Relevant bibliographies by topics / Nozzle guide vane

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  1. Journal articles
  2. Dissertations / Theses
  3. Books
  4. Book chapters
  5. Conference papers

Academic literature on the topic 'Nozzle guide vane'

Author: Grafiati
Published: 4 June 2021
Last updated: 21 February 2023

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Journal articles on the topic "Nozzle guide vane"

1

Plante, Robert D. "The Nozzle Guide Vane Problem." Operations Research 36, no. 1 (February 1988): 18–33. http://dx.doi.org/10.1287/opre.36.1.18.

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2

Yang, Dengfeng, Ce Yang, Dazhong Lao, and Tao Zeng. "A detailed investigation of a variable nozzle turbine with novel forepart rotation guide vane." Proceedings of the Institution of Mechanical Engineers, Part D: Journal of Automobile Engineering 233, no. 4 (February 25, 2018): 994–1007. http://dx.doi.org/10.1177/0954407018757244.

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One of the disadvantages of a variable nozzle turbine in practical application is the stage performance degradation due to nozzle endwall leakage flow at small nozzle openings. Aiming at restricting the nozzle leakage flow rate to improve turbine stage performance, a novel forepart rotation guide vane has been proposed and numerically studied in present work. First, the numerical results of baseline turbine were validated by experimental data to ensure the accuracy of numerical methods. Then steady and unsteady simulations were performed on both baseline and forepart rotation guide vane turbines to demonstrate the effectiveness of the novel vane and to study the characteristics of nozzle leakage flow, respectively. Results indicate that there is up to 13.5% peak efficiency improvement that has been achieved at 10% nozzle opening with the forepart rotation guide vane design; besides, rotor–stator interaction for forepart rotation guide vane is also mitigated due to the reduced nozzle leakage flow rate, thus the intensity of loading fluctuation on rotor blades is weakened significantly, which is beneficial to improve rotor blade forced response.
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3

Shaikh, Faisal, and Budimir Rosic. "Unsteady phenomena at the combustor-turbine interface." Journal of the Global Power and Propulsion Society 5 (November 23, 2021): 202–15. http://dx.doi.org/10.33737/jgpps/143042.

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The combustor-turbine interface in a gas turbine is characterised by complex, highly unsteady flows. In a combined experimental and large eddy simulation (LES) study including realistic combustor geometry, the standard model of secondary flows in the nozzle guide vanes (NGV) is found to be oversimplified. A swirl core is created in the combustion chamber which convects into the first vane passages. Four main consequences of this are identified: variation in vane loading; unsteady heat transfer on vane surfaces; unsteadiness at the leading edge horseshoe vortex, and variation in the position of the passage vortex. These phenomena occur at relatively low frequencies, from 50–300 Hz. It seems likely that these unsteady phenomena result in non-optimal film cooling, and that by reducing unsteadiness designs with greater cooling efficiency could be achieved. Measurements were performed in a high speed test facility modelling a large industrial gas turbine with can combustors, including nozzle guide vanes and combustion chambers. Vane surfaces and endwalls of a nozzle guide vane were instrumented with 384 high speed thin film heat flux gauges, to measure unsteady heat transfer. The high resolution of measurements was such to allow direct visualisation in time of large scale turbulent structures over the endwalls and vane surfaces. A matching LES simulation was carried out in a domain matching experimental conditions including upstream swirl generators and transition duct. Data reduction allowed time-varying LES data to be recorded for several cycles of the unsteady phenomena observed. The combination of LES and experimental data allows physical explanation and visualisation of flow events.
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4

Yang, Dengfeng, Kai Wang, Huaiyu Wang, Qian Zhang, Xinguo Lei, and Leon Hu. "An Investigation of the Performance and Internal Flow of Variable Nozzle Turbines with Split Sliding Guide Vanes." Machines 10, no. 11 (November 16, 2022): 1084. http://dx.doi.org/10.3390/machines10111084.

