Shroud Influence on Gas Turbine Airblast Atomizer Swirler Flowfields

1995 ◽  
Author(s):  
Gerald J. Micklow ◽  
Karthikeyan Shivaraman ◽  
Insoo Cho
Keyword(s):  
Velocity Profile ◽  
Velocity Component ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Pollutant Emissions ◽  
Swirl Number ◽  
Exit Velocity ◽  
Fuel Atomization ◽  
Swirl Vane

The performance of high shear airblast fuel injectors for advanced gas turbine combustors is highly dependent on the design of the swirl vanes. The vanes may be of the straight or curved type. Curved vanes usually exhibit lower losses but straight vanes are also used due to lower cost and ease of manufacture. These type of vanes often operate under highly stalled conditions with high total pressure loss and a highly non-uniform exit velocity profile. This may produce poor fuel atomization with a non-uniform combustor fuel distribution resulting in lowered combustor efficiency and increased pollutant emissions. Constant turning curved vanes can also be easily manufactured. Properly designed vanes result in a greatly reduced total pressure loss. The exit velocity distribution is more uniform and higher in magnitude which can result in improved fuel atomization and distribution in the combustor. Further, the presence of a shroud is seen to have a major effect on the downstream flowfield. The present study compares standard helical flat vane performance with a low loss curved vane designed by the author for idle, and cruise conditions both with and without a shroud. The results from a three dimensional viscous numerical flow simulation show the curved swirl vane to be clearly superior to the standard flat helical swirl vane. The curved vane has a much lower total pressure loss with a more uniform exit velocity profile. This may result in improved combustor and engine performance and reduced pollutant emissions. The effect of the shroud was seen to reduce the size of the stall cell found in the vane passage for the helical vane. This resulted in a decrease in the magnitude of the axial velocity component in the outer vane passage and a decrease in the circumferential velocity component. This may result in a decrease in the swirl number. For the curved vane however, an increase in the magnitude of all velocity components was found which will result in a higher swirl number and better nozzle performance.

Three Dimensional Performance Prediction of Advanced Swirl Vanes for Gas Turbine Airblast Atomizers

1993 ◽  
Author(s):  
Gerald J. Micklow ◽  
Karthikeyan Shivaraman ◽  
H. Lee Nguyen
Keyword(s):  
Velocity Profile ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Total Pressure Loss ◽  
Pollutant Emissions ◽  
Exit Velocity ◽  
Fuel Atomization ◽  
Swirl Vane

The performance of high shear airblast fuel injectors for advanced gas turbine combustors is highly dependent on the design of the swirl vanes. The vanes may be of the straight or curved type. Curved vanes usually exhibit lower losses but straight vanes are also used due to lower cost and ease of manufacture. These type of vanes often operate under highly stalled conditions with high total pressure loss and a highly non-uniform exit velocity profile. This can produce poor fuel atomization with a non-uniform combustor fuel distribution resulting in lowered combustor efficiency and increased pollutant emissions. Constant turning curved vanes can also be easily manufactured. Properly designed vanes result in a greatly reduced total pressure loss. The exit velocity distribution is more uniform and higher in magnitude resulting in improved fuel atomization and distribution in the combustor. The present study compares standard helical flat vane performance with a low loss curved vane designed by the author for idle, takeoff and cruise conditions. The results from a three dimensional viscous numerical flow simulation show the curved swirl vane to be clearly superior to the standard flat helical swirl vane. The curved vane has a much lower total pressure loss with a more uniform exit velocity profile. This will result in improved combustor and engine performance and reduced pollutant emissions.

scholarly journals Three Dimensional Analysis of Advanced Swirl Vane/Nozzle Assemblies

1996 ◽  
Author(s):  
Gerald J. Micklow ◽  
Michael Benjamin
Keyword(s):  
Pressure Gradient ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Adverse Pressure Gradient ◽  
Total Pressure Loss ◽  
Linear Variation ◽  
Exit Velocity ◽  
Fuel Atomization ◽  
Vane Tip

