system. [for turbine cooling

    NASA Technical Reports Server (NTRS)

    Liebert, C. H.

    1978-01-01

    Spectral emittance measurements were made on a two-layer ceramic thermal barrier coating system consisting of a metal substrate, a NiCrAly bond coating and a yttria-stabilized zirconia ceramic coating. Spectral emittance data were obtained for the coating system at temperatures of 300 to 1590 K, ceramic thickness of zero to 0.076 centimeter, and wavelengths of 0.4 to 14.6 micrometers. The data were transformed into total hemispherical emittance values and correlated with respect to ceramic coating thickness and temperature using multiple regression curve fitting techniques. The results show that the ceramic thermal barrier coating system is highly reflective and significantly reduces radiation heat loads on cooled gas turbine engine components. Calculation of the radiant heat transfer within the nonisothermal, translucent ceramic coating material shows that the gas-side ceramic coating surface temperature can be used in heat transfer analysis of radiation heat loads on the coating system.

  2. Method for forming a liquid cooled airfoil for a gas turbine

    DOEpatents

    Grondahl, Clayton M.; Willmott, Leo C.; Muth, Myron C.

    1981-01-01

    A method for forming a liquid cooled airfoil for a gas turbine is disclosed. A plurality of holes are formed at spaced locations in an oversized airfoil blank. A pre-formed composite liquid coolant tube is bonded into each of the holes. The composite tube includes an inner member formed of an anti-corrosive material and an outer member formed of a material exhibiting a high degree of thermal conductivity. After the coolant tubes have been bonded to the airfoil blank, the airfoil blank is machined to a desired shape, such that a portion of the outer member of each of the composite tubes is contiguous with the outer surface of the machined airfoil blank. Finally, an external skin is bonded to the exposed outer surface of both the machined airfoil blank and the composite tubes.

  3. Optimization of engines for a commercial Mach 0.85 transport using advanced turbine cooling methods

    NASA Technical Reports Server (NTRS)

    Kraft, G. A.

    1972-01-01

    A parametric study was made of a group of separate-flow-turbofan engines for use in advanced technology airplanes designed for a cruise Mach number of 0.85 at 40,000 feet. The three-engined airplanes were sized to carry 200 passengers 3000 nautical miles. Supercritical aerodynamics were assumed. Film-cooled turbines were used and sea-level-static turbine-rotor-inlet temperature was always 2600 F. The optimum cycle depends on the noise goal assumed. Without a noise goal the best fan pressure ratio (FPR) is about 1.90. At noise goals of FAR 36, -10 EPNdb, and -20 EPNdb, the best FPR's are 1.85, 1.76, and 1.70, respectively, at cruise. The take-off FPR's are progressively less than the cruise value as the noise goal approaches -20 EPNdb. The penalties in take-off gross weight incurred were 8.5, 19, and 64 percent at goals of FAR 36, -10 EPNdb, and -20 EPNdb, respectively.

  4. Gas turbine engine active clearance control

    NASA Technical Reports Server (NTRS)

    Deveau, Paul J. (Inventor); Greenberg, Paul B. (Inventor); Paolillo, Roger E. (Inventor)

    1985-01-01

    Method for controlling the clearance between rotating and stationary components of a gas turbine engine are disclosed. Techniques for achieving close correspondence between the radial position of rotor blade tips and the circumscribing outer air seals are disclosed. In one embodiment turbine case temperature modifying air is provided in flow rate, pressure and temperature varied as a function of engine operating condition. The modifying air is scheduled from a modulating and mixing valve supplied with dual source compressor air. One source supplies relatively low pressure, low temperature air and the other source supplies relatively high pressure, high temperature air. After the air has been used for the active clearance control (cooling the high pressure turbine case) it is then used for cooling the structure that supports the outer air seal and other high pressure turbine component parts.

  5. In-situ formation of multiphase air plasma sprayed barrier coatings for turbine components

    DOEpatents

    Subramanian, Ramesh

    2001-01-01

    A turbine component (10), such as a turbine blade, is provided which is made of a metal alloy (22) and a base, planar-grained thermal barrier layer (28) applied by air plasma spraying on the alloy surface, where a heat resistant ceramic oxide overlay material (32') covers the bottom thermal barrier coating (28), and the overlay material is the reaction product of the precursor ceramic oxide overlay material (32) and the base thermal barrier coating material (28).

  6. Gas-fired boiler and turbine air toxics summary report. Final report, January-September 1995

    SciTech Connect

    Rossi-Lane, C.; Stein, D.; Himes, R.

    1996-08-01

    The objective of the report is to provide a summary of the criteria pollutants and hazardous air pollutants (HAPs) emitted from a variety of gas-fired stationary sources including utility boilers, utility turbines, and turbines used for natural gas transmission. The report provides emission factors for each compound measured as a function of load to support general use during the preparation of Title V permit applications.

  7. Design and cold-air test of single-stage uncooled turbine with high work output

    NASA Technical Reports Server (NTRS)

    Moffitt, T. P.; Szanca, E. M.; Whitney, W. J.; Behning, F. P.

    1980-01-01

    A solid version of a 50.8 cm single stage core turbine designed for high temperature was tested in cold air over a range of speed and pressure ratio. Design equivalent specific work was 76.84 J/g at an engine turbine tip speed of 579.1 m/sec. At design speed and pressure ratio, the total efficiency of the turbine was 88.6 percent, which is 0.6 point lower than the design value of 89.2 percent. The corresponding mass flow was 4.0 percent greater than design.

  8. Effect of Film-Hole Shape on Turbine Blade Film Cooling Performance

    NASA Technical Reports Server (NTRS)

    Han, J. C.; Teng, S.

    2000-01-01

    The detailed heat transfer coefficient and film cooling effectiveness distributions as well as tile detailed coolant jet temperature profiles on the suction side of a gas turbine blade A,ere measured using a transient liquid crystal image method and a traversing cold wire and a traversing thermocouple probe, respectively. The blade has only one row of film holes near the gill hole portion on the suction side of the blade. The hole geometries studied include standard cylindrical holes and holes with diffuser shaped exit portion (i.e. fanshaped holes and laidback fanshaped holes). Tests were performed on a five-blade linear cascade in a low-speed wind tunnel. The mainstream Reynolds number based on cascade exit velocity was 5.3 x 10(exp 5). Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. The wake Strouhal number was kept at 0 or 0.1. Coolant blowing ratio was varied from 0.4 to 1.2. Results show that both expanded holes have significantly improved thermal protection over the surface downstream of the ejection location, particularly at high blowing ratios. However, the expanded hole injections induce earlier boundary layer transition to turbulence and enhance heat transfer coefficients at the latter part of the blade suction surface. In general, the unsteady wake tends to reduce film cooling effectiveness.

  9. Measurements of Heat Transfer, Flow, and Pressures in a Simulated Turbine Blade Internal Cooling Passage

    NASA Technical Reports Server (NTRS)

    Russell, Louis M.; Thurman, Douglas R.; Poinsatte, Philip E.; Hippensteele, Steven A.

    1998-01-01

    An experimental study was made to obtain quantitative information on heat transfer, flow, and pressure distribution in a branched duct test section that had several significant features of an internal cooling passage of a turbine blade. The objective of this study was to generate a set of experimental data that could be used for validation of computer codes that would be used to model internal cooling. Surface heat transfer coefficients and entrance flow conditions were measured at nominal entrance Reynolds numbers of 45,000, 335,000, and 726,000. Heat transfer data were obtained by using a steady-state technique in which an Inconel heater sheet is attached to the surface and coated with liquid crystals. Visual and quantitative flow-field data from particle image velocimetry measurements for a plane at midchannel height for a Reynolds number of 45,000 were also obtained. The flow was seeded with polystyrene particles and illuminated by a laser light sheet. Pressure distribution measurements were made both on the surface with discrete holes and in the flow field with a total pressure probe. The flow-field measurements yielded flow-field velocities at selected locations. A relatively new method, pressure sensitive paint, was also used to measure surface pressure distribution. The pressure paint data obtained at Reynolds numbers of 335,000 and 726,000 compared well with the more standard method of measuring pressures by using discrete holes.

  10. Method and apparatus for wind turbine air gap control

    DOEpatents

    Grant, James Jonathan; Bagepalli, Bharat Sampathkumaran; Jansen, Patrick Lee; DiMascio, Paul Stephen; Gadre, Aniruddha Dattatraya; Qu, Ronghai

    2007-02-20

    Methods and apparatus for assembling a wind turbine generator are provided. The wind turbine generator includes a core and a plurality of stator windings circumferentially spaced about a generator longitudinal axis, a rotor rotatable about the generator longitudinal axis wherein the rotor includes a plurality of magnetic elements coupled to a radially outer periphery of the rotor such that an airgap is defined between the stator windings and the magnetic elements and the plurality of magnetic elements including a radially inner periphery having a first diameter. The wind turbine generator also includes a bearing including a first member in rotatable engagement with a radially inner second member, the first member including a radially outer periphery, a diameter of the radially outer periphery of the first member being substantially equal to the first diameter, the rotor coupled to the stator through the bearing such that a substantially uniform airgap is maintained.

  11. Air Force fuel mainburner/turbine effects programs

    NASA Technical Reports Server (NTRS)

    Jackson, T. A.

    1980-01-01

    A program for the determination of fuel property effects on aircraft gas turbine engine mainburners and turbines is discussed. The six engines selected as test candidates are the J79, J85, J57, TF30, TF39, and F100. Fuels election is the responsibility of the contractors with two fuels as exceptions. The petroleum JP-4 is to be used as a baseline in all tests. The shale JP-4 is to be used in nearly all tests. Fuel properties are to be correlated with combustion system performance paramters. In addition, life predictions are to be made for combustor and turbine hardware. These predictions are to be based on a typical mission for each system, measured metal temperatures and temperature gradients, and oxidation/corrosion effects.