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In order to effectively weaken the leakage flow and shock intensity of traditional “swing” type guide vanes in a variable nozzle turbine, a new flow control device named the “split sliding guide vane” (SSGV) is studied in the present work. Steady and unsteady calculations were carried out on both the SSGV and base model at 10%, 40%, and 100% open positions. The shock test was performed to verify the accuracy of the numerical method. The results showed that at 10%, 40%, and 100% open positions, the leakage flow of the SSGV was 43%, 51%, and 40% of that of the base model, respectively. When 10% open, the turbine efficiency increased by 12%, compared with the base model, since the SSGV could effectively inhibit the clearance leakage flow. Due to the increased distance between the rotor and guide vane, the shock intensity of the SSGV was only 52% of that of the base model when 40% was open. The SSGV could reduce the static pressure loss on the guide vane pressure surface, but for the guide vane suction surface, the static pressure distribution appeared in a “W” shape due to the influence of the vane profile. Finally, the flow in the rotor was studied, which showed that the weakening of the shock and reduction of the clearance leakage flow in the guide vane were also beneficial for the strength of downstream rotor blades.
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5

Balakrishnan, Anantaram, Robert Plante, and Richard Wong. "The nozzle guide vane problem: Partitioning a heterogeneous inventory." European Journal of Operational Research 35, no. 3 (June 1988): 328–38. http://dx.doi.org/10.1016/0377-2217(88)90223-8.

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6

Pujari, Arun Kumar, B. V. S. S. S. Prasad, and Nekkanti Sitaram. "Conjugate Heat Transfer Analysis on the Interior Surface of Nozzle Guide Vane with Combined Impingement and Film Cooling." International Journal of Turbo & Jet-Engines 37, no. 4 (November 18, 2020): 327–42. http://dx.doi.org/10.1515/tjj-2017-0026.

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AbstractThe effect of conjugate heat transfer is investigated on a first stage nozzle guide vane (NGV) of a high pressure gas turbine which has both impingement and film cooling holes. The study is carried out computationally by considering a linear cascade domain, having two passages formed between the vanes, with a chord length of 228 mm and spacing of 200 mm. The effect of (i) coolant and mainstream Reynolds numbers, (ii) thermal conductivity (iii) temperature difference between the mainstream and coolant at the internal surface of the nozzle guide vane are investigated under conjugate thermal condition. The results show that, with increasing coolant Reynolds number the lower conducting material shows larger percentage decrease in surface temperature as compared to the higher conducting material. However, the internal surface temperature is nearly independent of mainstream Reynolds number variation but shows significant variation for higher conducting material. Further, the temperature gradient within the solid thickness of NGV is higher for the lower conductivity material.
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7

Sargison, J. E., S. M. Guo, M. L. G. Oldfield, G. D. Lock, and A. J. Rawlinson. "A Converging Slot-Hole Film-Cooling Geometry—Part 2: Transonic Nozzle Guide Vane Heat Transfer and Loss." Journal of Turbomachinery 124, no. 3 (July 1, 2002): 461–71. http://dx.doi.org/10.1115/1.1459736.

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This paper presents the first experimental measurements on an engine representative nozzle guide vane, of a new film-cooling hole geometry, a con¯vergings¯lot-hole¯ or console. The patented console geometry is designed to improve the heat transfer and aerodynamic performance of turbine vane and rotor blade cooling systems. These experiments follow the successful validation of the console design in low-speed flat-plate tests described in Part 1 of this paper. Stereolithography was used to manufacture a resin model of a transonic, engine representative nozzle guide vane in which seven rows of previously tested fan-shaped film-cooling holes were replaced by four rows of consoles. This vane was mounted in the annular vane ring of the Oxford cold heat transfer tunnel for testing at engine Reynolds numbers, Mach numbers and coolant to mainstream momentum flux ratios using a heavy gas to simulate the correct coolant to mainstream density ratio. Heat transfer data were measured using wide-band thermochromic liquid crystals and a modified analysis technique. Both surface heat transfer coefficient and the adiabatic cooling effectiveness were derived from computer-video records of hue changes during the transient tunnel run. The cooling performance, quantified by the heat flux at engine temperature levels, of the console vane compares favourably with that of the previously tested vane with fan-shaped holes. The new console film-cooling hole geometry offers advantages to the engine designer due to a superior aerodynamic efficiency over the fan-shaped hole geometry. These efficiency measurements are demonstrated by results from midspan traverses of a four-hole pyramid probe downstream of the nozzle guide vane.
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8

Krishnamoorthy, V., B. R. Pai, and S. P. Sukhatme. "Influence of Upstream Flow Conditions on the Heat Transfer to Nozzle Guide Vanes." Journal of Turbomachinery 110, no. 3 (July 1, 1988): 412–16. http://dx.doi.org/10.1115/1.3262212.

