The non-reacting flow characteristics of pylon and wall injections in a dual combustion ramjet engine

2021 ◽  
pp. 146808742110178
Author(s):  
J Sarathkumar Sebastin ◽  
S Jeyakumar ◽  
K Karthik
Keyword(s):  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Reacting Flow ◽  
Gas Generator ◽  
Coaxial Jets ◽  
Ramjet Engine ◽  
Exit Nozzle ◽  
Wall Injection

The influence of pylon and wall injection in coaxial jets of a Dual Combustion Ramjet engine is numerically investigated in a non-reacting flow field. The supersonic combustor is modeled and analyzed using the commercial CFD software ANSYS 18.0. The three-dimensional compressible Reynolds-averaged Navier-Stokes (RANS) equations coupled with the SST k-ω turbulence model have been used to analyze the coaxial mixing characteristics of the jets. The numerical study is validated with the experimental data of the wall static pressures measured in the combustor’s flow direction. The pylon and wall injectors are located symmetrically at the gas generator’s exit nozzle, and the air is used as the injectant to simulate gaseous fuel. Three injection pressures are used for the study to understand the flow field characteristics in the injector regime. Also, the gas generator downstream direction is investigated. The shock waves generated from the gas generator nozzle enhance the mixing of the coaxial jets with minimum total pressure loss. The shock wave interactions are noticed with reducing intensity within the supersonic combustor for pylon injection, leading to higher total pressure loss than the wall injection. The pylon injection provides the spatial distribution of fuels compared to the wall injection in the coaxial supersonic flow field.

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 of Deposition Phenomena on High-Pressure Turbine Vane Using UPACS

2019 ◽  
Author(s):  
Kenta Mizutori ◽  
Koji Fukudome ◽  
Makoto Yamamoto ◽  
Masaya Suzuki
Keyword(s):  
Numerical Simulation ◽  
High Pressure ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
High Pressure Turbine ◽  
Pressure Loss Coefficient ◽  
Turbine Vane ◽  
Total Pressure Loss Coefficient

Abstract We performed numerical simulation to understand deposition phenomena on high-pressure turbine vane. Several deposition models were compared and the OSU model showed good adaptation to any flow field and material, so it was implemented on UPACS. After the implementation, the simulations of deposition phenomenon in several cases of the flow field were conducted. From the results, particles adhere on the leading edge and the trailing edge side of the pressure surface. Also, the calculation of the total pressure loss coefficient was conducted after computing the flow field after deposition. The total pressure loss coefficient increased after deposition and it was revealed that the deposition deteriorates aerodynamic performance.

Probabilistic and Numerical Modelling of a Lobed Mixer at Windmilling Conditions

2013 ◽  
Author(s):  
Maxime Lecoq ◽  
Nicholas Grech ◽  
Pavlos K. Zachos ◽  
Vassilios Pachidis
Keyword(s):  
Flow Field ◽  
Response Surface ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Total Pressure ◽  
Engine Performance ◽  
Operating Conditions ◽  
Total Pressure Loss ◽  
Operating Point ◽  
Mixing Losses

Aero-gas turbine engines with a mixed exhaust configuration offer significant benefits to the cycle efficiency relative to separate exhaust systems, such as increase in gross thrust and a reduction in fan pressure ratio required. A number of military and civil engines have a single mixed exhaust system designed to mix out the bypass and core streams. To reduce mixing losses, the two streams are designed to have similar total pressures. In design point whole engine performance solvers, a mixed exhaust is modelled using simple assumptions; momentum balance and a percentage total pressure loss. However at far off-design conditions such as windmilling and altitude relights, the bypass and core streams have very dissimilar total pressures and momentum, with the flow preferring to pass through the bypass duct, increasing drastically the bypass ratio. Mixing of highly dissimilar coaxial streams leads to complex turbulent flow fields for which the simple assumptions and models used in current performance solvers cease to be valid. The effect on simulation results is significant since the nozzle pressure affects critical aspects such as the fan operating point, and therefore the windmilling shaft speeds and air mass flow rates. This paper presents a numerical study on the performance of a lobed mixer under windmilling conditions. An analysis of the flow field is carried out at various total mixer pressure ratios, identifying the onset and nature of recirculation, the flow field characteristics, and the total pressure loss along the mixer as a function of the operating conditions. The data generated from the numerical simulations is used together with a probabilistic approach to generate a response surface in terms of the mass averaged percentage total pressure loss across the mixer, as a function of the engine operating point. This study offers an improved understanding on the complex flows that arise from mixing of highly dissimilar coaxial flows within an aero-gas turbine mixer environment. The total pressure response surface generated using this approach can be used as look-up data for the engine performance solver to include the effects of such turbulent mixing losses.

