scholarly journals Modification of Prandtl Wind Tunnel

2018 ◽  
Author(s):  
Bohua Sun
Keyword(s):  
Wind Tunnel ◽  
Pressure Loss ◽  
Total Pressure ◽  
Noise Pollution ◽  
Variable Cross Section ◽  
Total Pressure Loss ◽  
Variable Cross ◽  
Circular Tunnel ◽  
Tunnel Design ◽  
Wind Tunnel Design

Wind tunnels are devices that enable researchers to study the flow over objects of interest, the forces acting on them and their interaction with the flow, which is nowadays playing an increasingly important role due to noise pollution. Since the first closed circuit wind tunnel with variable cross-section was built in G¨ottingen, its Prandtl configuration has little change. The wind tunnel with Prandtl configuration has four corners and vanes, more than 50% of the total pressure loss are caused by the corners and vanes. How to reduce the total pressure loss is a world class problem in the wind tunnel design. This study attempts to propose a novel configuration of wind tunnel, where the corners have been replaced by semi-circular tunnel. Sun wind tunnel 2 has only two corners and vanes, while Sun wind tunnel 1 has no  corners and vanes at all. It is expected the new wind tunnel can reduce the total pressure loss from 50% to 10%.

Numerical and Experimental Study of Swirling Flow in a Short Annular to Round Diffuser/Nozzle

2014 ◽  
Author(s):  
A. Namet-Allah ◽  
A. M. Birk
Keyword(s):  
Wind Tunnel ◽  
Pressure Loss ◽  
Total Pressure ◽  
Swirling Flow ◽  
Velocity Profiles ◽  
Back Pressure ◽  
Total Pressure Loss ◽  
Cfd Simulations ◽  
Measured Velocity ◽  
Numerical Predictions

The objective of the current paper is to gain an understanding of the effects of inlet swirling flow on the flow field through short annular transition diffusers and nozzles. These devices are representative of the primary driving nozzles for certain exhaust ejector systems. It is known that strongly swirling flow can degrade ejector performance due to core separation. It is believed that minor changes in driving nozzle shape can improve ejector performance significantly. Two configurations of a diffuser/nozzle were tested experimentally and numerically under different swirl strengths. The two configurations were mounted on an annular flow wind tunnel. Two shapes of the annulus’ centre body end; square and elliptical, were used. Based on the hydraulic inlet diameter, average velocity and temperature in the annulus of the wind tunnel, the measurements were carried out at Mach range of 0.21 to 0.26 with Reynolds number of 2.3 to 2.7×105. Ansys14 was used for the CFD simulations. The measured velocity profiles in the annulus were used as input flow conditions in the CFD investigation. The RNG k-ε turbulence model was used in the CFD simulations. The measured velocity profiles at the device exit, and measured surface pressures on the annulus, duct and nozzle walls were compared with the CFD predictions. The measured back pressure coefficient and total pressure loss through the diffuser systems were compared with the CFD predictions. A reasonable agreement between the experimental data and numerical predictions was observed. It was found computationally that the size of the central recirculation zone behind the annulus centre body has negative effects on the diffuser performance under different swirl numbers. The square shape of the annulus’ centre body end increased the back pressure and total pressure loss coefficients over the elliptical shape. However, the flow uniformity at the duct and nozzle exits improved with the square shape over the elliptical end. These differences may have a significant effect on ejector pumping.

Lift and Drag Distributions of Yacht Sails Using Wake Surveys

1996 ◽  
Vol 118 (2) ◽  
pp. 346-351 ◽  
Author(s):  
N. J. Locke ◽  
P. S. Jackson ◽  
R. G. J. Flay
Keyword(s):  
Wind Tunnel ◽  
Numerical Methods ◽  
Flow Structure ◽  
Stream Function ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Lift And Drag

The theory and numerical methods used to determine the lift and drag of a yacht model from wind-tunnel wake measurements are discussed. The results from wake surveys taken downstream of a l/40th scale IACC model yacht are presented. The lift and drag distributions are given as well as plots of the total pressure loss, vorticity, and stream-function, which are helpful in understanding the flow structure.