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The influence of a combustor located just upstream of a nozzle guide vane cascade on the heat flux distribution to the nozzle guide vane was experimentally investigated. The surface temperature distribution around the convectively cooled vane of the cascade was obtained by locating the cascade, firstly in a low-turbulence uniform hot gas stream, secondly in a high-turbulence, uniform hot gas stream, and thirdly in a high-turbulence, nonuniform hot gas stream present just downstream of the combustor exit. The results indicate that the increased blade surface temperatures observed for the cascade placed just downstream of the combustor can be accounted for by the prevailing turbulence level measured at cascade inlet in cold-flow conditions and the average gas temperature at the cascade inlet.
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9

Flaszynski, Pawel, Michal Piotrowicz, and Tommaso Bacci. "Clocking and Potential Effects in Combustor–Turbine Stator Interactions." Aerospace 8, no. 10 (October 2, 2021): 285. http://dx.doi.org/10.3390/aerospace8100285.

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Investigations of combustors and turbines separately have been carried out for years by research institutes and aircraft engine companies, but there are still many questions about the interaction effect. In this paper, a prediction of a turbine stator’s potential effect on flow in a combustor and the clocking effect on temperature distribution in a nozzle guide vane are discussed. Numerical simulation results for the combustor simulator and the nozzle guide vane (NGV) of the first turbine stage are presented. The geometry and flow conditions were defined according to measurements carried out on a test section within the framework of the EU FACTOR (full aerothermal combustor–turbine interactions research) project. The numerical model was validated by a comparison of results against experimental data in the plane at a combustor outlet. Two turbulence models were employed: the Spalart–Allmaras and Explicit Algebraic Reynolds Stress models. It was shown that the NGV potential effect on flow distribution at the combustor–turbine interface located at 42.5% of the axial chord is weak. The clocking effect due to the azimuthal position of guide vanes downstream of the swirlers strongly affects the temperature and flow conditions in a stator cascade.
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Baines, N. C., M. L. G. Oldfield, J. P. Simons, and J. M. Wright. "The Aerodynamic Development of a Highly Loaded Nozzle Guide Vane." Journal of Turbomachinery 108, no. 2 (October 1, 1986): 261–68. http://dx.doi.org/10.1115/1.3262046.

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A series of high-pressure turbine nozzle guide vanes has been designed for progressively increasing blade loading and reduction in blade solidity without additional loss penalty. Early members of the series achieved this by changes to the suction surface contour, but for the latest design the pressure surface contour was extensively modified to reduce the velocities on this surface substantially. Cascade testing revealed that this vane had a higher loss than its predecessor, and this appears to be largely due to a long region of boundary layer growth on the suction surface and possibly also an unsteady separation. These tests demonstrated the value of a flattened pitot tube held against the blade surface in determining the boundary layer state. By using a pitot probe of only modest frequency response (of order 100 Hz) it was possible to observe significant qualitative differences in the raw signals from laminar, transitional and turbulent boundary layers, which have previously been observed only with much higher frequency instruments. The test results include a comparison of boundary layer measurements on the same cascade test section in two different high-speed wind tunnels. This comparison suggests that freestream turbulence can have a large effect on boundary layer development and growth.
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Dissertations / Theses on the topic "Nozzle guide vane"

1

Agricola, Lucas. "Nozzle Guide Vane Sweeping Jet Impingement Cooling." The Ohio State University, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=osu1525436077557298.

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Harvey, Neil William. "Heat transfer on nozzle guide vane end walls." Thesis, University of Oxford, 1991. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.293454.

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3

Rowbury, David. "Discharge coefficients of nozzle guide vane film cooling holes." Thesis, University of Oxford, 1998. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.365838.

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4

Cresci, Irene. "High pressure nozzle guide vane cooling system flow characteristics." Thesis, University of Oxford, 2016. http://ora.ox.ac.uk/objects/uuid:b8826eb5-f4ad-4fe8-8730-9134fd9fd183.