  12. Flow Field in a Single-Stage Model Air Turbine With Seal Rings and Pre-Swirled Purge Flow

    NASA Astrophysics Data System (ADS)

    Dunn, Dennis M.

    Modern gas turbines operate at high mainstream gas temperatures and pressures, which requires high durability materials. A method of preventing these hot gases from leaking into the turbine cavities is essential for improved reliability and cost reduction. Utilizing bleed-off air from the compressor to cool internal components has been a common solution, but at the cost of decreasing turbine performance. The present work thoroughly describes the complex flow field between the mainstream gas and a single rotor-stator disk cavity, and mechanisms of mainstream gas ingestion. A combined approach of experimental measurement and numerical simulation are performed on the flow in a single-stage model gas turbine. Mainstream gas ingestion into the cavity is further reduced by utilizing two axially overlapping seal rings, one on the rotor disk and the other on the stator wall. Secondary purge air is injected into the rotor-stator cavity pre-swirled through the stator radially inboard of the two seal rings. Flow field predictions from the simulations are compared against experimental measurements of static pressure, velocity, and tracer gas concentration acquired in a nearly identical model configuration. Operational conditions were performed with a main airflow Reynolds number of 7.86e4 and a rotor disk speed of 3000rpm. Additionally the rotational Reynolds number was 8.74 e5 with a purge air nondimensional flow rate cw=4806. The simulation models a 1/14 rotationally periodic sector of the turbine rig, consisting of four rotor blades and four stator vanes. Gambit was used to generate the three-dimensional unstructured grids ranging from 10 to 20 million cells. Effects of turbulence were modeled using the single-equation Spalart-Allmaras as well as the realizable k-epsilon models. Computations were performed using FLUENT for both a simplified steady-state and subsequent time-dependent formulation. Simulation results show larger scale structures across the entire sector angle

  13. Fuel-air mixing apparatus for reducing gas turbine combustor exhaust emissions

    NASA Technical Reports Server (NTRS)

    Zupanc, Frank J. (Inventor); Yankowich, Paul R. (Inventor)

    2006-01-01

    A fuel-air mixer for use in a combustion chamber of a gas turbine engine is provided. The fuel air mixing apparatus comprises an annular fuel injector having a plurality of discrete plain jet orifices, a first swirler wherein the first swirler is located upstream from the fuel injector and a second swirler wherein the second swirler is located downstream from the fuel injector. The plurality of discrete plain jet orifices are situated between the highly swirling airstreams generated by the two radial swirlers. The distributed injection of the fuel between two highly swirling airstreams results in rapid and effective mixing to the desired fuel-air ratio and prevents the formation of local hot spots in the combustor primary zone. A combustor and a gas turbine engine comprising the fuel-air mixer of the present invention are also provided as well as a method using the fuel-air mixer of the present invention.

  14. Simulations of Turbine Cooling Flows Using a Multiblock-Multigrid Scheme

    NASA Technical Reports Server (NTRS)

    Steinthorsson, Erlendur; Ameri, Ali A.; Rigby, David L.

    1996-01-01

    Results from numerical simulations of air flow and heat transfer in a 'branched duct' geometry are presented. The geometry contains features, including pins and a partition, as are found in coolant passages of turbine blades. The simulations were performed using a multi-block structured grid system and a finite volume discretization of the governing equations (the compressible Navier-Stokes equations). The effects of turbulence on the mean flow and heat transfer were modeled using the Baldwin-Lomax turbulence model. The computed results are compared to experimental data. It was found that the extent of some regions of high heat transfer was somewhat under predicted. It is conjectured that the underlying reason is the local nature of the turbulence model which cannot account for upstream influence on the turbulence field. In general, however, the comparison with the experimental data is favorable.

  15. 16 CFR Appendix H to Part 305 - Cooling Performance and Cost for Central Air Conditioners

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 16 Commercial Practices 1 2011-01-01 2011-01-01 false Cooling Performance and Cost for Central Air Conditioners H Appendix H to Part 305 Commercial Practices FEDERAL TRADE COMMISSION REGULATIONS UNDER SPECIFIC... RULEâ) Pt. 305, App. H Appendix H to Part 305—Cooling Performance and Cost for Central Air...

  16. 16 CFR Appendix H to Part 305 - Cooling Performance and Cost for Central Air Conditioners

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 16 Commercial Practices 1 2010-01-01 2010-01-01 false Cooling Performance and Cost for Central Air Conditioners H Appendix H to Part 305 Commercial Practices FEDERAL TRADE COMMISSION REGULATIONS UNDER SPECIFIC... RULEâ) Pt. 305, App. H Appendix H to Part 305—Cooling Performance and Cost for Central Air...

  17. Flame holding tolerant fuel and air premixer for a gas turbine combustor

    DOEpatents

    York, William David; Johnson, Thomas Edward; Ziminsky, Willy Steve

    2012-11-20

    A fuel nozzle with active cooling is provided. It includes an outer peripheral wall, a nozzle center body concentrically disposed within the outer wall in a fuel and air pre-mixture. The fuel and air pre-mixture includes an air inlet, a fuel inlet and a premixing passage defined between the outer wall in the center body. A gas fuel flow passage is provided. A first cooling passage is included within the center body in a second cooling passage is defined between the center body and the outer wall.

  18. Air/fuel supply system for use in a gas turbine engine

    DOEpatents

    Fox, Timothy A; Schilp, Reinhard; Gambacorta, Domenico

    2014-06-17

    A fuel injector for use in a gas turbine engine combustor assembly. The fuel injector includes a main body and a fuel supply structure. The main body has an inlet end and an outlet end and defines a longitudinal axis extending between the outlet and inlet ends. The main body comprises a plurality of air/fuel passages extending therethrough, each air/fuel passage including an inlet that receives air from a source of air and an outlet. The fuel supply structure communicates with and supplies fuel to the air/fuel passages for providing an air/fuel mixture within each air/fuel passage. The air/fuel mixtures exit the main body through respective air/fuel passage outlets.

  19. Self Adaptive Air Turbine for Wave Energy Conversion Using Shutter Valve and OWC Heoght Control System

    SciTech Connect

    Di Bella, Francis A

    2014-09-29

    An oscillating water column (OWC) is one of the most technically viable options for converting wave energy into useful electric power. The OWC system uses the wave energy to “push or pull” air through a high-speed turbine, as illustrated in Figure 1. The turbine is typically a bi-directional turbine, such as a Wells turbine or an advanced Dennis-Auld turbine, as developed by Oceanlinx Ltd. (Oceanlinx), a major developer of OWC systems and a major collaborator with Concepts NREC (CN) in Phase II of this STTR effort. Prior to awarding the STTR to CN, work was underway by CN and Oceanlinx to produce a mechanical linkage mechanism that can be cost-effectively manufactured, and can articulate turbine blades to improve wave energy capture. The articulation is controlled by monitoring the chamber pressure. Funding has been made available from the U.S. Department of Energy (DOE) to CN (DOE DE-FG-08GO18171) to co-share the development of a blade articulation mechanism for the purpose of increasing energy recovery. However, articulating the blades is only one of the many effective design improvements that can be made to the composite subsystems that constitute the turbine generator system.

  20. Experimental feasibility study of radial injection cooling of three-pad radial air foil bearings

    NASA Astrophysics Data System (ADS)

    Shrestha, Suman K.

    Air foil bearings use ambient air as a lubricant allowing environment-friendly operation. When they are designed, installed, and operated properly, air foil bearings are very cost effective and reliable solution to oil-free turbomachinery. Because air is used as a lubricant, there are no mechanical contacts between the rotor and bearings and when the rotor is lifted off the bearing, near frictionless quiet operation is possible. However, due to the high speed operation, thermal management is one of the very important design factors to consider. Most widely accepted practice of the cooling method is axial cooling, which uses cooling air passing through heat exchange channels formed underneath the bearing pad. Advantage is no hardware modification to implement the axial cooling because elastic foundation structure of foil bearing serves as a heat exchange channels. Disadvantage is axial temperature gradient on the journal shaft and bearing. This work presents the experimental feasibility study of alternative cooling method using radial injection of cooling air directly on the rotor shaft. The injection speeds, number of nozzles, location of nozzles, total air flow rate are important factors determining the effectiveness of the radial injection cooling method. Effectiveness of the radial injection cooling was compared with traditional axial cooling method. A previously constructed test rig was modified to accommodate a new motor with higher torque and radial injection cooling. The radial injection cooling utilizes the direct air injection to the inlet region of air film from three locations at 120° from one another with each location having three axially separated holes. In axial cooling, a certain axial pressure gradient is applied across the bearing to induce axial cooling air through bump foil channels. For the comparison of the two methods, the same amount of cooling air flow rate was used for both axial cooling and radial injection. Cooling air flow rate was

  1. Coaxial fuel and air premixer for a gas turbine combustor

    DOEpatents

    York, William D; Ziminsky, Willy S; Lacy, Benjamin P

    2013-05-21

    An air/fuel premixer comprising a peripheral wall defining a mixing chamber, a nozzle disposed at least partially within the peripheral wall comprising an outer annular wall spaced from the peripheral wall so as to define an outer air passage between the peripheral wall and the outer annular wall, an inner annular wall disposed at least partially within and spaced from the outer annular wall, so as to define an inner air passage, and at least one fuel gas annulus between the outer annular wall and the inner annular wall, the at least one fuel gas annulus defining at least one fuel gas passage, at least one air inlet for introducing air through the inner air passage and the outer air passage to the mixing chamber, and at least one fuel inlet for injecting fuel through the fuel gas passage to the mixing chamber to form an air/fuel mixture.