Experimental Study of Combustion on a Micro Gas Turbine Combustor

2008 ◽  
Author(s):  
Feng-Shan Wang ◽  
Wen-Jun Kong ◽  
Bao-Rui Wang
Keyword(s):  
Gas Turbine ◽  
Pressure Loss ◽  
Loss Factor ◽  
Total Pressure ◽  
Combustion Efficiency ◽  
Total Pressure Loss ◽  
Test Facility ◽  
Inlet Temperature ◽  
Micro Gas Turbine ◽  
Oil Fuel

A research program is in development in China as a demonstrator of combined cooling, heating and power system (CCHP). In this program, a micro gas turbine with net electrical output around 100kW is designed and developed. The combustor is designed for natural gas operation and oil fuel operation, respectively. In this paper, a prototype can combustor for the oil fuel was studied by the experiments. In this paper, the combustor was tested using the ambient pressure combustor test facility. The sensors were equipped to measure the combustion performance; the exhaust gas was sampled and analyzed by a gas analyzer device. From the tests and experiments, combustion efficiency, pattern factor at the exit, the surface temperature profile of the outer liner wall, the total pressure loss factor of the combustion chamber with and without burning, and the pollutants emission fraction at the combustor exit were obtained. It is also found that with increasing of the inlet temperature, the combustion efficiency and the total pressure loss factor increased, while the exit pattern factor coefficient reduced. The emissions of CO and unburned hydrogen carbon (UHC) significantly reduced, but the emission of NOx significantly increased.

Numerical investigations on the hydrogen jet pressure variations in a strut based scramjet combustor

2021 ◽  
Vol ahead-of-print (ahead-of-print) ◽  
Author(s):  
Jeyakumar Suppandipillai ◽  
Jayaraman Kandasamy ◽  
R. Sivakumar ◽  
Mehmet Karaca ◽  
Karthik K.
Keyword(s):  
Shock Wave ◽  
Supersonic Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Injection Pressure ◽  
Total Pressure Loss ◽  
Normal Shock ◽  
Content Type ◽  
Normal Shock Wave ◽  
Scramjet Combustor

Purpose This paper aims to study the influences of hydrogen jet pressure on flow features of a strut-based injector in a scramjet combustor under-reacting cases are numerically investigated in this study. Design/methodology/approach The numerical analysis is carried out using Reynolds Averaged Navier Stokes (RANS) equations with the Shear Stress Transport k-ω turbulence model in contention to comprehend the flow physics during scramjet combustion. The three major parameters such as the shock wave pattern, wall pressures and static temperature across the combustor are validated with the reported experiments. The results comply with the range, indicating the adopted simulation method can be extended for other investigations as well. The supersonic flow characteristics are determined based on the flow properties, combustion efficiency and total pressure loss. Findings The results revealed that the augmentation of hydrogen jet pressure via variation in flame features increases the static pressure in the vicinity of the strut and destabilize the normal shock wave position. Indeed, the pressure of the mainstream flow drives the shock wave toward the upstream direction. The study perceived that once the hydrogen jet pressure is reached 4 bar, the incoming flow attains a subsonic state due to the movement of normal shock wave ahead of the strut. It is noticed that the increase in hydrogen jet pressure in the supersonic flow field improves the jet penetration rate in the lateral direction of the flow and also increases the total pressure loss as compared with the baseline injection pressure condition. Practical implications The outcome of this research provides the influence of fuel injection pressure variations in the supersonic combustion phenomenon of hypersonic vehicles. Originality/value This paper substantiates the effect of increasing hydrogen jet pressure in the reacting supersonic airstream on the performance of a scramjet combustor.

Film Cooling and Aerodynamic Performance on Multi-Cavity Squealer Tip of a Turbine Blade

2021 ◽  
Author(s):  
Feng Li ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Cooling Performance ◽  
Pressure Loss Coefficient ◽  
Blade Tip ◽  
Total Pressure Loss Coefficient

Abstract The blade tip region of the shroud-less high-pressure gas turbine is exposed to an extremely operating condition with combined high temperature and high heat transfer coefficient. It is critical to design new tip structures and apply effective cooling method to protect the blade tip. Multi-cavity squealer tip has the potential to reduce the huge thermal loads and improve the aerodynamic performance of the blade tip region. In this paper, numerical simulations were performed to predict the aerothermal performance of the multi-cavity squealer tip in a heavy-duty gas turbine cascade. Different turbulence models were validated by comparing to the experimental data. It was found that results predicted by the shear-stress transport with the γ-Reθ transition model have the best precision. Then, the film cooling performance, the flow field in the tip gap and the leakage losses were presented with several different multi-cavity squealer tip structures, under various coolant to mainstream mass flow ratios (MFR) from 0.05% to 0.15%. The results show that the ribs in the multi-cavity squealer tip could change the flow structure in the tip gap for that they would block the coolant and the leakage flow. In this study, the case with one-cavity (1C) achieves the best film cooling performance under a lower MFR. However, the cases with multi-cavity (2C, 3C, 4C) show higher film cooling effectiveness under a higher MFR of 0.15%, which are 32.6%%, 34.2%% and 41.0% higher than that of the 1C case. For the aerodynamic performance, the case with single-cavity has the largest total pressure loss coefficient in all MFR studied, whereas the case with two-cavity obtains the smallest total pressure loss coefficient, which is 7.6% lower than that of the 1C case.

