NUMERICAL CALCULATIONS OF GASEOUS REACTING FLOWS IN A MODEL OF GAS TURBINE COMBUSTORS

This paper describes the numerical calculations of gaseous reaction flows in a model of gas turbine combustors. The profiles of hydrodynamic and thermodynamic patterns in a three-dimensional combustor model are obtained by solving the governing differential transport equations. The well-established numerical prediction algorithm SIMPLE, the modified k-ε turbulence model and k-ε-g turbulent diffusion flame model have been adopted in computations. The β function has been selected as probability density function. The effect of combustion process on flow patterns has been investigated. The calculated results have been verified by experiments. They are in remarkably good agreement.

Effects of Lobe Peak-to-Trough Width Ratio on Mixing and Combustion Performance in ATR Combustors

The air-turbo-rocket(ATR)engine is a promising propulsion plant for achieving numerous surface and air launched missile missions.The application of lobed mixer in the ATR combustor can promote the mixing of the fuelrich gas and the air,thus improving the engine performance significantly.The numerical simulation method was conducted to explore the effects of lobe peak-to-trough width ratio on mixing and combustion performance in ATR combustors.Results show that:For a given peak lobe width b1,the combustion efficiency and total pressure loss decrease with the increase of trough lobe width b2;For a given b2,the combustion efficiency and total pressure loss decrease with the increase of b1;The fan-type lobed mixer with smaller b2has a better effect on promoting the combustion efficiency in the region near the ATR combustor center line than that with a pair of parallel side walls.The total pressure recovery coefficient reaches more than 0.99 at the exit of combustor in nonreactive combustion while the total pressure loss reaches more than 4%in the reacting combustion.Compared with the mixing process,more than80%of the total pressure loss is caused during combustion.

An analytical theory of heated duct flows in supersonic combustors

One-dimensional analytical theory is developed for supersonic duct flow with variation of cross section, wall friction, heat addition, and relations between the inlet and outlet flow parameters are obtained. By introducing a self- similar parameter, effects of heat releasing, wall friction, and change in cross section area on the flow can be normalized and a self-similar solution of the flow equations can be found. Based on the result of self-similar solution, the sufficient and necessary condition for the occurrence of thermal choking is derived. A re- lation of the maximum heat addition leading to thermal choking of the duct flow is derived as functions of area ratio, wall friction, and mass addition, which is an extension of the classic Rayleigh flow theory, where the effects of wall friction and mass addition are not considered. The present work is expected to provide fundamentals for developing an integral analytical theory for ramjets and scramjets.

An FV-EE model to predict lean blowout limits for gas turbine combustors with different structures and sprays

The occurrence of Lean Blowout(LBO)is a disadvantage that endangers a stable operation of gas turbines.A determination of LBO limits is essential in the design of gas turbine combustors.A semiempirical model is one of the most widely used methods to predict LBO limits.Among the existing semiempirical models for predicting LBO limits,Lefebvre’s LBO model and the Flame Volume(FV)model are particularly suitable for gas turbine combustors.On the basis of Lefebvre’s and FV models,the concept of effective evaporation efficiency is introduced in this paper,and a Flame Volume-Evaporation Efficiency(FV-EE)model is derived and validated.LBO experiments are carried out in a model combustor with 23 different structures and 10 different sprays.The prediction uncertainty of the FV-EE model is less than±13%for all of these 33 structures and sprays,compared with±50%for the FV model and±60%for Lefebvre’s model.Furthermore,the prediction uncertainty of the FV-EE model is also less than±13%for other combustors from available literature.

Hybrid method based on flame volume concept for lean blowout limits prediction of aero engine combustors

The Lean Blowout(LBO)limit is crucial for the aircraft engines.The semi-empirical(such as Lefebvre’s LBO model and Flame Volume(FV)model),numerical and hybrid methods are widely utilized for the LBO limit quick prediction.An innovative hybrid method based on the FV concept is proposed.This method can be classified as a semi-empirical/physical based hybrid prediction method.In this hybrid method,it is assumed that the flame volume varies nearly linearly with the fuel/air ratio near the LBO.The flame volume is obtained directly by the numerical simulation using the threshold value of the visible flame boundary as 900 K.Then the final LBO limits is determined by the FV model.On the premise of keeping the good generality of prediction,the hybrid method based on the FV concept can further improve the prediction accuracy.The comparison with the prediction of the existing available methods on fifteen combustors shows that the hybrid method based on the FV concept achieves better prediction accuracy.The prediction uncertainties between the experimental results and the predicted values by the hybrid method based on the FV concept are within about±10%.