  2. Cold-air performance of compressor-drive turbine of Department of Energy upgraded automobile gas turbine engine. 2: Stage performance

    NASA Technical Reports Server (NTRS)

    Roelke, R. J.; Haas, J. E.

    1982-01-01

    The aerodynamic performance of the compressor-drive turbine of the DOE upgraded gas turbine engine was determined in low temperature air. The as-received cast rotor blading had a significantly thicker profile than design and a fairly rough surface finish. Because of these blading imperfections a series of stage tests with modified rotors were made. These included the as-cast rotor, a reduced-roughness rotor, and a rotor with blades thinned to near design. Significant performance changes were measured. Tests were also made to determine the effect of Reynolds number on the turbine performance. Comparisons are made between this turbine and the compressor-drive turbine of the DOE baseline gas turbine engine.

  3. Validation of the RVACS (Reactor Vessel Auxiliary Cooling System)/RACS (Reactor Air Cooling System) model in SASSYS-1

    SciTech Connect

    Dunn, F.E.

    1987-01-01

    The SASSYS-1 LMR systems analysis code contains a model for transient analysis of heat removal by a RVACS (Reactor Vessel Auxiliary Cooling System) or a RACS (Reactor Air Cooling System) in an LMR (Liquid Metal Reactor). This model has been validated by comparisons of model predictions with experimental data from a large scale RVACS/RACS simulation experiment performed at Argonne National Laboratory. 4 refs., 1 fig.

  4. Mechanical Design of a Performance Test Rig for the Turbine Air-Flow Task (TAFT)

    NASA Technical Reports Server (NTRS)

    Forbes, John C.; Xenofos, George D.; Farrow, John L.; Tyler, Tom; Williams, Robert; Sargent, Scott; Moharos, Jozsef

    2004-01-01

    To support development of the Boeing-Rocketdyne RS84 rocket engine, a full-flow, reaction turbine geometry was integrated into the NASA-MSFC turbine air-flow test facility. A mechanical design was generated which minimized the amount of new hardware while incorporating all test and instrumentation requirements. This paper provides details of the mechanical design for this Turbine Air-Flow Task (TAFT) test rig. The mechanical design process utilized for this task included the following basic stages: Conceptual Design. Preliminary Design. Detailed Design. Baseline of Design (including Configuration Control and Drawing Revision). Fabrication. Assembly. During the design process, many lessons were learned that should benefit future test rig design projects. Of primary importance are well-defined requirements early in the design process, a thorough detailed design package, and effective communication with both the customer and the fabrication contractors.

  5. Air Corrosivity in U.S. Outdoor-Air-Cooled Data Centers is Similar to That in Conventional Data Centers

    SciTech Connect

    Coles, Henry C.; Han, Taewon; Price, Phillip N.; Gadgil, Ashok J.; Tschudi, William F.

    2011-07-17

    There is a concern that environmental-contamination caused corrosion may negatively affect Information Technology (IT) equipment reliability. Nineteen data centers in the United States and two in India were evaluated using Corrosion Classification Coupons (CCC) to assess environmental air quality as it may relate IT equipment reliability. The data centers were of two basic types: closed and outside-air cooled. A closed data center provides cool air to the IT equipment using air conditioning in which only a small percent age of the recirculation air is make-up air continuously supplied from outside to meet human health requirements. An outside-air cooled data center uses outside air directly as the primary source for IT equipment cooling. Corrosion measuring coupons containing copper and silver metal strips were placed in both closed and outside-air cooled data centers. The coupons were placed at each data center (closed and outside-air cooled types) with the location categorized into three groups: (1) Outside - coupons sheltered, located near or at the supply air inlet, but located before any filtering, (2) Supply - starting just after initial air filtering continuing inside the plenums and ducts feeding the data center rooms, and (3) Inside located inside the data center rooms near the IT equipment. Each coupon was exposed for thirty days and then sent to a laboratory for a corrosion rate measurement analysis. The goal of this research was to investigate whether gaseous contamination is a concern for U.S. data center operators as it relates to the reliability of IT equipment. More specifically, should there be an increased concern if outside air for IT equipment cooling is used To begin to answer this question limited exploratory measurements of corrosion rates in operating data centers in various locations were undertaken. This study sought to answer the following questions: (1) What is the precision of the measurements (2) What are the approximate statistical

  6. Citywide Impacts of Cool Roof and Rooftop Solar Photovoltaic Deployment on Near-Surface Air Temperature and Cooling Energy Demand

    NASA Astrophysics Data System (ADS)

    Salamanca, F.; Georgescu, M.; Mahalov, A.; Moustaoui, M.; Martilli, A.

    2016-04-01

    Assessment of mitigation strategies that combat global warming, urban heat islands (UHIs), and urban energy demand can be crucial for urban planners and energy providers, especially for hot, semi-arid urban environments where summertime cooling demands are excessive. Within this context, summertime regional impacts of cool roof and rooftop solar photovoltaic deployment on near-surface air temperature and cooling energy demand are examined for the two major USA cities of Arizona: Phoenix and Tucson. A detailed physics-based parametrization of solar photovoltaic panels is developed and implemented in a multilayer building energy model that is fully coupled to the Weather Research and Forecasting mesoscale numerical model. We conduct a suite of sensitivity experiments (with different coverage rates of cool roof and rooftop solar photovoltaic deployment) for a 10-day clear-sky extreme heat period over the Phoenix and Tucson metropolitan areas at high spatial resolution (1-km horizontal grid spacing). Results show that deployment of cool roofs and rooftop solar photovoltaic panels reduce near-surface air temperature across the diurnal cycle and decrease daily citywide cooling energy demand. During the day, cool roofs are more effective at cooling than rooftop solar photovoltaic systems, but during the night, solar panels are more efficient at reducing the UHI effect. For the maximum coverage rate deployment, cool roofs reduced daily citywide cooling energy demand by 13-14 %, while rooftop solar photovoltaic panels by 8-11 % (without considering the additional savings derived from their electricity production). The results presented here demonstrate that deployment of both roofing technologies have multiple benefits for the urban environment, while solar photovoltaic panels add additional value because they reduce the dependence on fossil fuel consumption for electricity generation.

  7. Optimum design of bipolar plates for separate air flow cooling system of PEM fuel cells stacks

    NASA Astrophysics Data System (ADS)

    Franco, Alessandro

    2015-12-01

    The paper discusses about thermal management of PEM fuel cells. The objective is to define criteria and guidelines for the design of the air flow cooling system of fuel cells stacks for different combination of power density, bipolar plates material, air flow rate, operating temperature It is shown that the optimization of the geometry of the channel permits interesting margins for maintaining the use of separate air flow cooling systems for high power density PEM fuel cells.

  8. A two-dimensional cascade solution using minimized surface singularity density distributions - with application to film cooled turbine blades

    NASA Technical Reports Server (NTRS)

    Mcfarland, E.; Tabakoff, W.; Hamed, A.

    1977-01-01

    An investigation of the effects of coolant injection on the aerodynamic performance of cooled turbine blades is presented. The coolant injection is modeled in the inviscid irrotational adiabatic flow analysis through the cascade using the distributed singularities approach. The resulting integral equations are solved using a minimized surface singularity density criteria. The aerodynamic performance was evaluated using this solution in conjunction with an existing mixing theory analysis. The results of the present analysis are compared with experimental measurements in cold flow tests.

  9. Numerical simulation and analysis of the internal flow in a Francis turbine with air admission

    NASA Astrophysics Data System (ADS)

    Yu, A.; Luo, X. W.; Ji, B.

    2015-01-01

    In case of hydro turbines operated at part-load condition, vortex ropes usually occur in the draft tube, and consequently generate violent pressure fluctuation. This unsteady flow phenomenon is believed harmful to hydropower stations. This paper mainly treats the internal flow simulation in the draft tube of a Francis turbine. In order to alleviate the pressure fluctuation induced by the vortex rope, air admission from the main shaft center is applied, and the water-air two phase flow in the entire flow passage of a model turbine is simulated based on a homogeneous flow assumption and SST k-ω turbulence model. It is noted that the numerical simulation reasonably predicts the pressure fluctuations in the draft tube, which agrees fairly well with experimental data. The analysis based on the vorticity transport equation shows that the vortex dilation plays a major role in the vortex evolution with air admission in the turbine draft tube, and there is large value of vortex dilation along the vortex rope. The results show that the aeration with suitable air volume fraction can depress the vortical flow, and alleviate the pressure fluctuation in the draft tube.

  10. A Combined Experimental/Computational Study of Flow in Turbine Blade Cooling Passage

    NASA Technical Reports Server (NTRS)

    Tse, D. G. N.; Kreskovsky, J. P.; Shamroth, S. J.; Mcgrath, D. B.

    1994-01-01

    Laser velocimetry was utilized to map the velocity field in a serpentine turbine blade cooling passage at Reynolds and Rotation numbers of up to 25.000 and 0.48. These results were used to assess the combined influence of passage curvature and Coriolis force on the secondary velocity field generated. A Navier-Stokes code (NASTAR) was validated against incompressible test data and then used to simulate the effect of buoyancy. The measurements show a net convection from the low pressure surface to high pressure surface. The interaction of the secondary flows induced by the turns and rotation produces swirl at the turns, which persisted beyond 2 hydraulic diameters downstream of the turns. The incompressible flow field predictions agree well with the measured velocities. With radially outward flow, the buoyancy force causes a further increase in velocity on the high pressure surface and a reduction on the low pressure surface. The results were analyzed in relation to the heat transfer measurements of Wagner et al. (1991). Predicted heat transfer is enhanced on the high pressure surfaces and in turns. The incompressible flow simulation underpredicts heat transfer in these locations. Improvements observed in compressible flow simulation indicate that the buoyancy force may be important.

