Detailed Performance Comparison of a Dump and Short Faired Combustor Diffuser System

1994 ◽  
Vol 116 (3) ◽  
pp. 517-526 ◽  
Author(s):  
J. F. Carrotte ◽  
P. A. Denman ◽  
A. P. Wray ◽  
P. Fry
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Performance Comparison ◽  
Total Pressure Loss ◽  
Compressor Blades ◽  
High Loss ◽  
Flame Tube ◽  
Design Concepts ◽  
Recirculating Flow ◽  
Flame Tube Head

A rectangular model simulating four sectors of a combustion chamber was used to compare the performance of a standard dump diffuser, of overall length 180 mm, with that of a faired design 25.5 mm shorter. The performance of each system was assessed in terms of total pressure loss and static pressure recovery between prediffuser inlet and the annuli surrounding the flame tube. Since the program objective was to test design concepts only, no allowance was made for the presence of burner feed arms or flame tube support pins. In addition, tests were performed with relatively low levels of inlet turbulence and no wake mixing effects from upstream compressor blades. Relative to the dump design, the mass weighted total pressure loss to the outer and inner annuli was reduced by 30 and 40 percent, respectively, for the faired diffuser. Measurements around the flame tube head were used to identify regions of high loss within each system and account for the differences in performance. Within a dump diffuser the flow separates at prediffuser exit resulting in a free surface diffusion around the flame tube head and a recirculating flow in the dump cavity. This source of loss is eliminated in the faired system where the flow remains attached to the casings. Furthermore, the faired system exhibited similar velocity magnitudes and gradients around the combustor head despite its shorter length.

scholarly journals Detailed Performance Comparison of a Dump and Short Faired Combustor Diffuser System

1993 ◽  
Author(s):  
J. F. Carrotte ◽  
P. A. Denman ◽  
A. P. Wray ◽  
P. Fry
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Performance Comparison ◽  
Total Pressure Loss ◽  
Compressor Blades ◽  
High Loss ◽  
Flame Tube ◽  
Design Concepts ◽  
Recirculating Flow ◽  
Flame Tube Head

A rectangular model simulating 4 sectors of a combustion chamber was used to compare the performance of a standard dump diffuser, of overall length 180mm, with that of a faired design 25.5mm shorter. The performance of each system was assessed in terms of total pressure loss and static pressure recovery between pre-diffuser inlet and the annuli surrounding the flame tube. Since the program objective was to test design concepts only, no allowance was made for the presence of burner feed arms or flame tube support pins. In addition, tests were performed with relatively low levels of inlet turbulence and no wake mixing effects from upstream compressor blades. Relative to the dump design, the mass weighted total pressure loss to the outer and inner annuli was reduced by 30% and 40% respectively for the faired diffuser. Measurements around the flame tube head were used to identify regions of high loss within each system and account for the differences in performance. Within a dump diffuser the flow separates at pre-diffuser exit resulting in a free surface diffusion around the flame tube head and a recirculating flow in the dump cavity. This source of loss is eliminated in the faired system where the flow remains attached to the inner casing. Furthermore, the faired system exhibited similar velocity magnitudes and gradients around the combustor head despite its shorter length.

Performance Evaluation of Linear Turbine Cascades Using Three-Dimensional Viscous Flow Calculations

1985 ◽  
Vol 107 (4) ◽  
pp. 969-975 ◽  
Author(s):  
J. Moore ◽  
J. G. Moore
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Surface Flow ◽  
Total Pressure Loss ◽  
Suction Surface ◽  
Cross Sectional ◽  
High Loss ◽  
Flow Development ◽  
Turbine Cascades

The overall performance of two geometrically similar linear turbine cascades is calculated using an elliptic flow program. The increase in the mass-averaged total pressure loss is calculated within and downstream of the cascades and the results show good agreement with the measured values. The buildup and decay of the secondary kinetic energy are also shown; measurements are available for one of the cascades near and downstream of the trailing edge and these are in close agreement with the calculated values. Details of the flow development are also compared with measurements. Calculated velocity vectors near the endwall show the overturning revealed by surface flow visualization and similarly near the suction surface the strong spanwise flow is well calculated. Calculated contours of total pressure loss in cross-sectional planes confirm the important interaction of the passage vortex with the profile boundary layer at midspan. Regions of high loss near midspan are calculated downstream of both cascades; this three-dimensional flow development is followed in the calculations.

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.

scholarly journals Numerical and Experimental Investigations on Optimized 3D Compressor Airfoils

2018 ◽  
Author(s):  
Jan Mihalyovics ◽  
Christian Brück ◽  
Dieter Peitsch ◽  
Ilias Vasilopoulos ◽  
Marcus Meyer
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Experimental Studies ◽  
Total Pressure Loss ◽  
Design Parameters ◽  
Flow Conditions ◽  
Flow Angle ◽  
Flow Phenomena ◽  
The Impact ◽  
Stator Design

The objective of the presented work is to perform numerical and experimental studies on compressor stators. This paper presents the modification of a baseline stator design using numerical optimization resulting in a new 3D stator. The Rolls Royce in-house compressible flow solver HYDRA was employed to predict the 3D flow, solving the steady RANS equations with the Spalart-Allmaras turbulence model, and its corresponding discrete adjoint solver. The performance gradients with respect to the input design parameters were used to optimize the stator blade with respect to the total pressure loss over a prescribed incidence range, while additionally minimizing the flow deviation from the axial direction at the stator exit. Non-uniform profile boundary conditions, being derived from the experimental measurements, have been defined at the inlet of the CFD domain. The presented results show a remarkable decrease in the axial exit flow angle deviation and a minor decrease in the total pressure loss. Experiments were conducted on two compressor blade sets investigating the three-dimensional flow in an annular compressor stator cascade. Comparing the baseline flow of the 42° turning stator shows that the optimized stator design minimizes the secondary flow phenomena. The experimental investigation discusses the impact of steady flow conditions on each stator design while focusing on the comparison of the 3D optimized design to the baseline case. The flow conditions were investigated using five-hole probe pressure measurements in the wake of the blades. Furthermore, oil-flow visualization was applied to characterize flow phenomena. These experimental results are compared with the CFD calculations.

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