CFD predictions of LBO limits for aero-engine combustors using fuel iterative approximation

Lean blow-out （LBO） is critical to operational performance of combustion systems in propulsion and power generation. Current predictive tools for LBO limits are based on decadesold empirical correlations that have limited applicability for modern combustor designs. According to the Lefebvre＇s model for LBO and classical perfect stirred reactor （PSR） concept, a load parameter （LP） is proposed for LBO analysis of aero-engine combustors in this paper. The parameters contained in load parameter are all estimated from the non-reacting flow field of a combustor that is obtained by numerical simulation. Additionally, based on the load parameter, a method of fuel iterative approximation （FIA） is proposed to predict the LBO limit of the combustor. Compared with experimental data for 19 combustors, it is found that load parameter can represent the actual combustion load of the combustor near LBO and have good relativity with LBO fuel/air ratio （FAR）. The LBO FAR obtained by FIA shows good agreement with experimental data, the maximum prediction uncertainty of FIA is about ±17.5%. Because only the non-reacting flow is simulated, the time cost of the LBO limit prediction using FIA is relatively low （about 6 h for one combustor with computer equipment of CPU 2.66 GHz · 4 and 4 GB memory）, showing that FIA is reliable and efficient to be used for practical applications.

An Integrated Approach for Optimal Design of Micro Gas Turbine Combustors

The present work presents an approach for the optimized design of small gas turbine combustors, that integrates a0-D code, CFD analyses and an advanced game theory multi-objective optimization algorithm. The output of the0-D code is a baseline design of the combustor, given the required fuel characteristics, the basic geometry (tubularor annular) and the combustion concept (i.e. lean premixed primary zone or diffusive processes). For the optimizationof the baseline design a simplified parametric CAD/mesher model is then defined and submitted to a CFDcode. Free parameters of the optimization process are position and size of the liner hole arrays, their total area andthe shape of the exit duct, while different objectives are the minimization of NOx emissions, pressure losses andcombustor exit Pattern Factor. A 3D simulation of the optimized geometry completes the design procedure. As afirst demonstrative example, the integrated design process was applied to a tubular combustion chamber with alean premixed primary zone for a recuperative methane-fuelled small gas turbine of the 100 kW class.

Numerical study on ethylene-air continuous rotating detonation in annular combustors with different widths

To investigate the impact of combustor width on continuous rotating detonation(CRD)fueled by ethylene and air,a series of 3 D simulations are conducted by changing the inner cylinder radius of an annular combustor while retaining the same outer cylinder radius.The results show that the CRD wave propagates more steadily and faster as the combustor width increases.The high-temperature zone at the backward-facing step preheats the propellants and contributes to the steady propagation of the CRD wave in 25-and 30-mm wide combustors.The highest and the lowest velocities are obtained in the30-and 15-mm wide combustors at,respectively,1880.27 and 1681.01 m/s.On the other hand,the average thrust decreases as the combustor width increases.The highest thrust is obtained in the 15-mm wide combustor while the lowest is in the 30-mm wide combustor,at 758.06 and 525.93 N,respectively.Nevertheless,the thrust is much more stable in the 25-and 30-mm wide combustors than in the 15-and 20-mm wide combustors.

PIV MEASUREMENT FOR SWIRLER FLOW FIELD IN GAS TURBINE COMBUSTOR

The characteristics of swirler flow field, including cold flow field and combustion flow field, in gas tur- bine combustor with two-stage swirler are studied by using particle image velocimetry （PIV）. Velocity compo- nents, fluctuation velocity, Reynolds stress and recirculation zone length are obtained, respectively. Influences of geometric parameter of primary hole, arrangement of primary hole, inlet air temperature, first-stage swirler an- gle and fuel/air ratio on flow field are investigated, respectively. The experimental results reveal that the primary recirculation zone lengths of combustion flow field are shorter than those of cold flow field, and the primary reeir- culation zone lengths decrease with the increase of inlet air temperature and fuel/air ratio. The change of the geo- metric parameter of primary hole casts an important influence on the swirler flow field in two-stage swirler com- bustor.