  11. Heat/mass transfer and flow characteristics of pin fin cooling channels in turbine blades

    NASA Astrophysics Data System (ADS)

    Lau, S. C.; Saxena, A.

    Experiments studied the local heat/mass transfer distributions and pressure drops in pin fin channels that modeled internal cooling passages in gas turbine blades. Heat/mass transfer distributions were determined for a straight flow through a pin fin channel (H/D = 1.0, X/D = S/D = 2.5) and a flow through the pin fin channel with trailing edge flow ejection. The overall friction factor and local pressure drop results were obtained for various configurations and lengths of the trailing edge ejection holes. The results show that, when there is trailing edge flow ejection, the main flow stream turns toward the trailing edge ejection holes. The wake regions downstream of the pins and the regions affected by secondary flow shift toward the ejection holes. The local channel wall heat/mass transfer is generally high immediately upstream of a pin, in the wake region downstream of a pin, and in the regions affected by secondary flow. In the case with trailing edge flow ejection, the heat/mass transfer generally decreases in the radial direction as a result of the reducing radial mass flow rate. The overall friction is higher when the trailing edge ejection holes are longer and when they are configured such that more flow is forced further downstream in the pin fin channel before exiting through the ejection holes.

  12. A combined experiment/computational study of flow in turbine blade cooling passage

    NASA Astrophysics Data System (ADS)

    Tse, D. G. N.; Kreskovsky, J. P.; Shamroth, S. J.; McGrath, D. B.

    1994-05-01

    Laser velocimetry was utilized to map the velocity field in a serpentine turbine blade cooling passage at Reynolds and Rotation numbers of up to 25.000 and 0.48. These results were used to assess the combined influence of passage curvature and Coriolis force on the secondary velocity field generated. A Navier-Stokes code (NASTAR) was validated against incompressible test data and then used to simulate the effect of buoyancy. The measurements show a net convection from the low pressure surface to high pressure surface. The interaction of the secondary flows induced by the turns and rotation produces swirl at the turns, which persisted beyond 2 hydraulic diameters downstream of the turns. The incompressible flow field predictions agree well with the measured velocities. With radially outward flow, the buoyancy force causes a further increase in velocity on the high pressure surface and a reduction on the low pressure surface. The results were analyzed in relation to the heat transfer measurements of Wagner et al. (1991). Predicted heat transfer is enhanced on the high pressure surfaces and in turns. The incompressible flow simulation underpredicts heat transfer in these locations. Improvements observed in compressible flow simulation indicate that the buoyancy force may be important.

  13. Investigation of the flow field downstream of a turbine trailing edge cooled nozzle guide vane

    SciTech Connect

    Sieverding, C.H.; Arts, T.; Denos, R.; Martelli, F.

    1996-04-01

    A trailing edge cooled low aspect ratio transonic turbine guide vane is investigated in the VKI Compression Tube Cascade Facility at an outlet Mach number {bar M}{sub 2,is} = 1.05 and a coolant flow rate {dot m}c/{dot m}g = 3 percent. The outlet flow field is surveyed by combined total-directional pressure probes and temperature probes. Special emphasis is put on the development of low blockage probes. Additional information is provided by oil flow visualizations and numerical flow visualizations with a three-dimensional Navier-Stokes code. The test results describe the strong differences in the axial evolution of the hub and tip endwall and secondary flows and demonstrate the self-similarity of the midspan wake profiles. According to the total pressure and temperature profiles, the wake mixing appears to be very fast in the near-wake but very slow in the far-wake region. The total pressure wake profile appears to be little affected by the coolant flow ejection.

  14. Evaluation of an Integrated Gas-Cooled Reactor Simulator and Brayton Turbine-Generator

    NASA Technical Reports Server (NTRS)

    Hissam, David Andy; Stewart, Eric T.

    2006-01-01

    A closed-loop brayton cycle, powered by a fission reactor, offers an attractive option for generating both planetary and in-space electric power. Non-nuclear testing of this type of system provides the opportunity to safely work out integration and system control challenges for a modest investment. Recognizing this potential, a team at Marshall Space Flight Center has evaluated the viability of integrating and testing an existing gas-cooled reactor simulator and a modified commercially available, off-the-shelf, brayton turbine-generator. Since these two systems were developed independently of one another, this evaluation had to determine if they could operate together at acceptable power levels, temperatures, and pressures. Thermal, fluid, and structural analyses show that this combined system can operate at acceptable power levels and temperatures. In addition, pressure drops across the reactor simulator, although higher than desired, are also viewed as acceptable. Three potential working fluids for the system were evaluated: N2, He/Ar, and He/Xe. Other potential issues, such as electrical breakdown in the generator and the operation of the brayton foil bearings using various gas mixtures, were also investigated.

  15. CACD (Complex Air Cleaning Devices) of the GTE (Gas turbine electrostation)-110: Problems and solutions

    NASA Astrophysics Data System (ADS)

    Budakov, I. V.; Neuimin, V. M.

    2015-12-01

    The paper considers CACD of the compressor of the GTE-110 gas turbine. The CACD efficiency has been tested under different conditions of the GTE-325 of the Ivanovo combined cycle plant (CCP) JSC INTER RAO-Electrogeneration Exploitation. It sets out the requirements for the dust collector, de-icing system, and air intake tract CACD. De-icing and air preparation methods are shown.

  16. Dynamic performance testing of prototype 3 ton air-cooled carrier absorption chiller

    SciTech Connect

    Borst, R.R.; Wood, B.D.

    1985-05-01

    The performance of a prototype 3 ton cooling capacity air-cooled lithium bromide/water absorption chiller was tested using an absorption chiller test facility which was modified to expand its testing capabilities to include air-cooled chillers in addition to water-cooled chillers. Temperatures of the three externally supplied fluid loops: hot water, chilled water, and cooling air, were varied in order to determine the effects this would have on the two principal measures of chiller performance: cooling capacity and thermal coefficient of performance (COP). A number of interrelated factors were identified as contributing to less than expected performance. For comparison, experimental correlations of other investigators for this and other similar absorption chillers are presented. These have been plotted as both contour and three-dimensional performance maps in order to more clearly show the functional dependence of the chiller performance on the fluid loop temperatures.

  17. Dynamic performance testing of prototype 3 ton air-cooled carrier absorption chiller

    NASA Astrophysics Data System (ADS)

    Borst, R. R.; Wood, B. D.

    1985-05-01

    The performance of a prototype three ton cooling capacity air-cooled lithium bromide/water absorption chiller was tested using an absorption chiller test facility which was modified to expand its testing capabilities to include air-cooled chillers in addition to water-cooled chillers. Temperatures of the three externally supplied fluid loops: hot water, chilled water, and cooling air, were varied in order to determine the effects this would have on the two principal measures of chiller performance: cooling capacity and thermal coefficient of performance (COP). A number of interrelated factors were identified as contributing to less than expected performance. For comparison, experimental correlations of other investigators for this and other similar absorption chillers are presented. These have been plotted as both contour and three-dimensional performance maps in order to more clearly show the functional dependence of the chiller performance on the fluid loop temperatures.

  18. Mid-section of a can-annular gas turbine engine with an improved rotation of air flow from the compressor to the turbine

    DOEpatents

    Little, David A.; Schilp, Reinhard; Ross, Christopher W.

    2016-03-22

    A midframe portion (313) of a gas turbine engine (310) is presented and includes a compressor section with a last stage blade to orient an air flow (311) at a first angle (372). The midframe portion (313) further includes a turbine section with a first stage blade to receive the air flow (311) oriented at a second angle (374). The midframe portion (313) further includes a manifold (314) to directly couple the air flow (311) from the compressor section to a combustor head (318) upstream of the turbine section. The combustor head (318) introduces an offset angle in the air flow (311) from the first angle (372) to the second angle (374) to discharge the air flow (311) from the combustor head (318) at the second angle (374). While introducing the offset angle, the combustor head (318) at least maintains or augments the first angle (372).

  19. Intercooler cooling-air weight flow and pressure drop for minimum drag loss

    NASA Technical Reports Server (NTRS)

    Reuter, J George; Valerino, Michael F

    1944-01-01

    An analysis has been made of the drag losses in airplane flight of cross-flow plate and tubular intercoolers to determine the cooling-air weight flow and pressure drop that give a minimum drag loss for any given cooling effectiveness and, thus, a maximum power-plant net gain due to charge-air cooling. The drag losses considered in this analysis are those due to (1) the extra drag imposed on the airplane by the weight of the intercooler, its duct, and its supports and (2) the drag sustained by the cooling air in flowing through the intercooler and its duct. The investigation covers a range of conditions of altitude, airspeed, lift-drag ratio, supercharger-pressure ratio, and supercharger adiabatic efficiency. The optimum values of cooling air pressure drop and weight flow ratio are tabulated. Curves are presented to illustrate the results of the analysis.

  20. Reducing secondary losses by blowing cold air in a turbine

    NASA Technical Reports Server (NTRS)

    Koschel, W.

    1977-01-01

    Local blowing on the profile suction side of the turbine guide wheel blades can be effective in preventing the propagation of secondary flows that is, the transport of casing and hub boundary layers by pressure gradients. Some preliminary results on how the blowing should be accomplished in order to influence the secondary flows in the desired manner are given. The effectiveness of blowing is demonstrated. Blowing is also seen to be more effective than using boundary layer slots as far as diminishing losses in the rim zones is concerned.

