406 Effect of Pressure Ratio and Exit Mach Number for Hysteresis Phenomenon of Shock Wave in Axisymmetric Over-Expanded Supersonic Jet(1)

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. CFD predictions of blade tip heat transfer are compared to test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; they are flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25, 2.0, and 2.75% of blade span. The tip heat transfer results of the numerical models agree fairly well with the data and are comparable to other CFD predictions in the open literature.

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Investigation of the 3D Unsteady Rotor Pressure Field in a HP Turbine Stage

Volume 5: Turbo Expo 2002, Parts A and B ◽

10.1115/gt2002-30365 ◽

2002 ◽

Cited By ~ 7

Author(s):

E. Valenti ◽

J. Halama ◽

R. De´nos ◽

T. Arts

Keyword(s):

Reynolds Number ◽

Mach Number ◽

Pressure Field ◽

Pressure Ratio ◽

Turbine Stage ◽

Time Resolved ◽

Mach Number Distribution ◽

Rotor Surface ◽

Exit Mach Number ◽

Unsteady Pressure Measurements

This paper presents steady and unsteady pressure measurements at three span locations (15, 50 and 85%) on the rotor surface of a transonic turbine stage. The data are compared with the results of a 3D unsteady Euler stage calculation. The overall agreement between the measurements and the prediction is satisfactory. The effects of pressure ratio and Reynolds number are discussed. The rotor time-averaged Mach number distribution is very sensitive to the pressure ratio of the stage since the incidence of the flow changes as well as the rotor exit Mach number. The time-resolved pressure field is dominated by the vane trailing edge shock waves. The incidence and intensity of the shock strongly varies from hub to tip due to the radial equilibrium of the flow at the vane exit. The decrease of the pressure ratio attenuates significantly the amplitude of the fluctuations. An increase of the pressure ratio has less significant effect since the change in the vane exit Mach number is small. The effect of the Reynolds number is weak for both the time-averaged and the time-resolved rotor static pressure at mid-span, while it causes an increase of the pressure amplitudes at the two other spans.

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Effects of Shock Wave Development on Secondary Flow Behavior in Linear Turbine Cascade at Transonic Condition

Volume 2B: Turbomachinery ◽

10.1115/gt2020-15772 ◽

2020 ◽

Author(s):

Hoshio Tsujita ◽

Masanao Kaneko

Keyword(s):

Shock Wave ◽

Mach Number ◽

Secondary Flow ◽

Turbine Blade ◽

Gas Turbines ◽

Flow Behavior ◽

Turbine Cascade ◽

Blade Loading ◽

Passage Vortex ◽

Exit Mach Number

Abstract Gas turbines widely applied to power generation and aerospace propulsion systems are continuously enhanced in efficiency for the reduction of environmental load. The energy recovery efficiency from working fluid in a turbine component constituting gas turbines can be enhanced by the increase of turbine blade loading. However, the increase of turbine blade loading inevitably intensifies the secondary flows, and consequently increases the associated loss generation. The development of the passage vortex is strongly influenced by the pitchwise pressure gradient on the endwall in the cascade passage. In addition, a practical high pressure turbine stage is generally driven under transonic flow conditions where the shock wave strongly influences the pressure distribution on the endwall. Therefore, it becomes very important to clarify the effects of the shock wave formation on the secondary flow behavior in order to increase the turbine blade loading without the deterioration of efficiency. In this study, the two-dimensional and the three-dimensional transonic flows in the HS1A linear turbine cascade at the design incidence angle were analyzed numerically by using the commercial CFD code with the assumption of steady compressible flow. The isentropic exit Mach number was varied from the subsonic to the supersonic conditions in order to examine the effects of development of shock wave caused by the increase of exit Mach number on the secondary flow behavior. The increase of exit Mach number induced the shock across the passage and increased its obliqueness. The increase of obliqueness reduced the cross flow on the endwall by moving the local minimum point of static pressure along the suction surface toward the trailing edge. As a consequence, the increase of exit Mach number attenuated the passage vortex.

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Development of a Supersonic Ejector for Capturing Very Low-Pressure Vent Gases and Re-Injection Into a High-Pressure Gas Stream

Volume 1: Project Management; Design and Construction; Environmental Issues; GIS/Database Development; Innovative Projects and Emerging Issues; Operations and Maintenance; Pipelining in Northern Environments; Standards and Regulations ◽

10.1115/ipc2006-10061 ◽

2006 ◽

Author(s):

K. K. Botros ◽

J. Geerligs ◽

H. Imran ◽

W. Thompson

Keyword(s):

Mach Number ◽

Optimum Design ◽

Nozzle Exit ◽

Pressure Ratio ◽

Supersonic Nozzle ◽

Low Pressure ◽

Fuel Gas ◽

Dry Gas Seal ◽

Supersonic Ejector ◽

Exit Mach Number

The purpose of the ejector device is to capture the gas leakage from a dry-gas seal at low pressure, and re-inject it into the fuel gas line to the gas generator (without the use of compressors or rotating elements), hence providing a means to utilize the gas that would otherwise be vented to atmosphere. Implementation of this device will also have the benefit of reducing greenhouse gas emissions to the atmosphere. The primary challenge to achieve the above goal lies in the fact that the leakage gas pressure is in the range of 70–340 kPag, while the minimum pressure required upstream of the fuel gas regulator is in the range of 2400–3300 kPag. The device consists of a two-stage supersonic ejector. The first stage is highly supersonic (nozzle exit Mach number ≃ 2.54), while the second stage is moderately supersonic (nozzle exit Mach number ≃ 1.72). Several tests where conducted on various configurations of the two stages on natural gas in order to arrive at the optimum design and operating parameters. The optimum design gave an expansion pressure ratio (motive/suction) of the order of 14.0 and compression pressure ratio (discharge/suction) of around 8.1. These ratios would meet the requirement of the minimum suction and discharge pressure mentioned above. This paper presents the optimum configuration arrived at after several iterations of different geometries of the supersonic nozzles, particularly for the first stage ejector, and presents the performance test results of the integrated system. The results indicate that the device would meet the requirements of capturing the low pressure, low flow dry gas seal leakage and re-inject it into the fuel gas stream with an overall ejector efficiency (based on thermodynamic availability) of 80%.