Reconstruction of Turbulent Swirling Flow in a Dump Combustor

Experimental data of the continuous evolution of fluid flow characteristics in a dump combustor is very useful and essential for better and optimum designs of gas turbine combustors and ramjet engines. Unfortunately, experimental techniques such as 2D and/or 3D LDV （Laser Doppler Velocimetry） measurements provide only limited discrete information at given points; especially, for the cases of complex flows such as dump combustor swirling flows. For this type of flows, usual numerical interpolating schemes appear to be unsuitable. Recently, neural networks have emerged as viable means of expanding a finite data set of experimental measurements to enhance better understanding of a particular complex phenomenon. This study showed that generalized feed forward network is suitable for the prediction of turbulent swirling flow characteristics in a model dump combustor. These techniques are proposed for optimum designs of dump combustors and ramjet engines.

Numerical Analysis of Pollutant Formation in a Liquid Spray Sudden Expansion Combustor

A SUE （sudden expansion） combustor is analyzed using CFD （computational fluid dynamics） methods. Corresponding CO and NOx emissions are computed for various operating conditions of the SUE combustor with a can type geometrical configuration. The goal of this work is to see if the SUE combustor is a viable alternative to conventional combustors which utilize swirlers. It is found the can type combustor＇s NOx emissions are quite low compared to other combustor types but the CO emissions are fairly high. Emissions can be improved by providing better mixing of the fuel and oxidizer in the primary combustion zone. The SUE combustor design needs to be further refined in order for it to be a viable alternative to conventional combustors with swirters.

Experimental Investigation of Hydrocarbon-fuel Ignition in Scramjet Combustor

The direct-connected supersonic combustor experiment is finished for kerosene fuel ignition in H_2/O_2 preheated impulse facility. The entrance parameter of combustor corresponds to scramjet flight Mach number 3.5. Kerosene ignition is realized by using hydrogen as pilot flame. Wall pressure distributions of combustion are measured and flame photographs of ultraviolet ray are got. Experiment indicates that it is very difficult for kerosene fuel to realize self-ignition at low entrance temperature (below 900K) in supersonic combustor. Hydrogen pilot flame is one of the efficient methods for realizing kerosene ignition.

Experimental Study on Effects of Fuel Injection on Scramjet Combustor Performance

In order to investigate the effects of fuel injection distribution on the scrarnjet combustor performance, there are conducted three sets of test on a hydrocarbon fueled direct-connect scramjet test facility. The results of Test A, whose fuel injection is carried out with injectors located on the top-wall and the bottom-wall, show that the fuel injection with an appropriate close-front and centralized distribution would be of much help to optimize combustor performances. The results of Test B, whose fuel injection is performed at the optimal injection locations found in Test A, with a given equivalence ratio and different injection proportions for each injector, show that this injection mode is of little benefit to improve combustor performances. The results of Test C with a circumferential fuel injection distribution displaies the possibility of ameliorating combustor performance. By analyzing the effects of injection location parameters on combustor performances on the base of the data of Test C, it is clear that the injector location has strong coupled influences on combustor performances. In addition, an irmer-force synthesis specific impulse is used to reduce the errors caused by the disturbance of fuel supply and working state of air heater while assessing combustor performances.

Influence of propagation direction on operation performance of rotating detonation combustor with turbine guide vane

Due to the pressure gain combustion characteristics,the rotating detonation combustor(RDC)can enhance thermodynamic cycle efficiency.Therefore,the performance of gas-turbine engine can be further improved with this combustion technology.In the present study,the RDC operation performance with a turbine guide vane(TGV)is experimentally investigated.Hydrogen and air are used as propellants while hydrogen and air mass flow rate are about 16.1 g/s and 500 g/s and the equivalence ratio is about 1.0.A pre-detonator is used to ignite the mixture.High-frequency dynamic pressure transducers and silicon pressure sensors are employed to measure pressure oscillations and static pressure in the combustion chamber.The experimental results show that the steady propagation of rotating detonation wave(RDW)is observed in the combustion chamber and the mean propagation velocity is above 1650 m/s,reaching over 84%of theoretical Chapman-Jouguet detonation velocity.Clockwise and counterclockwise propagation directions of RDW are obtained.For clockwise propagation direction,the static pressure is about 15%higher in the combustor compared with counterclockwise propagation direction,but the RDW dominant frequency is lower.When the oblique shock wave propagates across the TGV,the pressure oscillations reduces significantly.In addition,as the detonation products flow through the TGV,the static pressure drops up to 32%and 43%for clockwise and counterclockwise propagation process respectively.