  1. Effect of Ambient Design Temperature on Air-Cooled Binary Plant Output

    SciTech Connect

    Dan Wendt; Greg Mines

    2011-10-01

    Air-cooled binary plants are designed to provide a specified level of power production at a particular air temperature. Nominally this air temperature is the annual mean or average air temperature for the plant location. This study investigates the effect that changing the design air temperature has on power generation for an air-cooled binary plant producing power from a resource with a declining production fluid temperature and fluctuating ambient temperatures. This analysis was performed for plants operating both with and without a geothermal fluid outlet temperature limit. Aspen Plus process simulation software was used to develop optimal air-cooled binary plant designs for specific ambient temperatures as well as to rate the performance of the plant designs at off-design operating conditions. Results include calculation of annual and plant lifetime power generation as well as evaluation of plant operating characteristics, such as improved power generation capabilities during summer months when electric power prices are at peak levels.

  2. Multiple piece turbine airfoil

    DOEpatents

    Kimmel, Keith D; Wilson, Jr., Jack W.

    2010-11-02

    A turbine airfoil, such as a rotor blade or a stator vane, for a gas turbine engine, the airfoil formed as a shell and spar construction with a plurality of dog bone struts each mounted within openings formed within the shell and spar to allow for relative motion between the spar and shell in the airfoil chordwise direction while also forming a seal between adjacent cooling channels. The struts provide the seal as well as prevent bulging of the shell from the spar due to the cooling air pressure.

  3. 40 CFR 92.108 - Intake and cooling air measurements.

    Code of Federal Regulations, 2014 CFR

    2014-07-01

    ... practice J244 (incorporated by reference at § 92.5) are allowed. (b) Humidity and temperature measurements. (1) Air that has had its absolute humidity altered is considered humidity-conditioned air. For this type of intake air supply, the humidity measurements must be made within the intake air supply...

  4. 40 CFR 92.108 - Intake and cooling air measurements.

    Code of Federal Regulations, 2012 CFR

    2012-07-01

    ... recommended practice J244 (incorporated by reference at § 92.5) are allowed. (b) Humidity and temperature measurements. (1) Air that has had its absolute humidity altered is considered humidity-conditioned air. For this type of intake air supply, the humidity measurements must be made within the intake air...

  5. 40 CFR 92.108 - Intake and cooling air measurements.

    Code of Federal Regulations, 2011 CFR

    2011-07-01

    ... recommended practice J244 (incorporated by reference at § 92.5) are allowed. (b) Humidity and temperature measurements. (1) Air that has had its absolute humidity altered is considered humidity-conditioned air. For this type of intake air supply, the humidity measurements must be made within the intake air...

  6. Cooling of Poultry Using Immersion or air chilling

    Technology Transfer Automated Retrieval System (TEKTRAN)

    During processing, poultry carcasses must be cooled to 40 F or below within 4 to 8 hours after slaughter to retard growth of pathogenic and spoilage microorganisms. In the U.S., poultry has traditionally been cooled using immersion chilling because this method is both economical and efficient; howe...

  7. TACT1, a computer program for the transient thermal analysis of a cooled turbine blade or vane equipped with a coolant insert. 1. Users manual

    NASA Technical Reports Server (NTRS)

    Gaugler, R. E.

    1978-01-01

    A computer program to calculate transient and steady state temperatures, pressures, and coolant flows in a cooled, axial flow turbine blade or vane with an impingement insert is described. Coolant side heat transfer coefficients are calculated internally in the program, with the user specifying either impingement or convection heat transfer at each internal flow station. Spent impingement air flows in a chordwise direction and is discharged through the trailing edge and through film cooling holes. The ability of the program to handle film cooling is limited by the internal flow model. Sample problems, with tables of input and output, are included in the report. Input to the program includes a description of the blade geometry, coolant supply conditions, outside thermal boundary conditions, and wheel speed. The blade wall can have two layers of different materials, such as a ceramic thermal barrier coating over a metallic substrate. Program output includes the temperature at each node, the coolant pressures and flow rates, and the inside heat-transfer coefficients.

  8. Use of Air2Air Technology to Recover Fresh-Water from the Normal Evaporative Cooling Loss at Coal-Based Thermoelectric Power Plants

    SciTech Connect

    Ken Mortensen

    2009-06-30

    This program was undertaken to build and operate the first Air2Air{trademark} Water Conservation Cooling Tower at a power plant, giving a validated basis and capability for water conservation by this method. Air2Air{trademark} water conservation technology recovers a portion of the traditional cooling tower evaporate. The Condensing Module provides an air-to-air heat exchanger above the wet fill media, extracting the heat from the hot saturated moist air leaving in the cooling tower and condensing water. The rate of evaporate water recovery is typically 10%-25% annually, depending on the cooling tower location (climate).

  9. Experimental investigation of heat transfer and flow using V and broken V ribs within gas turbine blade cooling passage

    NASA Astrophysics Data System (ADS)

    Kumar, Sourabh; Amano, R. S.

    2015-05-01

    Gas turbines are extensively used for aircraft propulsion, land-based power generation, and various industrial applications. With an increase in turbine rotor inlet temperatures, developments in innovative gas turbine cooling technology enhance the efficiency and power output; these advancements of turbine cooling have allowed engine designs to exceed normal material temperature limits. For internal cooling design, techniques for heat extraction from the surfaces exposed to hot stream of gas are based on an increase in the heat transfer areas and on the promotion of turbulence of the cooling flow. In this study, an improvement in performance is obtained by casting repeated continuous V- and broken V-shaped ribs on one side of the two pass square channels into the core of the blade. A detailed experimental investigation is done for two pass square channels with a 180° turn. Detailed heat transfer distribution occurring in the ribbed passage is reported for a steady state experiment. Four different combinations of 60° V- and broken 60° V-ribs in a channel are considered. A series of thermocouples are used to obtain the temperature on the channel surface and local heat transfer coefficients are obtained for Reynolds numbers 16,000, 56,000 and 85,000 within the turbulent flow regime. Area averaged data are calculated in order to compare the overall performance of the tested ribbed surface and to evaluate the degree of heat transfer enhancement induced by the rib. Flow within the channels is characterized by heat transfer enhancing ribs, bends, rotation and buoyancy effects. A series of experimental measurements is performed to predict the overall performance of the channel. This paper presents an attempt to collect information about the Nusselt number, the pressure drop and the overall performance of the eight different ribbed ducts at the specified Reynolds number. The main contribution of this study is to evaluate the best combination of rib arrangements

  10. Aerothermal shape optimization for a double row of discrete film cooling holes on the suction surface of a turbine vane

    NASA Astrophysics Data System (ADS)

    El Ayoubi, Carole; Ghaly, Wahid; Hassan, Ibrahim

    2015-10-01

    A multiple-objective optimization is implemented for a double row of staggered film holes on the suction surface of a turbine vane. The optimization aims to maximize the film cooling performance, which is assessed using the cooling effectiveness, while minimizing the corresponding aerodynamic loss, which is measured with a mass-averaged total pressure coefficient. Three geometric variables defining the hole shape are optimized: the conical expansion angle, compound angle and length to diameter ratio of the non-diffused portion of the hole. The optimization employs a non-dominated sorting genetic algorithm coupled with an artificial neural network to generate the Pareto front. Reynolds-averaged Navier-Stokes simulations are employed to construct the neural network and investigate the aerodynamic and thermal optimum solutions. The optimum designs exhibit improved performance in comparison to the reference design. The optimization methodology allowed investigation into the impact of varying the geometric variables on the cooling effectiveness and the aerodynamic loss.

  11. Turbine vane gas film cooling with injection in the leading edge region from a single row of spanwise angled holes

    NASA Technical Reports Server (NTRS)

    Lecuyer, M. R.; Hanus, G. J.

    1976-01-01

    An experimental study of gas film cooling was conducted on a 3X size model turbine vane. Injection in the leading edge region was from a single row of holes angled in a spanwise direction. Measurements of the local heat flux downstream from the row of coolant holes, both with and without film coolant flow, were used to determine the film cooling performance presented in terms of the Stanton number ratio. Results for a range of coolant blowing ratio, M = 0 to 2.0, indicate a reduction in heat flux of up to 15 to 30 percent at a point 10 to 11 hole diameters downstream from injection. An optimum coolant blowing ratio corresponds to a coolant-to-freestream velocity ratio in the range of 0.5. The shallow injection angle resulted in superior cooling performance for injection closest to stagnation, while the effect of injection angle was insignificant for injection further from stagnation.

  12. Effects of intensive evaporative cooling on performance characteristics of land-based gas turbine

    SciTech Connect

    Utamura, Motoaki; Kuwahara, Takaaki; Murata, Hidetaro; Horii, Nobuyuki

    1999-07-01

    Injection of finely-atomized water droplets at inlet to compressor is demonstrated to increase the power and augment the efficiency of gas turbine using a 115MW simple cycle commercial power plant. Power-up mechanism of the present system is identified to be a composite of three existing methods. Design requirement on droplet diameter is discussed in view of blade erosion as well as evaporation efficiency within the compressor. Special spray nozzle to generate water droplets with sauter mean diameter of 10 {micro} m is developed and applied to demonstration test. Experiments show that injection of spray water of 1% to air mass ratio would increase power output by about 10% and thermal efficiency by 3% (relative) respectively. A newly introduced incremental efficiency defined as the ratio of incremental power to additional fuel energy is found to be in excess of 10% (absolute) over thermal efficiency in case without water injection and to be independent of spray amount. It is also revealed that the operation of water spraying suppresses dust deposition on compressor blades under proper control of water quality, which mitigates the deterioration of compressor adiabatic efficiency.

