scholarly journals Выбор геометрических характеристик фронтового устройства и длины камеры сгорания прямоточного типа

2021 ◽  
pp. 19-28
Author(s):  
Дмитрий Викторович Козел
Keyword(s):  
Combustion Chamber ◽  
Pressure Loss ◽  
Total Pressure ◽  
Combustion Efficiency ◽  
Total Pressure Loss ◽  
Stable Operation ◽  
Operating Characteristics ◽  
Combustion Chambers ◽  
Geometric Characteristics ◽  
Optimization Parameters

A method has been developed for selecting the geometric characteristics of the front and the length of the direct-flow combustion chamber. Afterburner combustion chambers are of the ramjet type and are used for a short-term increase in the thrust of a gas turbine engine during takeoff, for overcoming the sound barrier by an aircraft and for flying at supersonic speed, and for making maneuvers. As part of ramjet engines, ramjet combustion chambers are used as the main combustion chambers in which the process of fuel combustion and heat supply to the working fluid is ensured. The developed method for selecting the geometric characteristics consists in optimizing the main operating characteristics of the combustion chamber. Mathematical models are proposed for describing the dependence of the total pressure loss, the combustion efficiency and the range of stable operation of the combustion chamber against the parameters of the flow at the inlet to the combustion chamber and the geometric characteristics of the front device and the length of the combustion chamber. The analysis of the dependences of the combustion chamber working characteristics on the geometric characteristics of the front-line device and its length is carried out. As a result of the analysis of mathematical models, a list of the main geometric characteristics of the front device was determined, on which the total pressure loss, the combustion efficiency and the range of stable operation of the combustion chamber depend. Optimization parameters, optimization criterion and limits for solving the optimization problem are determined. As an implementation of the optimization method, it is proposed to use a diagram of the combustion chamber performance in the coordinates of the optimization parameters. The developed method makes it possible to ensure the optimal basic operating characteristics of the combustion chamber - total pressure loss, combustion efficiency and combustion stability limits.

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.

Effect of Casing Geometry on Flow Characteristics in a Model Can-Combustor

2012 ◽  
Author(s):  
Abdur Rahim ◽  
Dhirgham Alkhafagiy ◽  
Prabal Talukdar
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Numerical Study ◽  
Flow Characteristics ◽  
Total Pressure Loss ◽  
Stable Operation ◽  
Flow Uniformity ◽  
Flow Through ◽  
Combustor Liner

In a gas turbine combustor, it is necessary to use a diffuser to decelerate the high velocity air stream delivered by the compressor and thus avoid high total pressure loss. The interaction between the diffuser and combustor external flows plays a key role in controlling the pressure loss, air flow distribution around the combustor liner. Flow through casing-liner annulus is crucial as it feeds air to the primary, secondary and dilution holes. It is important that the annulus flow has sufficient static pressure to achieve adequate penetration of the jets. Moreover, the correct proportion of air enters the combustor liner through the dome and the various ports to maintain stable operation and good quality outlet condition. Length of combustor can be reduced if a provision is made for sufficient diffusion in the dump region. In the present numerical study, three can-combustor models of different geometry with a constant dump-gap have been analyzed with emphasis on the flow through annulus. A comparison has been made amongst the three models in terms of flow uniformity, static pressure recovery and total pressure loss. It is observed that flow uniformity in the annulus region is improved if a small divergence in length and a curved shape step height casing is made.

Combination of 1D Code and CFD for Performance Analysis of a Silo Type Gas Turbine Combustor

2010 ◽  
Author(s):  
N. Rasooli ◽  
S. Besharat Shafiei ◽  
H. Khaledi
Keyword(s):  
Combustion Chamber ◽  
Gas Turbine ◽  
Mass Distribution ◽  
Gas Turbines ◽  
Pressure Loss ◽  
Combustion Efficiency ◽  
Operating Conditions ◽  
Empirical Correlations ◽  
Combustion Chambers ◽  
Cfd Simulations