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The continuous demand from the airlines for reduced jet engine fuel consumption results in increasingly challenging high pressure turbine nozzle guide vane (NGV) working conditions. The capability to reproduce representative boundary conditions in a rig at the combustor-turbine interaction plane is a key feature when testing NGVs in an engine-representative environment. A large scale linear cascade rig to investigate NGV leading edge cooling systems has been designed with particular attention being paid to creating engine representative conditions at the NGV inlet plane. The combustor simulator replicates the main features of a rich-burn design including large dilution jets and extensive endwall film cooling. CFD simulations have been used to develop the design which matches Reynolds number and mainstream-to-dilution jet momentum flux ratio. Detailed measurements of velocity, turbulence and temperature have been acquired at the NGV inlet plane. A thermo-couple was manufactured from 12.7 Î1⁄4m diameter wire and carefully calibrated to obtain its time constant in the velocity range of interest. The results are compared to CFD predictions and data in the literature. The time-averaged measurements show that the flow field conditions are dominated by the endwall cooling flows. The time-resolved data show that the measured turbulence length scale reflects the scale of the relevant upstream jets while the spectrum of temperature fluctuations reports a thermal cascade independent of any geometrical features. Attention was also focused on the flow field downstream of different endwall film cooling holes configurations: three arrangements of a double row of staggered cylindrical holes (lateral pitch-to-diameter ratio of 2 - 3 - 6) and one with intersecting holes (intersecting angle of 90o) were experimentally and numerically analyzed. The research quantified the extent by which closer spaced hole configurations provide more effective film coverage. It was found that the turbulent integral length scales are strongly connected to the hole diameter and spacing. It was also found that intersecting holes can potentially reduce the amount of required coolant at a fixed pressure ratio, but offer worst film performance than cylindrical holes. RANS simulations proved successful at predicting the main trends shown by the measurements. A new concept to increase the pressure margin across the film cooling holes in a specific region of vane LE coolant passage was introduced and developed: an insert was used to cover the area with the highest risk of ingestion, slowing down the flow and increasing the local static pressure. Numerical simulations were initially used to compare different designs and to analyse the impact of the insert on the overall coolant flow distribution. In particular, the effect on the static pressure downstream of the insert was identified as a critical factor that needs to be taken into account during the design process in order to avoid hot gas ingestion in other areas. The experimental campaign proved the ability of this new design to significantly increase the pressure margin in the covered region.
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Ceci, Alessandro. "Transonic Flow Features in a Nozzle Guide Vane Passage." Thesis, KTH, Farkost och flyg, 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-213986.

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The entropy noise in modern engines is mainly originating from two types of mechanisms.First, chemical reactions in the combustion chamber lead to unsteady heat releasewhich is responsible of the direct combustion noise. Second, hot and cold blobsof air coming from the combustion chamber are advected and accelerated throughturbine stages, giving rise to the so-called entropy noise (or indirect combustionnoise). In the present work, numerical characterization of indirect combustion noiseof a Nozzle Guide Vane passage was assessed using three-dimensional Large EddySimulations. The study was conducted on a simplified topology of a real turbinestator passage, for which experimental data were available in transonic operatingconditions. First, a baseline case was reproduced to validate a numerical finite volumesolver against the experimental measurements. Then, the same solver is used toreproduce the effects of incoming entropy waves from the combustion chamber andto characterize the additional generated acoustic power. Periodic temperature fluctuationsare imposed at the inlet, permitting to simulate hot and cold packets of aircoming from the unsteady combustion. The incoming waves are characterized bytheir characteristic wavelength; therefore, a parametric study has been conductedvarying the inlet temperature of the passage, generating entropy waves of greaterwavelengths. The study proves that the generated indirect combustion noise canbe significant. Moreover, the generated indirect combustion noise increases as thewavelength of the incoming disturbances increases. Finally, the present work suggeststhat, in transonic conditions, there might be flow features which enhance theindirect combustion noise generation mechanism.
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Lai, Cheng-Chyuan. "Fully film cooled nozzle guide vane heat transfer measurement and prediction." Thesis, University of Oxford, 1999. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.312115.

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Boccadamo, Danilo. "Numerical investigation of a transonic nozzle guide vane under elevated loading." Thesis, KTH, Energiteknik, 2016. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-200803.

