Heat Transfer Measurements Downstream of a Two-Dimensional Jet Entering a Crossflow

1987 ◽  
Vol 109 (4) ◽  
pp. 572-578 ◽  
Author(s):  
S. Wittig ◽  
V. Scherer
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Momentum Flux ◽  
Static Pressure ◽  
Flux Ratio ◽  
Main Flow ◽  
Two Dimensional ◽  
Reattachment Point ◽  
Static Pressure Distribution

Nusselt and Stanton numbers have been evaluated in and behind the recirculating zone produced by a two-dimensional jet entering a crossflow. The momentum flux ratio of the jet to the main flow was varied from 1.44–8.4 and measurements of the static pressure distribution and of the flow field by a five-hole probe were performed. A relation between the location of the reattachment point of the flow and the maximum of heat transfer was observed. Comparisons with available data are made. The experiments are intended for the verification of calculational codes.

Boundary-Layer Explorations Over a Two-Dimensional Mast/Sail Geometry

1990 ◽  
Vol 27 (04) ◽  
pp. 250-256
Author(s):  
Stuart Wilkinson
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Skin Friction ◽  
Static Pressure ◽  
Separated Flow ◽  
Velocity Profiles ◽  
Two Dimensional ◽  
Representative Data ◽  
Static Pressure Distribution

An experimental aerodynamic boundary-layer investigation is performed over the suction surfaces of a typical two-dimensional mast/sail geometry. Velocity profiles are obtained at a number of locations which, together with visualization data and the corresponding static pressure distribution, are used to describe the fundamental nature of the complex partially separated flow field associated with such geometries. The velocity profiles are fully analyzed to provide thickness parameters and skin friction coefficients, suitable for use as representative data in the development of predictive theories involving viscid-inviscid interactions. The chordwise variations of the thickness parameters are graphically presented and discussed.

The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 1—A Unique Experimental Facility

1987 ◽  
Vol 109 (4) ◽  
pp. 520-526 ◽  
Author(s):  
S. Deutsch ◽  
W. C. Zierke
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Static Pressure ◽  
Compressor Blade ◽  
Boundary Layer Transition ◽  
Two Dimensional ◽  
Suction Surface ◽  
Pressure Surface ◽  
Static Pressure Distribution

A unique cascade facility is described which permits the use of laser-Doppler velocimetry (LDV) to measure blade boundary layer profiles. Because of the need for a laser access window, the facility cannot reply on continuous blade pack suction to achieve two-dimensional, periodic flow. Instead, a strong suction upstream of the blade pack is used in combination with tailboards to control the flow field. The distribution of the upstream suction is controlled through a complex baffling system. A periodic, two–dimensional flow field is achieved at a chord Reynolds number of 500,000 and an incidence angle of 5 deg on a highly loaded, double circular arc, compressor blade. Inlet and outlet flow profiles, taken using five-hole probes, and the blade static-pressure distribution are used to document the flow field for use with the LDV measurements (see Parts 2 and 3). Inlet turbulence intensity is measured, using a hot wire, to be 0.18 percent. The static-pressure distribution suggests both separated flow near the trailing edge of the suction surface and an initially laminar boundary layer profile near the leading edge of the pressure surface. Probe measurements are supplemented by sublimation surface visualization studies. The sublimation studies place boundary layer transition at 64.2 ± 3.9 percent chord on the pressure surface, and indicate separation on the suction surface at 65.6 percent ± 3.5 percent chord.

Influence of Opposing Dilution Jets on Effusion Cooling

2021 ◽  
Author(s):  
M. Riley Creer ◽  
Karen A. Thole
Keyword(s):  
Gas Turbine ◽  
Momentum Flux ◽  
Static Pressure ◽  
Combustion Process ◽  
Flux Ratio ◽  
High Momentum ◽  
Cooling Effectiveness ◽  
Static Pressure Distribution ◽  
The Impact ◽  
Combustor Liner