  13. Complementary velocity and heat transfer measurements in a rotating turbine cooling passage

    NASA Astrophysics Data System (ADS)

    Bons, Jeffrey Peter

    An experimental investigation was conducted on the internal flowfield of a simulated turbine blade cooling passage. The passage is of a square cross-section and was manufactured from quartz for optical accessibility. Velocity measurements were taken using Particle Image Velocimetry for both heated and non-heated cases. Thin film resistive heaters on the four passage walls allow heat to be added to the coolant flow without obstructing laser access. Under the same conditions, an infrared detector with associated optics collected wall temperature data for use in calculating local Nusselt number. The test section was operated with radial outward flow and at values of Reynolds number, Rotation number, and density ratio typical of applications. Velocity data for the non-heated case document the evolution of the Coriolis-induced double vortex. The vortex has the effect of increasing the leading side boundary layer thickness while decreasing the trailing side boundary layer thickness. Also, the streamwise component of the Coriolis acceleration creates a thinned side wall boundary layer. These data reveal an unsteady, turbulent flowfield in the cooling passage. Velocity data for the heated case show a strongly distorted streamwise profile indicative of a buoyancy effect on the leading side. The Coriolis vortex is the mechanism for the accumulation of stagnant flow on the leading side of the passage. Heat transfer data show a maximum factor of two difference in the Nusselt number from trailing side to leading side. An estimate of this heat transfer disparity based on the measured boundary layer edge velocity yields approximately the same factor of two. A momentum integral model was developed for data interpretation which accounts for Coriolis and buoyancy effects. Calculated streamwise profiles and secondary flows match the experimental data well. The model, the velocity data, and the heat transfer data combine to suggest the presence of separated flow on the leading wall

  14. Interim Report: Air-Cooled Condensers for Next Generation Geothermal Power Plants Improved Binary Cycle Performance

    SciTech Connect

    Daniel S. Wendt; Greg L. Mines

    2010-09-01

    As geothermal resources that are more expensive to develop are utilized for power generation, there will be increased incentive to use more efficient power plants. This is expected to be the case with Enhanced Geothermal System (EGS) resources. These resources will likely require wells drilled to depths greater than encountered with hydrothermal resources, and will have the added costs for stimulation to create the subsurface reservoir. It is postulated that plants generating power from these resources will likely utilize the binary cycle technology where heat is rejected sensibly to the ambient. The consumptive use of a portion of the produced geothermal fluid for evaporative heat rejection in the conventional flash-steam conversion cycle is likely to preclude its use with EGS resources. This will be especially true in those areas where there is a high demand for finite supplies of water. Though they have no consumptive use of water, using air-cooling systems for heat rejection has disadvantages. These systems have higher capital costs, reduced power output (heat is rejected at the higher dry-bulb temperature), increased parasitics (fan power), and greater variability in power generation on both a diurnal and annual basis (larger variation in the dry-bulb temperature). This is an interim report for the task ‘Air-Cooled Condensers in Next- Generation Conversion Systems’. The work performed was specifically aimed at a plant that uses commercially available binary cycle technologies with an EGS resource. Concepts were evaluated that have the potential to increase performance, lower cost, or mitigate the adverse effects of off-design operation. The impact on both cost and performance were determined for the concepts considered, and the scenarios identified where a particular concept is best suited. Most, but not all, of the concepts evaluated are associated with the rejection of heat. This report specifically addresses three of the concepts evaluated: the use of

  15. Gas turbine engine adapted for use in combination with an apparatus for separating a portion of oxygen from compressed air

    DOEpatents

    Bland, Robert J.; Horazak, Dennis A.

    2012-03-06

    A gas turbine engine is provided comprising an outer shell, a compressor assembly, at least one combustor assembly, a turbine assembly and duct structure. The outer shell includes a compressor section, a combustor section, an intermediate section and a turbine section. The intermediate section includes at least one first opening and at least one second opening. The compressor assembly is located in the compressor section to define with the compressor section a compressor apparatus to compress air. The at least one combustor assembly is coupled to the combustor section to define with the combustor section a combustor apparatus. The turbine assembly is located in the turbine section to define with the turbine section a turbine apparatus. The duct structure is coupled to the intermediate section to receive at least a portion of the compressed air from the compressor apparatus through the at least one first opening in the intermediate section, pass the compressed air to an apparatus for separating a portion of oxygen from the compressed air to produced vitiated compressed air and return the vitiated compressed air to the intermediate section via the at least one second opening in the intermediate section.

  16. Preliminary investigation of cooling-air ejector performance at pressure ratios from 1 to 10

    NASA Technical Reports Server (NTRS)

    Ellis, C W; Hollister, D P; Sargent, A F , Jr

    1951-01-01

    Preliminary investigation was made of conical cooling air ejector at primary pressure ratios from 1 to 10. The cooling-air flow was maintained at zero and the resulting pressure variation in the shroud indicated pumping ability. The cooling-air flow was maintained at zero and the resulting pressure variation in the shroud indicated pumping ability. The gross thrust of the ejector and nozzle were compared. Several ratios of the spacing between the nozzle and shroud exit to the nozzle exit diameter were investigated for several shroud to nozzle exit diameter ratios. Maximum gross thrust loss occurred under conditions of zero cooling-air flow and was as much as 35 percent below nozzle jet thrust. For minimum thrust loss, ejector should be designed with as low diameter and spacing ratio as possible.

  17. A new opportunity for hydro: Using air turbines for generating electricity

    SciTech Connect

    Gorlov, A.M. )

    1992-09-01

    A concept that uses hydropower to compress air could increase the number of locations where hydro is economically and environmentally feasible. The idea is being tested in a demonstration project in the northeastern U.S. The hydroelectric industry could experience substantial growth in low-head hydro facilities if a concept now being developed proves successful. This concept aims to enable power developers to generate electricity economically at sites currently not feasible for hydropower because water heads are too low. Many areas of North America are studded with low-head dams that could provide considerable hydro capacity if low-head generation were economically feasible. The six New England states in the US, for example, contain approximately 15,000 dams that have never been used to generate electric power because they impound water with heads ranging from 3 to 13 feet. Conventional facilities are not economically practical for generating electricity at these low heads. However, a promising alternative approach is to use water at these low-head dams to compress air, and then to use the air to power an air turbine-generator that produces electricity. The concept, called hydropneumatic generation, can be visualized by imagining a container, such as a large teacup, inverted and submerged in tidal waters. As the tide rises, the water compresses the air trapped inside the container. When the tide ebbs, the pressure decreases, putting the air into a partial vacuum. If a vent pipe were installed from the container to the atmosphere, air would flow out of the container as the water depth increased, and flow back in as the water depth decreased. Hydropneumatic energy is generated by installing an air-powered turbine to harness the energy of this airflow through the vent pipe. The turbine can be installed to rotate in the same direction at all times, even though the airflow reverses direction.

  18. The start-up of a gas turbine engine using compressed air tangentially fed onto the blades of the basic turbine

    NASA Technical Reports Server (NTRS)

    Slobodyanyuk, L. K.; Dayneko, V. I.

    1983-01-01

    The use of compressed air was suggested to increase the reliability and motor lifetime of a gas turbine engine. Experiments were carried out and the results are shown in the form of the variation in circumferential force as a function of the entry angle of the working jet onto the turbine blade. The described start-up method is recommended for use with massive rotors.

  19. 16 CFR Appendix H to Part 305 - Cooling Performance and Cost for Central Air Conditioners

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 16 Commercial Practices 1 2014-01-01 2014-01-01 false Cooling Performance and Cost for Central Air Conditioners H Appendix H to Part 305 Commercial Practices FEDERAL TRADE COMMISSION REGULATIONS UNDER SPECIFIC... CONSERVATION ACT (âENERGY LABELING RULEâ) Pt. 305, App. H Appendix H to Part 305—Cooling Performance and...

  20. Dehumidifying Air for Cooling & Refrigeration: Nanotechnology Membrane-based Dehumidifier

    SciTech Connect

    2010-10-01

    Broad Funding Opportunity Announcement Project: Dais is developing a product called NanoAir which dehumidifies the air entering a building to make air conditioning more energy efficient. The system uses a polymer membrane that allows moisture but not air to pass through it. A vacuum behind the membrane pulls water vapor from the air, and a second set of membranes releases the water vapor outside. The membrane’s high selectivity translates into reduced energy consumption for dehumidification. Dais’ design goals for NanoAir are the use of proprietary materials and processes and industry-standard installation techniques. NanoAir is also complementary to many other energy saving strategies, including energy recovery.

  1. Mechanical Design of a Performance Test Rig for the Turbine Air-Flow Task (TAFT)

    NASA Technical Reports Server (NTRS)

    Xenofos, George; Forbes, John; Farrow, John; Williams, Robert; Tyler, Tom; Sargent, Scott; Moharos, Jozsef

    2003-01-01

    To support development of the Boeing-Rocketdyne RS84 rocket engine, a fill-flow, reaction turbine geometry was integrated into the NASA-MSFC turbine air-flow test facility. A mechanical design was generated which minimized the amount of new hardware while incorporating all test and instrUmentation requirements. This paper provides details of the mechanical design for this Turbine Air-Flow Task (TAFT) test rig. The mechanical design process utilized for this task included the following basic stages: Conceptual Design. Preliminary Design. Detailed Design. Baseline of Design (including Configuration Control and Drawing Revision). Fabrication. Assembly. During the design process, many lessons were learned that should benefit future test rig design projects. Of primary importance are well-defined requirements early in the design process, a thorough detailed design package, and effective communication with both the customer and the fabrication contractors. The test rig provided steady and unsteady pressure data necessary to validate the computational fluid dynamics (CFD) code. The rig also helped characterize the turbine blade loading conditions. Test and CFD analysis results are to be presented in another JANNAF paper.

  2. Method of predicting radiation heat transfer in turbine cooling test facilities

    NASA Technical Reports Server (NTRS)

    Gladden, H. J.; Liebert, C. H.

    1975-01-01

    A method is presented for calculating the average net radiation heat flux to turbine vanes and blades. The net radiation heat flux at a vane leading edge calculated by this method was compared with heat flux values independently determined from experimental tests on a vane in a cascade. The spectral emissivities of the turbine vane and the cascade wall were also measured.