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Despite many new investigations over the last years, there is no indication that alternative energy conversion technologies will overtake the place of turbomachinery. Hence thermal turbines are still the most dominant movers for electricity generation.Although this leadership in the energy production does not seem to be in danger, the current drivers in turbomachinery industry are to work towards both less fuel consumption and less pollution. In order to meet the future economic and environmental goals, researchers press towards highly loaded vanes and blades. This has to be performed at maintained or improved aerodynamic performances. Increased performances and blade loading lead in turn to increased velocities and larger regions of supersonic fluid velocities and consequently general increasing of shock intensities. The biggest problem dealing with supersonic flow and high shock intensities is that the boundary layer, when walking through these regions, experiences strong pressure gradients and intense shock-boundary layer interaction. This may lead the blade to stall meaning detachment of both boundary layer and cooling-film from the wall. These effects can evidently lead to catastrophic consequences since nowadays the materials used in turbomachinery applications have temperature strengths much lower than those coming from the combustion chamber. This thanks to very complex blade and vane cooling systems.There are even other features that may take benefit from increased velocities such as an attenuation in the boundary layer growth and the static pressure distribution on the blade surface. For helping researchers studying these new geometries, a cold air annular test rig designed by “Siemens Industrial Turbomachinery AB”, it has been built and placed at “Division of Heat and Power Technology” at KTH.The present thesis has the goal to provide a numerical model for CFD calculations, optimized for boundary layer studies, able to give a good prediction of detachment of the boundary layer and losses for different working cases. A previous model was provided with a commercial software for both ideal vane and real test rig. Recover of results and adaptions of the model were performed with a new version of the same software starting from the previous model. A comparison between numerical and experimental results have shown a good match for the subsonic and transonic case. Instead, problems were met for the supersonic case. Many attempts of different boundary condition at the inlet have been run. No reliable solution has been reached with realistic pressure profile at inlet while realistic results have been found using the mass flow rate as Inlet boundary condition. At the end, an analysis of shock and detachment is provided in terms of density gradient and static entropy distribution through the blade passage. Future works may aim to solve the “supersonic problem”.
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8

Bonilla, Carlos Humberto. "The Effect of Film Cooling on Nozzle Guide Vane Ash Deposition." The Ohio State University, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=osu1353961326.

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Rahim, Amir. "Effect of nozzle guide vane shaping on high pressure turbine stage performance." Thesis, University of Oxford, 2017. https://ora.ox.ac.uk/objects/uuid:35274ff0-0ea7-47bc-adc3-388f136b9555.

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This thesis presents a computational fluid dynamic (CFD) study of high pressure gas turbine blade design with different realistic inlet temperature and velocity boundary conditions. The effects of blade shaping and inlet conditions can only be fully understood by considering the aerodynamics and heat transfer concurrently; this is in contrast to the sequential method of blade design for aerodynamics followed by cooling. The inlet boundary conditions to the NGV simulations are governed by the existence of discrete fuel injectors in the combustion chamber. An appreciation of NGV shaping design under engine realistic inflow conditions will allow for an identification of the correct three dimensional shaping parameters that should be considered for design optimisation. The Rolls-Royce efficient Navier-Stokes solver, HYDRA, was employed in all computational results for a transonic turbine stage. The single passage unsteady method based on the Fourier Shape Correction is adopted. The solver is validated under both rich burn (hot steak only) and the case with swirl inlet profiles for aerothermal characteristics; good agreement is noted with the validation data. Post processing methods were used in order to obtain time-averaged results and blade visualisations. Subsequently, a surrogate design optimisation methodology using machine learning combined with a Genetic Algorithm is implemented and validated. A study of the effect of NGV compound lean on stage performance is carried out and contrasted for uniform and rich burn inlets, and subsequently for lean burn. Compound lean is shown to produce a tip uploading at the rotor inlet, which is beneficial for rich burn, but detrimental for lean burn. It is also found that for rich burn, fluid driving temperature is more dominant than HTC in determining rotor blade heat transfer, the opposite sense to the uniform inlet. Also, for a lean burn inlet, there is another role reversal, with HTC dominating fluid driving temperature in determining heat transfer. A novel NGV design methodology is proposed that seeks to mitigate the combined effects of inlet hot streak and swirling flow. In essence, the concept two NGVs in a pair are shaped independently of each other, thus allowing the inlet flow non uniformity to be suitably accommodated. Finally, two numerical NGV optimisation studies are undertaken for the combined hot streak and swirl inlet for two clocking positions; vane impinging and passage aligned. Due to the prohibitive cost of unsteady CFD simulations for an optimisation strategy, a suitable objective function at the NGV exit plane is used to minimise rotor tip heat flux. The optimised shape for the passage case resulted in the lowest tip heat flux distribution, however the optimum shape for the impinging case led to the highest gain in stage efficiency. This therefore suggests that NGV lean and clocking position should be a consideration for future optimisation and design of the HP stage.
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Nowlin, Scott Raymond. "The use of intersecting film cooling passages for nozzle guide vane cooling." Thesis, University of Oxford, 2009. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.670018.