Abstract The gas turbine combustion process reaches gas temperatures that exceed the melting temperature of the combustor liner materials. Cooling the liner is critical to combustor durability and is often accomplished with double-walled liners that contain both impingement and effusion holes. The liner cooling is complicated with the interruption of the effusion cooling by large dilution jets that facilitate the combustion process. Given the presence of the dilution jets, it is important to understand the effect that the dilution jet has on the opposing wall in respect to the effusion film. This research includes measurements of the local static pressure distribution for a range of dilution jet momentum flux ratios to investigate the impact that the opposing dilution jet has on the effusion film. The interactions with the effusion cooling were also evaluated by measuring the overall cooling effectiveness across the panel. Measurements show that the opposing dilution jets did impact the liner at dilution jet momentum flux ratios that were greater than 20. The impacts at high momentum flux ratios were indicated through increased local static pressures measured on the surface of the combustor liner. Furthermore, the dilution touchdown decreased the overall cooling effectiveness of the effusion cooling. Results also indicated that the opposing dilution jets changed position on the liner as the dilution jet momentum flux ratio changes.

Investigation of Flow Characteristics in a Combustor With Opposed Jets

2018 ◽  
Author(s):  
Yongbin Ji ◽  
Bing Ge ◽  
Shusheng Zang
Keyword(s):  
Flow Field ◽  
Gas Turbine ◽  
Momentum Flux ◽  
Flow Characteristics ◽  
Cross Flow ◽  
Flux Ratio ◽  
Main Flow ◽  
3D Flow ◽  
Momentum Flux Ratio ◽  
High Efficient

Jet-in-cross flow (JICF) has been investigated broadly because of its wide engineering application, for example in the gas turbine field, film cooling on the turbine vanes and blades, primary and dilution jets in the combustors and so on. In the gas turbine combustors, the main flow is generated by the swirlers to stabilize the flame, which induces complicated 3D flow characteristics. Different from uniform main flow, swirling cross flow has a strong tangential velocity component, which will deflect the jets in the circumferential direction as well as in the streamwise direction. So, the degradation behavior of the jets is more complex than that in the uniform cross flow. This paper presents PIV measurement of the flow field inside of a three-nozzle annular combustor with opposed quenching jets on the liner walls. Dry ice as a newly proposed flow tracer was proposed and tried. The momentum flux ratio and jet holes configuration are studied to evaluate their effects on the primary recirculation zone, downstream flow field. Finally, numerical simulation was also performed to reveal 3D flow characteristics as well as turbulent kinetic energy generation. The results show that momentum flux ratio has a dominant influence on flow characteristics in the combustor. Getting better understanding of jets behavior in the swirling cross flow helps optimization design of quenching or dilution holes geometry and arrangement for the gas turbine combustor, which turns to be very beneficial to the low-emission and high efficient combustor development.

Formation mechanism of static pressure circumferential double-peak distribution in a centrifugal compressor under the action of downstream boundary

2021 ◽  
pp. 095440702098307
Author(s):  
Botai Su ◽  
Ce Yang ◽  
Hanzhi Zhang ◽  
Xin Shi ◽  
Li Fu ◽  
...  
Keyword(s):  
Flow Field ◽  
Pressure Distribution ◽  
Mass Flow Rate ◽  
Formation Mechanism ◽  
Flow Rate ◽  
Centrifugal Compressor ◽  
Static Pressure ◽  
Double Peak ◽  
Static Pressure Distribution ◽  
Impeller Outlet

In the centrifugal compressor applied in the automobile turbochargers, the asymmetric structure of volute causes the non-uniform flow field in the impeller and compressor stall. The non-uniformity of the flow field in the compressor can be reflected by the casing-wall static pressure distribution. In this study, by removing the volute and directly imposing different simplified static pressure boundary conditions at the diffuser outlet, the formation mechanism of casing-wall static pressure circumferential double-peak distribution of the compressor is explored. It is found that the mass flow rate is redistributed at the impeller outlet due to local high static pressure induced by the volute tongue, which results in the formation of two airflow regions with high velocity in the diffuser, ultimately leading to the static pressure circumferential double-peak distribution in the diffuser and the impeller. Noted that because of the existence of the blades, the airflow regions with high and low velocity formed in the diffuser are locked within a limited range of one or more widths of the blade passage. When the number of blades in the compressor is large, the static pressure can appear as multi-peak distribution in the circumferential direction. Moreover, the result of the mass flow rate redistribution at the impeller outlet is determined by the static pressure distribution characteristics at the diffuser outlet.