  3. Wing-Nacelle-Propeller Tests - Comparative Tests of Liquid-Cooled and Air-Cooled Engine Nacelles

    NASA Technical Reports Server (NTRS)

    Wood, Donald H.

    1934-01-01

    This report gives the results of measurements of the lift, drag, and propeller characteristics of several wing and nacelle combinations with a tractor propeller. The nacelles were so located that the propeller was about 31% of the wing chord directly ahead of the leading edge of the wing, a position which earlier tests (NASA Report No. 415) had shown to be efficient. The nacelles were scale models of an NACA cowled nacelle for a radial air-cooled engine, a circular nacelle with the V-type engine located inside and the radiator for the cooling liquid located inside and the radiator for the type, and a nacelle shape simulating the housing which would be used for an extension shaft if the engine were located entirely within the wing. The propeller used in all cases was a 4-foot model of Navy No. 4412 adjustable metal propeller. The results of the tests indicate that, at the angles of attack corresponding to high speeds of flight, there is no marked advantage of one type of nacelle over the others as far as low drag is concerned, since the drag added by any of the nacelles in the particular location ahead of the wing is very small. The completely cowled nacelle for a radial air-cooled engine appears to have the highest drag, the liquid-cooled engine appears to have the highest drag, the liquid-cooled engine nacelle with external radiator slightly less drag. The liquid-cooled engine nacelle with radiator in the cowling hood has about half the drag of the cowled radial air-cooled engine nacelle. The extension-shaft housing shows practically no increase in drag over that of the wing alone. A large part of the drag of the liquid-cooled engine nacelle appears to be due to the external radiator. The maximum propulsive efficiency for a given propeller pitch setting is about 2% higher for the liquid-cooled engine nacelle with the radiator in the cowling hood than that for the other cowling arrangements.

  4. Thermo-economic comparative analysis of gas turbine GT10 integrated with air and steam bottoming cycle

    NASA Astrophysics Data System (ADS)

    Czaja, Daniel; Chmielnak, Tadeusz; Lepszy, Sebastian

    2014-12-01

    A thermodynamic and economic analysis of a GT10 gas turbine integrated with the air bottoming cycle is presented. The results are compared to commercially available combined cycle power plants based on the same gas turbine. The systems under analysis have a better chance of competing with steam bottoming cycle configurations in a small range of the power output capacity. The aim of the calculations is to determine the final cost of electricity generated by the gas turbine air bottoming cycle based on a 25 MW GT10 gas turbine with the exhaust gas mass flow rate of about 80 kg/s. The article shows the results of thermodynamic optimization of the selection of the technological structure of gas turbine air bottoming cycle and of a comparative economic analysis. Quantities are determined that have a decisive impact on the considered units profitability and competitiveness compared to the popular technology based on the steam bottoming cycle. The ultimate quantity that can be compared in the calculations is the cost of 1 MWh of electricity. It should be noted that the systems analyzed herein are power plants where electricity is the only generated product. The performed calculations do not take account of any other (potential) revenues from the sale of energy origin certificates. Keywords: Gas turbine air bottoming cycle, Air bottoming cycle, Gas turbine, GT10

  5. Turbine design review text

    NASA Technical Reports Server (NTRS)

    1975-01-01

    Three-volume publication covers theoretical, design, and performance aspects of turbines. Volumes cover thermodynamic and fluid-dynamic concepts, velocity diagram design, turbine blade aerodynamic design, turbine energy losses, supersonic turbines, radial-inflow turbines, turbine cooling, and aerodynamic performance testing.

  6. Turbo test rig with hydroinertia air bearings for a palmtop gas turbine

    NASA Astrophysics Data System (ADS)

    Tanaka, Shuji; Isomura, Kousuke; Togo, Shin-ichi; Esashi, Masayoshi

    2004-11-01

    This paper describes a turbo test rig to test the compressor of a palmtop gas turbine generator at low temperature (<100 °C). Impellers are 10 mm in diameter and have three-dimensional blades machined using a five-axis NC milling machine. Hydroinertia bearings are employed in both radial and axial directions. The performance of the compressor was measured at 50% (435 000 rpm) and 60% (530 000 rpm) of the rated rotational speed (870 000 rpm) by driving a turbine using compressed air at room temperature. The measured pressure ratio is lower than the predicted value. This could be mainly because impeller tip clearance was larger than the designed value. The measured adiabatic efficiency is unrealistically high due to heat dissipation from compressed air. During acceleration toward the rated rotational speed, a shaft crashed to the bearing at 566 000 rpm due to whirl. At that time, the whirl ratio was 8.

  7. Design considerations and experimental observations for the TAMU air-cooled reactor cavity cooling system for the VHTR

    SciTech Connect

    Sulaiman, S. A. Dominguez-Ontiveros, E. E. Alhashimi, T. Budd, J. L. Matos, M. D. Hassan, Y. A.

    2015-04-29

    The Reactor Cavity Cooling System (RCCS) is a promising passive decay heat removal system for the Very High Temperature Reactor (VHTR) to ensure reliability of the transfer of the core residual and decay heat to the environment under all off-normal circumstances. A small scale experimental test facility was constructed at Texas A and M University (TAMU) to study pertinent multifaceted thermal hydraulic phenomena in the air-cooled reactor cavity cooling system (RCCS) design based on the General Atomics (GA) concept for the Modular High Temperature Gas-Cooled Reactor (MHTGR). The TAMU Air-Cooled Experimental Test Facility is ⅛ scale from the proposed GA-MHTGR design. Groundwork for experimental investigations focusing into the complex turbulence mixing flow behavior inside the upper plenum is currently underway. The following paper illustrates some of the chief design considerations used in construction of the experimental test facility, complete with an outline of the planned instrumentation and data acquisition methods. Computational Fluid Dynamics (CFD) simulations were carried out to furnish some insights on the overall behavior of the air flow in the system. CFD simulations assisted the placement of the flow measurement sensors location. Preliminary experimental observations of experiments at 120oC inlet temperature suggested the presence of flow reversal for cases involving single active riser at both 5 m/s and 2.25 m/s, respectively and four active risers at 2.25 m/s. Flow reversal may lead to thermal stratification inside the upper plenum by means of steady state temperature measurements. A Particle Image Velocimetry (PIV) experiment was carried out to furnish some insight on flow patterns and directions.

  8. Design considerations and experimental observations for the TAMU air-cooled reactor cavity cooling system for the VHTR

    NASA Astrophysics Data System (ADS)

    Sulaiman, S. A.; Dominguez-Ontiveros, E. E.; Alhashimi, T.; Budd, J. L.; Matos, M. D.; Hassan, Y. A.

    2015-04-01

    The Reactor Cavity Cooling System (RCCS) is a promising passive decay heat removal system for the Very High Temperature Reactor (VHTR) to ensure reliability of the transfer of the core residual and decay heat to the environment under all off-normal circumstances. A small scale experimental test facility was constructed at Texas A&M University (TAMU) to study pertinent multifaceted thermal hydraulic phenomena in the air-cooled reactor cavity cooling system (RCCS) design based on the General Atomics (GA) concept for the Modular High Temperature Gas-Cooled Reactor (MHTGR). The TAMU Air-Cooled Experimental Test Facility is ⅛ scale from the proposed GA-MHTGR design. Groundwork for experimental investigations focusing into the complex turbulence mixing flow behavior inside the upper plenum is currently underway. The following paper illustrates some of the chief design considerations used in construction of the experimental test facility, complete with an outline of the planned instrumentation and data acquisition methods. Computational Fluid Dynamics (CFD) simulations were carried out to furnish some insights on the overall behavior of the air flow in the system. CFD simulations assisted the placement of the flow measurement sensors location. Preliminary experimental observations of experiments at 120oC inlet temperature suggested the presence of flow reversal for cases involving single active riser at both 5 m/s and 2.25 m/s, respectively and four active risers at 2.25 m/s. Flow reversal may lead to thermal stratification inside the upper plenum by means of steady state temperature measurements. A Particle Image Velocimetry (PIV) experiment was carried out to furnish some insight on flow patterns and directions.

  9. Water cooling system for an air-breathing hypersonic test vehicle

    NASA Technical Reports Server (NTRS)

    Petley, Dennis H.; Dziedzic, William M.

    1993-01-01

    This study provides concepts for hypersonic experimental scramjet test vehicles which have low cost and low risk. Cryogenic hydrogen is used as the fuel and coolant. Secondary water cooling systems were designed. Three concepts are shown: an all hydrogen cooling system, a secondary open loop water cooled system, and a secondary closed loop water cooled system. The open loop concept uses high pressure helium (15,000 psi) to drive water through the cooling system while maintaining the pressure in the water tank. The water flows through the turbine side of the turbopump to pump hydrogen fuel. The water is then allowed to vent. In the closed loop concept high pressure, room temperature, compressed liquid water is circulated. In flight water pressure is limited to 6000 psi by venting some of the water. Water is circulated through cooling channels via an ejector which uses high pressure gas to drive a water jet. The cooling systems are presented along with finite difference steady-state and transient analysis results. The results from this study indicate that water used as a secondary coolant can be designed to increase experimental test time, produce minimum venting of fluid and reduce overall development cost.

  10. Evaluation of Hybrid Air-Cooled Flash/Binary Power Cycle

    SciTech Connect

    Greg Mines

    2005-10-01

    Geothermal binary power plants reject a significant portion of the heat removed from the geothermal fluid. Because of the relatively low temperature of the heat source (geothermal fluid), the performance of these plants is quite sensitive to the sink temperature to which heat is rejected. This is particularly true of air-cooled binary plants. Recent efforts by the geothermal industry have examined the potential to evaporatively cool the air entering the air-cooled condensers during the hotter portions of a summer day. While the work has shown the benefit of this concept, air-cooled binary plants are typically located in regions that lack an adequate supply of clean water for use in this evaporative cooling. In the work presented, this water issue is addressed by pre-flashing the geothermal fluid to produce a clean condensate that can be utilized during the hotter portions of the year to evaporatively cool the air. This study examines both the impact of this pre-flash on the performance of the binary plant, and the increase in power output due to the ability to incorporate an evaporative component to the heat rejection process.