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Books on the topic "Nozzle guide vane"

1

Plante, Robert. The nozzle guide vane problem. West Lafayette, Ind: Institute for Research in the Behavioral, Economic, and Management Sciences, Krannert Graduate School of Management, Purdue University, 1987.

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2

Balakrishnan, Anantaram. The nozzle guide vane problem: Partitioning a heterogeneous inventory. West Lafayette, Ind: Institute for Research in the Behavioral, Economic, and Management Sciences, Krannert Graduate School of Management, Purdue University, 1986.

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3

Boyle, Robert J. Heat transfer predictions for two turbine nozzle geometries at high Reynolds and Mach numbers. [Washington, D.C.]: National Aeronautics and Space Administration, 1995.

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Boyle, Robert J. Heat transfer predictions for two turbine nozzle geometries at high Reynolds and Mach numbers. [Washington, D.C.]: National Aeronautics and Space Administration, 1995.

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Book chapters on the topic "Nozzle guide vane"

1

Redding, L. E., Christopher J. Hockley, R. Roy, and J. Menhen. "The Role of Maintenance, Repair, and Overhaul (MRO) Knowledge in Facilitating Service Led Design: A Nozzle Guide Vane Case Study." In Lecture Notes in Mechanical Engineering, 379–95. Cham: Springer International Publishing, 2015. http://dx.doi.org/10.1007/978-3-319-15536-4_31.

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Qi, Jianhui. "Development and Analysis of Design Trends for Supercritical $${\mathrm{{CO}}_{2}}$$ Radial Inflow Turbine Nozzle Guide Vanes." In Simulation Tools and Methods for Supercritical Carbon Dioxide Radial Inflow Turbine, 101–28. Singapore: Springer Nature Singapore, 2022. http://dx.doi.org/10.1007/978-981-19-2860-4_4.

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Jiang, Lei-Yong, Yinghua Han, and Prakash Patnaik. "Conjugate Heat Transfer of an Internally Air-Cooled Nozzle Guide Vane and Shrouds." In Heat Transfer - Models, Methods and Applications. InTech, 2018. http://dx.doi.org/10.5772/intechopen.74104.

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Conference papers on the topic "Nozzle guide vane"

1

Girardeau, Julian, Frederic Pardo, Jérôme Pailhès, and Jean-Pierre Nadeau. "Nozzle Guide Vane Cooling System Optimization." In ASME Turbo Expo 2012: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2012. http://dx.doi.org/10.1115/gt2012-69600.

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The authors would like to address improvements on cooling system optimization within a turboshaft Nozzle Guide Vane (NGV). Designing high performance cooling systems able to preserve the life duration of the NGV can lead to significant aerodynamic losses. Theses losses jeopardize the performance of the whole engine. In the same time, a low efficiency cooling system may affect engine Maintenance Repair and Overhaul (MRO) costs as component life decreases. To help turbine designers, the authors studied a vane and searched for an optimal cooling design by means of an evolutionary algorithm. The associated objective function is based on satisfaction indexes, using Harrington’s desirability curves and Antonsson’s aggregation functions. Evaluation and optimization methods will be presented as well as optimized designs.
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Farahani, Arash, and Peter Childs. "Nozzle Guide Vane Static Strip Seals." In ASME Turbo Expo 2006: Power for Land, Sea, and Air. ASMEDC, 2006. http://dx.doi.org/10.1115/gt2006-90185.

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Sealing of components where there is no relative motion between the elements concerned remains a significant challenge in many gas turbine engine applications. Loss of sealing and cooling air from the internal air system through seals impacts on specific fuel consumption and can lead to undesirable flow interactions with resultant cost implications. For gas turbines, various strip seal types have been developed for use between Nozzle Guide Vanes in order to limit the flow of gas between the main stream annulus and the internal air system. Many different types of design have been proposed for overcoming strip seal problems such as misalignment of the grooves due to manufacturing and assembly constraints. In this paper various methods, with a particular focus on patents, for minimising the amount of leakage caused by such problems for strip seals between nozzle guide vanes are reviewed. By considering the advantages and disadvantages of each technique it is concluded that although apparently new strip seal designs for NGVs have improved performance, none of them can be considered to be ideal. This paper reviews the techniques and makes recommendations for future designs.
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3

DeBarmore, Nick, Paul King, Fred Schauer, and John Hoke. "Nozzle Guide Vane Integration into Rotating Detonation Engine." In 51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2013. http://dx.doi.org/10.2514/6.2013-1030.