Ingestion Into the Upstream Wheelspace of an Axial Turbine Stage

1994 ◽  
Vol 116 (2) ◽  
pp. 327-332 ◽  
Author(s):  
T. Green ◽  
A. B. Turner
Keyword(s):  
Pressure Distribution ◽  
Static Pressure ◽  
Three Dimensional ◽  
Computer Code ◽  
Rotor Blades ◽  
Rotor Speed ◽  
Turbine Stage ◽  
Flow Rates ◽  
Static Pressure Distribution ◽  
Axial Clearance

The upstream wheelspace of an axial air turbine stage complete with nozzle guide vanes (NGVs) and rotor blades (430 mm mean diameter) has been tested with the objective of examining the combined effect of NGVs and rotor blades on the level of mainstream ingestion for different seal flow rates. A simple axial clearance seal was used with the rotor spun up to 6650 rpm by drawing air through it from atmospheric pressure with a large centrifugal compressor. The effect of rotational speed was examined for several constant mainstream flow rates by controlling the rotor speed with an air brake. The circumferential variation in hub static pressure was measured at the trailing edge of the NGVs upstream of the seal gap and was found to affect ingestion significantly. The hub static pressure distribution on the rotor blade leading edges was rotor speed dependent and could not be measured in the experiments. The Denton three-dimensional C.F.D. computer code was used to predict the smoothed time-dependent pressure field for the rotor together with the pressure distribution downstream of the NGVs. The level and distribution of mainstream ingestion, and thus the seal effectiveness, was determined from nitrous oxide gas concentration measurements and related to static pressure measurements made throughout the wheelspace. With the axial clearance rim seal close to the rotor the presence of the blades had a complex effect. Rotor blades in connection with NGVs were found to reduce mainstream ingestion seal flow rates significantly, but a small level of ingestion existed even for very high levels of seal flow rate.

Numerical and Experimental Investigation of Return Channel Vane Aerodynamics With Two-Dimensional and Three-Dimensional Vanes

2016 ◽  
Vol 139 (1) ◽  
Author(s):  
A. Hildebrandt ◽  
F. Schilling
Keyword(s):  
Experimental Investigation ◽  
Flow Field ◽  
Static Pressure ◽  
Three Dimensional ◽  
Pressure Rise ◽  
Field Performance ◽  
Two Dimensional ◽  
Flow Angle ◽  
Overall Performance ◽  
Return Channel

The present paper deals with the numerical and experimental investigation of the effect of return channel (RCH) dimensions of a centrifugal compressor stage on the aerodynamic performance. Three different return channel stages were investigated, two stages comprising three-dimensional (3D) return channel blades and one stage comprising two-dimensional (2D) RCH vanes. The analysis was performed regarding both the investigation of overall performance (stage efficiency, RCH total pressure loss coefficient) and detailed flow-field performance. For detailed experimental flow-field investigation at the stage exit, six circumferentially traversed three-hole probes were positioned downstream the return channel exit in order to get two-dimensional flow-field information. Additionally, static pressure wall measurements were taken at the hub and shroud pressure and suction side (SS) of the 2D and 3D return channel blades. The return channel system overall performance was calculated by measurements of the circumferentially averaged 1D flow field downstream the diffuser exit and downstream the stage exit. Dependent on the type of return channel blade, the numerical and experimental results show a significant effect on the flow field overall and detail performance. In general, satisfactory agreement between computational fluid dynamics (CFD)-prediction and test-rig measurements was achieved regarding overall and flow-field performance. In comparison with the measurements, the CFD-calculated stage performance (efficiency and pressure rise coefficient) of all the 3D-RCH stages was slightly overpredicted. Very good agreement between CFD and measurement results was found for the static pressure distribution on the RCH wall surfaces while small CFD-deviations occur in the measured flow angle at the stage exit, dependent on the turbulence model selected.

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