Conjugate Heat Transfer Characteristics of Double Wall Cooling With Gradient Diameter of Film and Impingement Holes

2021 ◽  
Author(s):  
Juan He ◽  
Qinghua Deng ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Flow Direction ◽  
Turbine Blades ◽  
Total Pressure Loss ◽  
Advanced Technique ◽  
Cooling Effectiveness ◽  
Uniform Pattern ◽  
Double Wall

Abstract Double wall cooling, consisting of internal impingement cooling and external film cooling, is believed to be the most advanced technique in modern turbine blades cooling. In this paper, to improve the uniformity of temperature distribution, a flat plate double wall cooling model with gradient diameter of film and impingement holes was proposed, and the heat transfer and flow characteristics were investigated by solving steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations with SST k-ω turbulence model. The influence of gradient diameter on overall cooling effectiveness and total pressure loss was studied by comparing with the uniform pattern at the blowing ratios ranging from 0.5 to 2. For gradient diameter of film hole patterns, results show that −10% film pattern always has the lowest film flow non-uniformity coefficient. The laterally averaged overall cooling effectiveness of uniform pattern lies between that of +10% and −10% film patterns, but the intersection of three patterns moves upstream from the middle of flow direction with the increase of blowing ratio. Therefore, the −10% film pattern exerts the highest area averaged cooling effectiveness, which is improved by up to 1.6% and 1% at BR = 0.5 and 1 respectively compared with a uniform pattern. However, at higher blowing ratios, the +10% film pattern maintains higher cooling effectiveness and lower total pressure loss. For gradient diameter of impingement hole patterns, the intersection of laterally averaged overall cooling effectiveness in three patterns is located near the middle of flow direction under all blowing ratios. The uniform pattern has the highest area averaged cooling effectiveness and the smallest non-uniform coefficient, but the −10% jet pattern has advantages of reducing pressure loss, especially in the laminated loss.

Second Vane Total Pressure Loss Due to Endwall Iceform Contouring

2008 ◽  
Author(s):  
Ronald S. LaFleur
Keyword(s):  
Heat Transfer ◽  
Secondary Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Drag ◽  
Pressure Coefficient ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Surface Shape ◽  
Exit Plane

The iceformation design method generates an endwall contour, altering the secondary flows that produce elevated endwall heat transfer load and total pressure losses. Iceformation is an analog to regions of metal melting where a hot fluid alters the isothermal surface shape of a part as it is maintained by a cooling fluid. The passage flow, heat transfer and geometry evolve together under the constraints of flow and thermal boundary conditions. The iceformation concept is not media dependent and can be used in analogous flows and materials to evolve novel boundary shapes. In the past, this method has been shown to reduce aerodynamic drag and total pressure loss in flows such as diffusers and cylinder/endwall junctures. A prior paper [1] showed that the Reynolds number matched iceform geometry had a 24% lower average endwall heat transfer than the rotationally symmetric endwall geometry of the Energy Efficiency Engine (E3). Comparisons were made between three endwall geometries: the ‘iceform’, the ‘E3’ and the ‘flat’ as a limiting case of the endwall design space. This paper adds to the iceformation design record by reporting the endwall aerodynamic performances. Second vane exit flow velocities and pressures were measured using an automated 2-D traverse of a 1.2 mm diameter five-hole probe. Exit plane maps for the three endwall geometries are presented showing the details of the total pressure coefficient contours and the velocity vectors. The formation of secondary flow vortices is shown in the exit plane and this results in an impact on exit plane total pressure loss distribution, off-design over- and under-turning of the exit flow. The exit plane contours are integrated to form overall measures of the total pressure loss. Relative to the E3 endwall, the iceform endwall has a slightly higher total pressure loss attributed to higher dissipation of the secondary flow within the passage. The iceform endwall has a closer-to-design exit flow pattern than the E3 endwall.

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