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Numerical Simulation of Supersonic Jet Control by Tabs with Slanted Perforation

International Journal of Turbo and Jet Engines ◽

10.1515/tjj-2019-0015 ◽

2019 ◽

Vol 0 (0) ◽

Author(s):

G. Ezhilmaran ◽

Suresh Chandra Khandai ◽

Yogesh Kumar Sinha ◽

S. Thanigaiarasu

Keyword(s):

Numerical Simulation ◽

Mach Number ◽

Near Field ◽

Pressure Ratio ◽

Supersonic Jet ◽

Pressure Gradients ◽

Nozzle Pressure Ratio ◽

Nozzle Pressure ◽

Jet Control ◽

Different Diameters

Abstract This paper presents the numerical simulation of Mach 1.5 supersonic jet with perforated tabs. The jet with straight perforation tab was compared with jets having slanted perforated tabs of different diameters. The perforation angles were kept as 0° and 10° with respect to the axis of the nozzle. The blockage areas of the tabs were 4.9 %, 4.9 % and 2.4 % for straight perforation, 10° slanted perforation ( {{{\Phi }}_{\ }} = 1.3 mm) and 10° slanted perforation ( {{{\Phi }}_{\ }} = 1.65 mm) respectively. The 3-D numerical simulations were carried out using the software. The mixing enhancements caused by these tabs were studied in the presence of adverse and favourable pressure gradients, corresponding to nozzle pressure ratio (NPR) of 3, 3.7 and 5. For Mach number 1.5 jet, NPR 3 corresponds to 18.92 % adverse pressure gradients and NPR 5 corresponds to 35.13 % favourable pressure gradients. The centerline Mach number of the jet with slanted perforations is found to decay at a faster rate than uncontrolled nozzle and jet with straight perforation tab. Mach number plots were obtained at both near-field and far field downstream locations. There is 25 % and 65 % reduction in jet core length were observed for the 0° and 10° perforated tabs respectively in comparison to uncontrolled jet.

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Numerical Modeling of Heat Transfer and Pressure Losses for an Uncooled Gas Turbine Blade Tip: Effect of Tip Clearance and Tip Geometry

Journal of Thermal Science and Engineering Applications ◽

10.1115/1.4000547 ◽

2009 ◽

Vol 1 (2) ◽

Cited By ~ 5

Author(s):

Lamyaa A. El-Gabry

Keyword(s):

Heat Transfer ◽

Mach Number ◽

Gas Turbine ◽

Computational Study ◽

Numerical Models ◽

Pressure Ratio ◽

Tip Clearance ◽

Pressure Losses ◽

Blade Tip ◽

Exit Mach Number

A computational study has been performed to predict the heat transfer distribution on the blade tip surface for a representative gas turbine first stage blade. Computational fluid dynamics (CFD) predictions of blade tip heat transfer are compared with test measurements taken in a linear cascade, when available. The blade geometry has an inlet Mach number of 0.3 and an exit Mach number of 0.75, pressure ratio of 1.5, exit Reynolds number based on axial chord of 2.57×106, and total turning of 110 deg. Three blade tip configurations were considered; a flat tip, a full perimeter squealer, and an offset squealer where the rim is offset to the interior of the tip perimeter. These three tip geometries were modeled at three tip clearances of 1.25%, 2.0%, and 2.75% of the blade span. The tip heat transfer results of the numerical models agree well with data. For the case in which side-by-side comparison with test measurements in the open literature is possible, the magnitude of the heat transfer coefficient in the “sweet spot” matches data exactly and shows 20–50% better agreement with experiment than prior CFD predictions of this same case.

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Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller

Journal of Turbomachinery ◽

10.1115/1.1791645 ◽

2004 ◽

Vol 126 (4) ◽

pp. 473-481 ◽

Cited By ~ 22

Author(s):

Hirotaka Higashimori ◽

Kiyoshi Hasagawa ◽

Kunio Sumida ◽

Tooru Suita

Keyword(s):

Boundary Layer ◽

Shock Wave ◽

Mach Number ◽

Centrifugal Compressor ◽

Transonic Flow ◽

Reverse Flow ◽

Pressure Ratio ◽

Layer Interaction ◽

Shock Boundary Layer Interaction ◽

Boundary Layer Interaction

Requirements for aeronautical gas turbine engines for helicopters include small size, low weight, high output, and low fuel consumption. In order to achieve these requirements, development work has been carried out on high efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. The centrifugal compressor is a high transonic compressor with an inlet Mach number of about 1.6. In high inlet Mach number compressors, the flow distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and Computational Fluid Dynamics (CFD). This report presents the measured results of the high transonic flow at the impeller inlet using Laser Doppler Velocimeter (LDV) and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution, and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460 m/s and the relative Mach number reaches about 1.6. Using a LDV, about 500 m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.

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