Experimental Investigation of Total Pressure Loss Development in a Highly Loaded Low-Pressure Turbine Cascade

2017 ◽  
Vol 140 (3) ◽  
Author(s):  
Philip Bear ◽  
Mitch Wolff ◽  
Andreas Gross ◽  
Christopher R. Marks ◽  
Rolf Sondergaard
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Low Pressure ◽  
Total Pressure Loss ◽  
Low Pressure Turbine ◽  
Pressure Transport ◽  
Secondary Flow Field ◽  
Highly Loaded

Improvements in turbine design methods have resulted in the development of blade profiles with both high lift and good Reynolds lapse characteristics. An increase in aerodynamic loading of blades in the low-pressure turbine (LPT) section of aircraft gas turbine engines has the potential to reduce engine weight or increase power extraction. Increased blade loading means larger pressure gradients and increased secondary losses near the endwall. Prior work has emphasized the importance of reducing these losses if highly loaded blades are to be utilized. The present study analyzes the secondary flow field of the front-loaded low-pressure turbine blade designated L2F with and without blade profile contouring at the junction of the blade and endwall. The current work explores the loss production mechanisms inside the LPT cascade. Stereoscopic particle image velocimetry (SPIV) data and total pressure loss data are used to describe the secondary flow field. The flow is analyzed in terms of total pressure loss, vorticity, Q-Criterion, turbulent kinetic energy, and turbulence production. The flow description is then expanded upon using an implicit large eddy simulation (ILES) of the flow field. The Reynolds-averaged Navier–Stokes (RANS) momentum equations contain terms with pressure derivatives. With some manipulation, these equations can be rearranged to form an equation for the change in total pressure along a streamline as a function of velocity only. After simplifying for the flow field in question, the equation can be interpreted as the total pressure transport along a streamline. A comparison of the total pressure transport calculated from the velocity components and the total pressure loss is presented and discussed. Peak values of total pressure transport overlap peak values of total pressure loss through and downstream of the passage suggesting that the total pressure transport is a useful tool for localizing and predicting loss origins and loss development using velocity data which can be obtained nonintrusively.

A Comprehensive Investigation of Blade Row Interaction Effects on Stator Loss Utilizing Vane Clocking

2017 ◽  
Author(s):  
Natalie R. Smith ◽  
Nicole L. Key
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Transition ◽  
Total Pressure Loss ◽  
Tip Leakage Flow ◽  
Inlet Flow ◽  
Blade Row ◽  
Vane Clocking

Blade row interactions drive the unsteady performance of high pressure compressors. Vane clocking is the relative circumferential positioning of consecutive stationary vane rows with the same vane count. By altering the upstream vane wake’s path with respect to the downstream vane, vane clocking changes the blade row interactions and results in a change in steady total pressure loss on the downstream vane. The open literature lacks a conclusive discussion of the flow physics governing these interactions in compressors. This paper presents the details of a comprehensive vane clocking study on the embedded stage of the Purdue 3-stage axial compressor. The steady loss results, including radial total pressure profiles and surface flow visualization, suggest a shift in the Stator 2 corner separations occurs between clocking configurations associated with the maximum and minimum total pressure loss. To better understand the flow mechanisms driving the vane clocking effects on the steady Stator 2 performance, time-resolved interrogations of the Stator 2 inlet flow field, surface pressure unsteadiness, and boundary layer response were conducted. The Stator 2 surface flows, both pressure unsteadiness and boundary layer transition, are influenced by vane clocking and interactions between Rotor 1 and Rotor 2, but neither of these results indicate a cause for the change in steady total pressure loss. Moreover, they are a result of upstream changes in the flow field: the interaction between the Stator 1 wake and Rotor 2 results in a circumferentially varying pattern which alters the inlet flow field for the downstream row, including the unsteadiness and frequency content in the tip and hub regions. Therefore, under different clocking configurations, Stator 2 experiences significantly different inlet blockage and unsteadiness from the Rotor 2 tip leakage flow and hub corner separation, which, in turn, shifts the radial blade loading distribution and subsequent loss development of Stator 2.