Whereas Gas Turbines are the most important producers of Propulsion and Power in the world and with attention to the importance of combustion chamber as one of the three basic components of Gas Turbine, various activities in different levels have been done on this component. Because of the environmental limitations and laws related to the pollutants such as NOx and CO, Lean Premixed Combustion Chambers are specially considered in gas turbine industries. This study is part of a Multi-Layer simulation of the whole gas turbine cycle in MPG Company. In this work, the combination of a general 1D code and CFD is used for deriving appropriate performance curves for a 1D and 0D gas turbine design, off-design and dynamic cycle code. This 1D code is a general code which has been developed for different combustion chambers; annular, can-annular, can type and silo type combustion chambers. The purpose of generating this 1D code is the possibility of fast analysis of combustors in different operating conditions and reaching required outputs. This 1D code is a part of a general simulation 1D code for gas turbine and was used for a silo type combustor performance prediction. This code generates required quantities such as pressure loss, exit temperature, liner temperature and mass distribution through the combustion chamber. Mass distribution and pressure loss are analyzed and determined with an electrical analogy. Results derived from 1D code are validated with empirical data available for different combustors. There is appropriate agreement between these experimental and analytical results. Drag coefficients for liner holes are available from experimental data and for burner are calculated as a curve with CFD simulations. What differs this code from other 1D codes for gas turbine combustors is the advantage of using combustion efficiencies evolved from numerical simulation results in different loads. These efficiencies are determined with CFD simulations and are available as maps and inserted into the gas temperature calculation algorithm of 1D code. In other 1D codes in this field, empirical correlations are used for combustion efficiency determination. Combustion efficiency curves for design and off-design conditions in this study are achieved by 2D and 3D simulation of combustion chamber with application of EBU/Finite Rate model and 8 step reactions of CH4 burning. Diffusion flame in low loads and premixed flame in high loads are considered. Flame stability and Lean Blow Out charts are evolved from CFD simulation and Heat transfer is applied with empirical correlations.

Effects of Inlet Parameters on Combustion Performance in Gas Turbine Combustor

2018 ◽  
Vol 35 (4) ◽  
pp. 339-350
Author(s):  
Yingwen Yan ◽  
Yunpeng Liu ◽  
Liang Huang ◽  
Jinghua Li
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Combustion Efficiency ◽  
Total Pressure Loss ◽  
Inlet Pressure ◽  
Inlet Temperature ◽  
Gas Turbine Combustor ◽  
Combustion Performance ◽  
Co Emission ◽  
No Emissions

Abstract The effects of different inlet parameters such as inlet temperature and pressure on combustion performance in a single-head combustor were experimentally investigated in this study. The combustion efficiency, total pressure loss, and CO and NO emissions at the outlet of a single-head rectangular combustor with different types of swirlers were separately measured. The experimental results showed that the inlet parameters had obvious effects on the combustion performance, with critical values of 600 K for the inlet temperature and 3.5 bar for the inlet pressure. The combustion efficiency noticeably increased with an increase in the inlet pressure or temperature below these values; however, when either of the inlet parameters was above the critical value, the combustion efficiency was approximately 100 %; that is, the combustion efficiency changed little with an increase in inlet temperate or pressure. When the inlet temperature or pressure increased, NO emission increased but CO emission decreased. By fitting curves to analyze the experimental data, the empirical relationships between the emissions and the inlet temperature were observed to be $CO\; \propto \;{e^{ - T}}, NO\; \propto \;{e^T}$, and those between the emissions and the inlet pressure were $CO\, \propto \,{e^{a + bP + c{P^2}}}, NO\, \propto \,{e^P}$. The total pressure loss increased with the inlet temperature.

Combustion Characteristics of Advanced Vortex Combustor Burning H2 Fuel

2014 ◽  
Author(s):  
Yangbo Deng ◽  
Xi Jiang ◽  
Fengmin Su
Keyword(s):  
Reaction Mechanism ◽  
Pressure Loss ◽  
Total Pressure ◽  
Equivalence Ratio ◽  
Combustion Efficiency ◽  
Combustion Characteristics ◽  
Main Flow ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Set Up

The combustion characteristics of the advanced vortex combustor (AVC) burning H2 fuel are studied numerically. First, using the 19-step reaction mechanism, the flame morphology of the pre-mixed H2/Air combustion under the different conditions, is computed. The calculation results are in agreement with experimental data from the literature. Second, a numerical model of a lean premixed H2/Air combustion is set up, based on the 19-step reaction mechanism. A numerical simulation is carried out to study the combustion characteristics of the AVC. The results show that the combustion can be steadily maintained, with the equivalence ratio of the H2/Air main flow kept at 0.45. At the same time, the total pressure loss coefficient is 2.77% and the combustion efficiency is 99.8%. The total pressure loss, vortex configuration and stability, combustion efficiency of the AVC are influenced by the equivalence ratio, total pressure and static pressure of main flow in the AVC.