  11. Experimental study on corrugated cross-flow air-cooled plate heat exchangers

    SciTech Connect

    Kim, Minsung; Baik, Young-Jin; Park, Seong-Ryong; Ra, Ho-Sang; Lim, Hyug

    2010-11-15

    Experimental study on cross-flow air-cooled plate heat exchangers (PHEs) was performed. The two prototype PHEs were manufactured in a stack of single-wave plates and double-wave plates in parallel. Cooling air flows through the PHEs in a crosswise direction against internal cooling water. The heat exchanger aims to substitute open-loop cooling towers with closed-loop water circulation, which guarantees cleanliness and compactness. In this study, the prototype PHEs were tested in a laboratory scale experiments. From the tests, double-wave PHE shows approximately 50% enhanced heat transfer performance compared to single-wave PHE. However, double-wave PHE costs 30% additional pressure drop. For commercialization, a wide channel design for air flow would be essential for reliable performance. (author)

  12. Improvement to Air2Air Technology to Reduce Fresh-Water Evaporative Cooling Loss at Coal-Based Thermoelectric Power Plants

    SciTech Connect

    Ken Mortensen

    2011-12-31

    This program was undertaken to enhance the manufacturability, constructability, and cost of the Air2Air{TM} Water Conservation and Plume Abatement Cooling Tower, giving a validated cost basis and capability. Air2Air{TM} water conservation technology recovers a portion of the traditional cooling tower evaporate. The Condensing Module provides an air-to-air heat exchanger above the wet fill media, extracting the heat from the hot saturated moist air leaving in the cooling tower and condensing water. The rate of evaporate water recovery is typically 10% - 25% annually, depending on the cooling tower location (climate). This program improved the efficiency and cost of the Air2Air{TM} Water Conservation Cooling Tower capability, and led to the first commercial sale of the product, as described.

  13. 40 CFR 92.108 - Intake and cooling air measurements.

    Code of Federal Regulations, 2010 CFR

    2010-07-01

    ... during the test. Overall measurement accuracy must be ±2 percent of full-scale value of the measurement... full-scale value. The Administrator must be advised of the method used prior to testing. (2... measurements. (1) Air that has had its absolute humidity altered is considered humidity-conditioned air....

  14. A comparison of humid air turbine (HAT) cycle and combined-cycle power plants

    SciTech Connect

    Rao, A.D.; Francuz, V.J.; Shen, J.C.; West, E.W. )

    1991-03-01

    The Humid Air Turbine (HAT) cycle is a combustion turbine-based power generating cycle that provides an alternative to combined-cycle power generation. The HAT cycle differs from combined cycles in that it eliminates the steam turbine bottoming cycle by vaporizing water into the turbine's combustion air with heat obtained from the combustion turbine exhaust and other heat sources. This report presents the results of a study conducted by Fluor Daniel, Inc. for EPRI in which the HAT cycle was compared with combined-cycle plants in integration with the Texaco coal gasification process, and in natural gas-fired plants. The comparison of the coal gasification-based power plants utilizing the HAT cycle with Texaco coal gasification-based combined-cycle plants indicate that HAT cycle-based plants are less expensive and produce less environmental emissions. Whereas the combined-cycle plants require the use of expensive syngas coolers to achieve high efficiencies, the HAT cycle plants can achieve similar high efficiencies without the use of such equipment, resulting in a significant savings in capital cost and a reduction in levelized cost of electricity of up to 15%. In addition, HAT cycle plants produce very low levels of NO{sub x} emissions, possibly as little as 6 ppmv (dry, 15% O{sub 2} basis) without requiring the use of control technologies such as selective catalytic reduction. In natural gas-fired plants, the HAT cycle was calculated to have as much as a 4 percentage point gain in efficiency over the combined cycle and a potential for substantial reductions in NO{sub x} emissions, CO{sub 2} emissions, and water consumption. 71 figs., 74 tabs.

  15. Correlation of Cooling Data from an Air-Cooled Cylinder and Several Multicylinder Engines

    NASA Technical Reports Server (NTRS)

    Pinkel, Benjamin; Ellerbrock, Herman H , Jr

    1940-01-01

    The theory of engine-cylinder cooling developed in a previous report was further substantiated by data obtained on a cylinder from a Wright r-1820-g engine. Equations are presented for the average head and barrel temperatures of this cylinder as functions of the engine and the cooling conditions. These equations are utilized to calculate the variation in cylinder temperature with altitude for level flight and climb. A method is presented for correlating average head and barrel temperatures and temperatures at individual points on the head and the barrel obtained on the test stand and in flight. The method is applied to the correlation and the comparison of data obtained on a number of service engines. Data are presented showing the variation of cylinder temperature with time when the power and the cooling pressure drop are suddenly changed.

  16. Optimization of heat transfer in cooled shell elements of gas-turbine engines

    NASA Astrophysics Data System (ADS)

    Rodionov, N. G.; Grinkrug, M. S.

    1985-08-01

    A theoretical solution is presented for the problem of finding an optimum distribution of the coefficients of heat transfer from the coolant in the shell structures of gas-turbine engines. The approach proposed here provides a way to efficiently use the mechanical properties of materials, to optimize coolant distribution over the shell surface, and, ultimately to improve the economy and performance of gas-turbine engines.

  17. Recent developments on Air Liquide advanced technologies turbines

    NASA Astrophysics Data System (ADS)

    Delcayre, Franck; Gondrand, Cecile; Drevard, Luc; Durand, Fabien; Marot, Gerard

    2012-06-01

    Air Liquide Advanced Technologies has developed for more than 40 years turboexpanders mainly for hydrogen and helium liquefiers and refrigerators and has in total more than 600 references of cryogenic turbo-expanders and cold compressors. The latest developments are presented in this paper. The key motivation of these developments is to improve the efficiency of the machines, and also to widen the range of operation. New impellers have been designed for low and high powers, the operation range is now between 200W and 200kW. The thrust bearings have been characterized in order to maximize the load which can be withstood and to increase the turbo-expander cold power. Considering low power machines, 3D open wheels have been designed and machined in order to increase the adiabatic efficiencies. A new type of machine, a turbobooster for methane liquefaction has been designed, manufactured and tested at AL-AT test facility.

  18. Nano-textured copper oxide nanofibers for efficient air cooling

    NASA Astrophysics Data System (ADS)

    An, Seongpil; Jo, Hong Seok; Al-Deyab, Salem S.; Yarin, Alexander L.; Yoon, Sam S.

    2016-02-01

    Ever decreasing of microelectronics devices is challenged by overheating and demands an increase in heat removal rate. Herein, we fabricated highly efficient heat-removal coatings comprised of copper oxide-plated polymer nanofiber layers (thorny devil nanofibers) with high surface-to-volume ratio, which facilitate heat removal from the underlying hot surfaces. The electroplating time and voltage were optimized to form fiber layers with maximal heat removal rate. The copper oxide nanofibers with the thorny devil morphology yielded a superior cooling rate compared to the pure copper nanofibers with the smooth surface morphology. This superior cooling performance is attributed to the enhanced surface area of the thorny devil nanofibers. These nanofibers were characterized with scanning electron microscopy, X-ray diffraction, atomic force microscopy, and a thermographic camera.

  19. Intelligent Engine Systems: Thermal Management and Advanced Cooling

    NASA Technical Reports Server (NTRS)

    Bergholz, Robert

    2008-01-01

    The objective of the Advanced Turbine Cooling and Thermal Management program is to develop intelligent control and distribution methods for turbine cooling, while achieving a reduction in total cooling flow and assuring acceptable turbine component safety and reliability. The program also will develop embedded sensor technologies and cooling system models for real-time engine diagnostics and health management. Both active and passive control strategies will be investigated that include the capability of intelligent modulation of flow quantities, pressures, and temperatures both within the supply system and at the turbine component level. Thermal management system concepts were studied, with a goal of reducing HPT blade cooling air supply temperature. An assessment will be made of the use of this air by the active clearance control system as well. Turbine component cooling designs incorporating advanced, high-effectiveness cooling features, will be evaluated. Turbine cooling flow control concepts will be studied at the cooling system level and the component level. Specific cooling features or sub-elements of an advanced HPT blade cooling design will be downselected for core fabrication and casting demonstrations.

  20. A miniaturized piezoelectric turbine with self-regulation for increased air speed range

    NASA Astrophysics Data System (ADS)

    Fu, Hailing; Yeatman, Eric M.

    2015-12-01

    This paper presents the design and demonstration of a piezoelectric turbine with self-regulation for increased air speed range. The turbine's transduction is achieved by magnetic "plucking" of a piezoelectric beam by the passing rotor. The increased speed range is achieved by the self-regulating mechanism which can dynamically adjust the magnetic coupling between the magnets on the turbine rotor and the piezoelectric beam using a micro-spring. The spring is controlled passively by the centrifugal force of the magnet on the rotor. This mechanism automatically changes the relative position of the magnets at different rotational speeds, making the coupling weak at low airflow speeds and strong at high speeds. Hence, the device can start up with a low airflow speed, and the output power can be ensured when the airflow speed is high. A theoretical model was established to analyse the turbine's performance, advantages, and to optimize its design parameters. A prototype was fabricated and tested in a wind tunnel. The start-up airflow speed was 2.34 m/s, showing a 30% improvement against a harvester without the mechanism.