scholarly journals Constrained Shape Optimization of Airfoil Cascades Using a Navier-Stokes Solver and a Genetic/SQP Algorithm

1999 ◽  
Author(s):  
Brian H. Dennis ◽  
George S. Dulikravich ◽  
Zhen-Xue Han
Keyword(s):  
Flow Field ◽  
Shape Optimization ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Field Analysis ◽  
Navier Stokes ◽  
Conic Section ◽  
Aerodynamic Shape ◽  
Flow Field Analysis

The objective in this aerodynamic shape design effort is to minimize total pressure loss across the two-dimensional linear airfoil cascade row while satisfying a number of constraints. They included fixed axial chord, total torque, inlet and exit flow angles, and blade cross-section area, while maintaining thickness distribution greater than a minimum specified value. The aerodynamic shape optimization can be performed by using any available flow-field analysis code. For the analysis of the performance of intermediate cascade shapes we used an unstructured grid based compressible Navier-Stokes flow-field analysis code with k-e turbulence model. A robust genetic optimization algorithm was used for optimization and a constrained sequential quadratic programming was used enforcement of certain constraints. The airfoil geometry was parameterized using conic section parameters and B-splines thus keeping the number of geometric design variables to a minimum while achieving a high degree of geometric flexibility and robustness. Significant reductions of the total pressure loss were achieved using this constrained method for a supersonic exit flow axial turbine cascade.

The Influence of Bluff-Body Structural Parameters and Incoming Velocities on Advanced Vortex Combustor

2013 ◽  
Vol 860-863 ◽  
pp. 1383-1387
Author(s):  
Zhi Kai Wang ◽  
Zhuo Xiong Zeng ◽  
Yi Hua Xu
Keyword(s):  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Bluff Body ◽  
Structural Parameters ◽  
Total Pressure Loss ◽  
Structural Parameter ◽  
Stoichiometric Ratio ◽  
Fuel Type ◽  
Cold Flow

The influence of bluff-body structural parameters and incoming velocity on the cold flow field of advanced vortex combustor (AVC) has been studied, and the results show that the vortex in cavity can be the most stable and the total pressure loss can be the least when structural parameter is B2/B1=0.7, L/B1=0.8. The incoming velocity should not be too high or too low, it should be combined with specific structural parameters of the combustor, the combustion flame speed, fuel type, and the stoichiometric ratio and other factors. If AVC design is reasonable, the cavity can not only stabilize the flame, but also not be sensitive to the speed changes.

Experimental Study of the Flow and Pressure Drop Performances in Advanced Combustor Distributor Diffuser

2013 ◽  
Author(s):  
H. X. Liang ◽  
J. Q. Suo ◽  
M. Li
Keyword(s):  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
High Performance ◽  
Area Ratio ◽  
Total Pressure Loss ◽  
Pressure Recovery ◽  
Flow Rates ◽  
Turbine Engine ◽  
Outlet Flow

Gas turbine engine uses diffuser system to decelerate the compressor exit flow velocity before it enters combustor, it is important to design the compact structure and high performance of the diffuser for gas turbine engine. The diffuser and combustor dome configurations are critical flow path parameters in the design of a low-pressure-loss, high-performance combustion system. With rising of the inlet Mach number of the combustor, dramatically increasing of the diffuser total pressure loss and flow separation. So a new distributor diffuser was designed. In this paper preliminary results from an experimental investigation into the aerodynamic performance on a rectangle combustor-diffuser system with seven distributor plates were presented. Measurements were taken in the diffuser section to assess the diffuser performance characteristics under various conditions, the appropriate outlet flow field can be attained by changing the plate area ratio and form. Tests were carried out to investigate the influence of distributor diffuser plate geometry. During these measurements for each parametric configuration, data were obtained at 24 different flow rates through the distributor diffuser, it gave the conclusion that the distributor diffuser area ratio could be more than traditional diffusers with shorter construction and higher pressure recovery performance, while the flow loss through it was not beyond the traditional limit. Overall static pressure recovery improves and overall total pressure loss reduces with increasing distributor diffuser area ratio, and the increased flow rates through the distributor diffuser gave rise to a higher total pressure loss. The total pressure loss fraction was less than 2.5% when Mach number changed from 0.3 to 0.38; if the area ratio was more than 2.1, the diffuser loss coefficient remained less than 0.3, pressure recovery coefficient more than 0.5 and area ratio up to 2.45. There exists an area ratio in 1.6∼2.0 which makes diffuser outlet flow field distribution more uniform; Baffle structure can adjust the flow field distribution of outlet diffuser. As a result, the distributor diffuser can be potentially satisfied with demands for high performance combustor.