The performance of high shear axial inflow/radial outflow airblast fuel injectors for advanced gas turbine combustors is highly dependent on the design of the swirl vanes. Curved vanes usually exhibit lower losses but straight vanes are also used due to lower cost and ease of manufacture. These type of vanes often operate under highly stalled conditions with high total pressure loss and a highly non-uniform exit velocity profile. This may produce poor fuel atomization with a non-uniform combustor fuel distribution resulting in lowered combustor efficiency and increased pollutant emissions. Properly designed vanes result in a greatly reduced total pressure loss. The exit velocity distribution is more uniform and higher in magnitude which can result in improved fuel atomization and distribution in the combustor. The present study investigates two curved swirler/nozzle shroud configurations operating at 1 and 10 atmospheres pressure for the same inlet temperature of 293°K. The first configuration was a twisted curved vane with thickness where the turning angle varied non-linearly from hub to tip with a maximum turning at the tip of 70 degrees. The second configuration was a curved vane with a linear variation of turning with 70 degrees turning at the tip. The results from a three dimensional viscous numerical flow simulation of these configurations shows similar performance for all cases investigated. The non-linear twisted vane however, had an approximately 3% higher mass flow rate than the vane with the linear variation in turning for the same exit static pressure at the hub. One problem which existed for all the conditions analyzed was a high loss region near the vane tip. This was due to the interaction with the shroud. As the flow exits the vane row and progresses along the nozzle outer lip, the flow area increases. This condition along with the streamline curvature effect of the outer nozzle lip causes an adverse pressure gradient to be formed in this region. This adverse pressure gradient causes the flow to separate from the vane suction surface. The problem initiated in the region of 70% span and increased in magnitude to the vane tip.

Numerical Simulation of Cold Swirl Field for Solid Fuel Ramjet with NACA Airfoil

2014 ◽  
Vol 716-717 ◽  
pp. 711-716
Author(s):  
Jie Yu ◽  
Xiong Chen ◽  
Hong Wen Li
Keyword(s):  
Numerical Simulation ◽  
Flow Field ◽  
Radial Velocity ◽  
Solid Fuel ◽  
Velocity Gradient ◽  
Pressure Loss ◽  
Total Pressure ◽  
Swirl Flow ◽  
Total Pressure Loss ◽  
Swirl Number

In order to study the swirl flow characteristics in the solid fuel ramjet chamber, a new type of annular vane swirler with NACA airfoil is designed. The cold swirl flow field in the chamber is numerically simulated with different camber and t attack angle, while the swirl number , swirl flow field structure, total pressure recovery coefficient were studied. According to numerical simulation result, the main factors in swirl number are camber and angle of attack, the greater angle of attack, the greater the camber ,the stronger swirl will be. Results show that the total pressure loss is mainly concentrated in the inlet section, the total pressure loss cause by vane swirler is small. Radial velocity gradient exists in swirling flow, and increases with the swirl number. With the influence of centrifugal force and combustion chamber structure, the radial velocity gradient increases.

Numerical Simulation on Circumferential Staging of an Aeroengine LPP Combustor

2013 ◽  
Author(s):  
Ronghai Mao ◽  
Mingtao Shang
Keyword(s):  
Numerical Simulation ◽  
Temperature Distribution ◽  
Pressure Loss ◽  
Total Pressure ◽  
Combustion Efficiency ◽  
Total Pressure Loss ◽  
Nox Emission ◽  
Pollutant Emissions ◽  
Recirculation Region ◽  
Outlet Temperature

With the increasing stringency of the CAEP regulation on the pollutant emissions, combustors in lean burn architecture are being widely developed by aero-engine manufacturers to achieve low NOx emission performance with competitive margins to CAEP thresholds. A three-dimensional numerical simulation has been carried out in the present investigation to study an LPP combustor and circumferential staging effects on its main stage, for potential application to ACAE CJ-1000A aeroengine. A realizable k-ε turbulent model has been employed by the simulation, together with a Discrete Phase model based on Lagrangian methodology for the two-phase flow. The generic performances of the combustor, mainly in terms of flow and flame structures, fuel-air enhanced mixing performance, total pressure loss, combustion efficiency, outlet temperature distribution, and pollutant emissions have been analyzed. It was found that a large-scale central recirculation region is formed in the flame tube, which is beneficial to the stability of the combustion. The total pressure loss of the combustor is insensitive to the circumferential staging. Under approach mode the circumferential staging enhances the combustion efficiency from 73.8% without staging to 93.8% with staging; meanwhile the local turbulent flame speed increases more than two times. However the OTDF deteriorates from 0.30 without staging to 0.78 with staging, although the RTDF is found to be insensitive to the circumferential staging. The radial temperature distribution profiles are found to be pretty flat during the whole LTO cycle. The NOx emission without circumferential staging is simulated to be 68% reduction relative to CAEP 6. The circumferential staging, however, increases NOx emission to 65% reduction relative to CAEP 6. While gaining higher combustion efficiency, the major drawbacks of the circumferential staging are degradations of OTDF and NOx emission. Although the numerical results seem to be quite encouraging, the uncertainty of CFD results especially the temperature distribution and emissions might be tremendous. Experimental work has to follow up for further clarification.

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.

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