Applications of little blades in a high-load compressor cascade with leaned blade

2021 ◽  
pp. 095440622110164
Author(s):  
Shan Ma ◽  
Xiaolin Sun
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Active Role ◽  
Low Energy ◽  
Total Pressure Loss ◽  
Stable Operation ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Lean Angle

The development of boundary layer affects the compressor cascade performance to a certain extent. Therefore, the compound lean and little blades are selected to redistribute the boundary layer, and the influences of these two flow control technologies on the axial compressor cascade performance are further studied. The calculated results showed that appropriate high pressure region on the blade suction surface near the end-wall is helpful to reduce the total pressure loss of compressor cascade, which can be achieved by positive lean technique. Meanwhile, the maximum stable operation boundary can be expanded by the application of positive leaned blade. On the other hand, the introduction of negative lean angle not only increases the total pressure loss of cascade, but reduces the stable operation range. As the little blades are introduced in the negative lean compressor cascade, the stable operation range is significantly improved by the introduction of little blades. Especially the cascade with −10° lean angle, the maximum stable operation boundary is increased from 1° to 6°. In the positive lean compressor cascade, although more low-energy fluid is accumulated on the blade suction surface near the mid-span, the little blades still show an active role in reducing the total pressure loss and expending the stable operation range, because the influence range of induced vortex reaches 30%span. The results provide a reference for improving the aerodynamic performance of compressor stator, especially when more low-energy fluid is blocked in the range near the mid-span.

Combustion efficiency and pressure loss balance for the supersonic combustor

2019 ◽  
Vol 234 (6) ◽  
pp. 1149-1156 ◽  
Author(s):  
AM Tahsini
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Cross Flow ◽  
Combustion Efficiency ◽  
Total Pressure Loss ◽  
Flight Performance ◽  
Compression Process ◽  
Design Point ◽  
Minimum Drag ◽  
Inlet Conditions

The purpose of this paper is to investigate the effects of intake’s compression process of the scramjet on its flight performance. The hydrogen injection to the supersonic cross-flow is considered as the problem configuration. The finite volume solver is developed to simulate the compressible reacting turbulent flow using the proper reaction mechanism as the finite rate chemistry. The combustion efficiency and the drag force are the most important parameters on the scramjet flight performance, and finding the design point to balance the higher combustion efficiency and the lower minimum drag, which depends on the total pressure loss, can be used to optimize the supersonic combustors. The performance of the supersonic intake is considered here using some oblique shock waves with equal flow-deflection angles to compute the combustor’s inlet condition. The variation of combustion efficiency and total pressure loss is presented for different combustor’s inlet conditions. The results are presented for the constant jet to inlet pressure ratios and also for the constant equivalence ratios, in which the last one is much appropriate and utilized to find the optimum design point of the intake and the combustor, for assumed flight condition.

Optimization study on the influence of little blades’ spatial position on a compressor cascade performance

2021 ◽  
pp. 095441002110443
Author(s):  
Shan Ma ◽  
Xiaolin Sun
Keyword(s):  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Spatial Position ◽  
Total Pressure Loss ◽  
Field Analysis ◽  
Stable Operation ◽  
Compressor Cascade ◽  
Optimal Position ◽  
End Wall

To reveal the importance of little blades’ spatial position to improve the cascade performance at different condition, the pitchwise and axial direction of the little blades on the end-wall are adopted as the optimization variables to complete a double-objective optimization. Meanwhile, the three-dimensional flow field characteristics of the cascade with and without little blades are analyzed comparatively. The study found that as the optimal solutions are obtained at the three bigger incidences (3°, 5°, and 7°), the optimal position is always close to the leading edge of blade and far away from the blade suction surface, and the more intuitive design suggestions are given in this article. Moreover, at the near design conditions (−1°, 0°, and 1°), little blades increase the total pressure loss and reduce the static pressure, which are considered unsuitable for improving the cascade performance. If the stable operation range are the main performance indicators, the optimization of the little blades’ spatial position should be completed at the near stall condition (7° incidence). If the conditions with mid-range incidences (2°< i <5°) are the main performance index, the parameter optimization of little blades should be achieved at 5°. Based on the further flow field analysis of the optimization results obtained at 3°, 5°, and 7° incidences (named Opt_Act3, Opt_Act5, and Opt_Act7), the induced vortices resist the effect of axial reverse pressure gradient and pass through the blade passage, which is the main reason for the total pressure loss reduction. Appropriate spatial position of little blades not only strengthens the capability to prevent the low-energy fluids accumulating in the corner region near the end-wall, but exhibits sufficient advantage to weaken the boundary layer.

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.

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