Influence of inlet swirl on pattern factor and pressure loss in an aero engine combustor

2021 ◽  
pp. 095440622110272
Author(s):  
KVL Narayana Rao ◽  
BVSSS Prasad ◽  
CH Kanna Babu ◽  
Girish K Degaonkar
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Linear Increase ◽  
Total Pressure Loss ◽  
Test Facility ◽  
Reacting Flow ◽  
Engine Test ◽  
Gas Temperature ◽  
Swirl Angle ◽  
Aero Engine

The effect of compressor exit swirl angle (θsw) at the intake of an aero engine combustor on the exit temperature non-uniformity (pattern factor) and combustor total pressure loss is investigated. Experiments are conducted in the engine test rig, measuring the gas temperature and pressure at the inlet and exit planes of the combustor. These parameters are measured at distinct locations along the circumferential and radial directions in the engine test facility. Simulations are carried out using RANS based turbulence modeling and reacting flow approach with Ansys CFX commercial code. The predicted results are validated with experimental data at 5° swirl angle. The swirl angle at the combustor intake is further varied from 0° to 15° and 4 cases (0°, 5°, 10° and 15°) has been considered to predict the effect on the combustor pattern factor and pressure loss. The changes in the flow structure inside the combustion chamber for all these 4 cases are reported in detail. The pattern factor varies from 0.34 to 0.49 as swirl angle changes from 0° to 15°. The lowest pattern factor of 0.34 occurs at 10° swirl angle. However a linear increase in combustor total pressure loss from 5.85% to 6.53% is predicted with the change in swirl angle from 0° to 15°.

Experimental Investigation of Total Pressure Loss Development in a Highly Loaded Low Pressure Turbine Cascade

2017 ◽  
Author(s):  
Philip Bear ◽  
Mitch Wolff ◽  
Andreas Gross ◽  
Christopher R. Marks ◽  
Rolf Sondergaard
Keyword(s):  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Low Pressure ◽  
Total Pressure Loss ◽  
Turbine Cascade ◽  
Low Pressure Turbine ◽  
Pressure Transport ◽  
Secondary Flow Field ◽  
Highly Loaded

Improvements in turbine design methods have resulted in the development of blade profiles with both high lift and good Reynolds lapse characteristics. An increase in aerodynamic loading of blades in the low pressure turbine section of aircraft gas turbine engines has the potential to reduce engine weight or increase power extraction. Increased blade loading means larger pressure gradients and increased secondary losses near the endwall. Prior work has emphasized the importance of reducing these losses if highly loaded blades are to be utilized. The present study analyzes the secondary flow field of the front-loaded low-pressure turbine blade designated L2F with and without blade profile contouring at the junction of the blade and endwall. The current work explores the loss production mechanisms inside the low pressure turbine cascade. Stereoscopic particle image velocimetry data and total pressure loss data are used to describe the secondary flow field. The flow is analyzed in terms of total pressure loss, vorticity, Q-Criterion, turbulent kinetic energy and turbulence production. The flow description is then expanded upon using an Implicit Large Eddy Simulation of the flow field. The RANS momentum equations contain terms with pressure derivatives. With some manipulation these equations can be rearranged to form an equation for the change in total pressure along a streamline as a function of velocity only. After simplifying for the flow field in question the equation can be interpreted as the total pressure transport along a streamline. A comparison of the total pressure transport calculated from the velocity components and the total pressure loss is presented and discussed. Peak values of total pressure transport overlap peak values of total pressure loss through and downstream of the passage suggesting that total pressure transport is a useful tool for localizing and predicting loss origins and loss development using velocity data which can be obtained non-intrusively.

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