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4

Kukutla, Pol Reddy, and B. V. S. S. S. Prasad. "Fluid Thermal Network Studies on Cooled Nozzle Guide Vane." In ASME 2019 Gas Turbine India Conference. American Society of Mechanical Engineers, 2019. http://dx.doi.org/10.1115/gtindia2019-2651.

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Abstract The aerothermal analysis is reported with the help of one-dimensional network modeling for the impingement cum film cooled gas turbine vane. The purpose of this one-dimensional simulation is to obtain the optimized film hole diameters of each row by analyzing the coolant flow distribution and overall effectiveness variations. FlownexR2017 commercial code is used to determine the detailed steady-state performance of the cooling vane. The results show that it is a useful simulation tool to obtain improved effectiveness of film cooling rows in a relatively short turn around time.
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5

Baines, N. C., M. L. G. Oldfield, J. P. Simons, and J. M. Wright. "The Aerodynamic Development of a Highly-Loaded Nozzle Guide Vane." In ASME 1986 International Gas Turbine Conference and Exhibit. American Society of Mechanical Engineers, 1986. http://dx.doi.org/10.1115/86-gt-229.

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A series of high pressure turbine nozzle guide vanes has been designed for progressively increasing blade loading and reduction in blade solidity without additional loss penalty. Early members of the series achieved this by changes to the suction surface contour, but for the latest design the pressure surface contour was extensively modified to reduce the velocities on this surface substantially. Cascade testing revealed that this vane had a higher loss than its predecessor, and this appears to be largely due to a long region of boundary layer growth on the suction surface and possibly also an unsteady separation. These tests demonstrated the value of a flattened pitot tube held against the blade surface in determining the boundary layer state. By using a pitot probe of only modest frequency response (of order 100 Hz) it was possible to observe significant qualitiative differences in the raw signals from laminar, transitional and turbulent boundary layers, which have previously been observed only with much higher frequency instruments. The test results include a comparison of boundary layer measurements on the same cascade test section in two different high-speed wind tunnels. This comparison suggests that freestream turbulence can have a large effect on boundary layer development and growth.
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6

Bonilla, C., C. Clum, M. Lawrence, B. Casaday, and J. P. Bons. "The Effect of Film Cooling on Nozzle Guide Vane Deposition." In ASME Turbo Expo 2013: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2013. http://dx.doi.org/10.1115/gt2013-95081.

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An accelerated deposition test facility was used to study the relationship between film cooling, surface temperature, and particle temperature on deposit formation. Tests were run at gas turbine representative inlet Mach numbers (0.1) and temperatures (1090°C). Deposits were created from lignite coal fly ash with mass median diameters of 1.3 and 8.8μm. Two CFM56-5B nozzle guide vane doublets, comprising three full passages and two half passages of flow, were utilized as the test articles. Tests were run with different levels of film cooling back flow margin and coolant temperature. Particle temperature upon impact with the vane surface was shown to be the leading factor in deposition. Since the particle must traverse the boundary layer of the cooled vane before impact, deposition is directly affected by the film and metal surface temperature as well. Film coolant jet strength showed only minor effect on deposit patterns on the leading edge. However, larger Stokes number (resulting in higher particle impact temperature) corresponded with increased deposit coverage area on the showerhead region. Additionally, infrared measurements showed a strong correlation between regions of greater deposits and elevated surface temperature on the pressure surface. Thickness distribution measurements also highlighted the effect of film cooling by showing reduced deposition immediately downstream of cooling holes. Implications for engine operation in particulate-laden environments are discussed.
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7

Casaday, B., A. Ameri, and J. P. Bons. "Numerical Investigation of Ash Deposition on Nozzle Guide Vane Endwalls." In ASME Turbo Expo 2012: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2012. http://dx.doi.org/10.1115/gt2012-68923.

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A computational study was performed to determine the factors that affect ash deposition rates on the endwalls in a nozzle guide vane passage. Deposition tests were simulated in flow around a flat plate with a cylindrical leading edge, as well as through a modern, high performance turbine vane passage. The flow solution was first obtained independent of the presence of particulates, and individual ash particles were subsequently tracked using a Langrangian tracking model. Two turbulence models were applied and their differences were discussed. The critical viscosity model was used to determine particle deposition. Features that contribute to endwall deposition, such as secondary flows, turbulent dispersion, or ballistic trajectories, were discussed and deposition was quantified. Particle sizes were varied, to reflect Stokes numbers ranging from 0.01 to 1.0, to determine the effect on endwall deposition. Results showed that endwall deposition rates can be as high as deposition on the leading edge for particles with a Stokes number less than 0.1, but endwall deposition rates for a Stokes numbers of 1.0 were less than 25% of the deposition rates on the leading edge or pressure surface of the turbine vane. Deposition rates on endwalls were largest near the leading edge stagnation region on both the cylinder and vane geometries, with significant deposition rates downstream showing a strong correlation to the secondary flows.
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8

Ragab, Kasem E., and Lamyaa El-Gabry. "Heat Transfer Analysis of the Surface of Nonfilm-Cooled and Film-Cooled Nozzle Guide Vanes in Transonic Annular Cascade." In ASME Turbo Expo 2017: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2017. http://dx.doi.org/10.1115/gt2017-64982.

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One of the approaches adopted to improve turbine efficiency and increase power to weight ratio is reducing vane count. In the current study, numerical analysis was performed for the heat transfer over the surface of nozzle guide vanes under the condition of reduced vane count using three dimensional computational fluid dynamics (CFD) models. The investigation has taken place in two stages: the baseline nonfilm-cooled nozzle guide vane, and the film-cooled nozzle guide vane. A finite volume based commercial code (ANSYS CFX 15) was used to build and analyze the CFD models. The investigated annular cascade has no heat transfer measurements available; hence in order to validate the CFD models against experimental data, two standalone studies were carried out on the NASA C3X vanes, one on the nonfilm-cooled C3X vane and the other on the film-cooled C3X vane. Different modelling parameters were investigated including turbulence models in order to obtain good agreement with the C3X experimental data, the same parameters were used afterwards to model the industrial nozzle guide vanes. Three Shear Stress Transport (SST) turbulence model variations were evaluated, the SST with Gamma-Theta transition model was found to yield the best agreement with the experimental results; model capabilities were demonstrated when the laminar to turbulent transition took place.
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9

Macoretta, Giuseppe, Bernardo Disma Monelli, Paolo Neri, Federico Bucciarelli, Damaso Checcacci, and Enrico Giusti. "Full-Scale Vibration Testing of Nozzle Guide Vanes." In ASME Turbo Expo 2021: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2021. http://dx.doi.org/10.1115/gt2021-59356.

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Abstract An increasing number of turboexpanders are equipped with Nozzle Guide Vane (NGV) as the first stator stage. By varying the throat area of the first stator vane the NGV enables an additional control methodology to the line-up power output allowing higher operational flexibility and higher efficiency at partial load and partial speed. The design of this component might become critical for enabling high expander availability considering its exposure to high temperature, thermal loading, and fluid induced vibrations. This is especially true also considering that the vibration frequencies of this sub-assembly are influenced by internal clearances and by the value of the friction coefficient, which leaves a relevant margin of error when using numerical methods (such as FEM) for predicting the actual structural behavior of this component. In this paper, the design of a full-scale test bench for the determination of both friction coefficients and modal behavior of a nozzle guide vane geometry is described. The bench enables us to simulate the pre-load due to aerodynamic forces on the NGV airfoil simulating the actual working conditions of bushes and bearings.
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10

Venkatasubramanya, S., S. A. Vasudev, and Sunil Chandel. "Experimental Evaluation of Cooling Effectiveness of High Pressure Turbine Nozzle Guide Vane." In ASME 2012 Gas Turbine India Conference. American Society of Mechanical Engineers, 2012. http://dx.doi.org/10.1115/gtindia2012-9558.

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High pressure turbine nozzle guide vane of a gas turbine engine, which operates at gas temperatures in excess of 1700 K, employs internal cooling, augmented convective cooling, impingement cooling and film cooling techniques to keep the vane in safe operating limits. Even though nozzle guide vanes are designed using heat transfer co-relations available in published papers and fundamental data, it is required to test the nozzle guide vane to ascertain the surface metal temperature and verify the adequacy of cooling. Adequacy of cooling is quantified by the term cooling effectiveness expressed and as percentage. The objective of the current work is to study the effect of gas to cooling air temperature ratio on cooling effectiveness. In the current study tests were first conducted to validate the test cascade in accordance with AGARD recommendations. Later tests were conducted to verify the constancy of cooling effectiveness across two gas temperatures and finally effect of gas to cooling air temperature ratio on cooling effectiveness was studied. The ratio was increased by a factor of 0.69 in leading edge and 0.72 in the trailing edge circuit and found that the cooling effectiveness remained constant.
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