THREE—DIMENSIONAL GAS TURBINE COMBUSTOR PERFORMANCE MODELING

Numerical analysis of three-dimensional(3-D)two-phase reacting flowfield in an annular combustor wity the dump diffuser is developed in arbitrary curvilinear coordi-nates.Combustor performances are estimated by the em-pirical-analytical desing method.Ths influence of three inlet velocity profiles of the prediffuser and two operating conditions on combustor preformance and flow character-istic is predicted.

Numerical study of convective heat transfer of a supersonic combustor with varied inlet flow conditions

Characteristics of convective heat transfer of a supersonic model combustor with variable inlet flow conditions were studied by numerical simulation in this paper.The three-dimensional flow and wall heat flux at different air inlet Mach numbers of 2.2,2.8 and 3.2 were studied numerically with Reynolds-averaged Navier-Stokes equations with a shear-stress transport(SST)k-ωturbulence model and a three-step reaction model.Meanwhile,ethylene was chosen as the fuel,and the fixed fuel-to-air equivalence ratio is 0.8 in all cases in this paper.The results of the simulations indicate that wall heat flux distribution of the combustor is very non-uniform with several peaks of wall heat flux at varied locations.For the low inlet Mach number of 2.2,a shock train structure is formed in the isolator,and three peaks of wall heat flux are located respectively on the backward face of the cavity,on the side wall near the fuel injection and on the bottom wall near the injection holes,and a maximum wall heat flux reaches 5.4 MW/m2.For the medium inlet Mach number of 2.8,there exists a much shorter shock structure with three peaks of wall heat flux similar to that of Mach number 2.2.However,as the inlet Mach number increased to 3.2,there is no shock structure upstream of fuel injections,and the combustor flow is in a supersonic mode with different locations and values of wall heat flux peaks.The statistical results of wall heat loading show that the change of total wall heat is not monotonic with the increase of inlet Mach number,and the maximum appears in the case of Mach number being 2.8.Meanwhile,for all the cases,the bottom wall takes up more than 50%of the total heat loading.

Numerical Analysis on Multi-Field Characteristics and Synergy in a Large-Size Annular Combustion Chamber with Double Swirlers

In order to comprehensively evaluate the flow and heat transfer performance of a large-size annular combustion chamber of a heavy-duty gas turbine,we carried out numerical computation and analyses on the velocity,temperature and pressure fields in the chamber with double swirlers.The mathematical model of the coupling combustion,gas flow,and heat transfer process was established.The influences of the inlet swirling strength,fuel-air ratio and temperature of the premixed gas on the multi-field characteristics and synergy were investigated on the basis of field synergy theory.The results showed that the central recirculation zone induced by the inlet swirling flow grows downstream in the combustion chamber.The velocity and temperature in the outlet section of the chamber tend to be uniform due to the upstream improved synergy.The outer swirl number of the premixed gas flow has a great influence on the comprehensive flow and heat transfer performance of the combustion chamber.The synergy angles change towards benefiting the synergy between velocity and temperature fields with the increasing swirl numbers and inlet gas temperature while the velocity-pressure synergy becomes poor.The increasing fuel-air ratio of premixed gas leads to different trends of the velocity-temperature synergy and velocity-pressure synergy.The comprehensive synergy representing the low-resistance heat transfer performance is evidently dominated mainly by the velocity-temperature synergy.

Studies of a combined way of flame stability in ramjet combustor

numerical simulation was conducted to study the influence of bleeding. The Euler-Lagrange method was used to investigate the two-phase turbulent combustion flow. Standard k-ε turbulent model was adopted in the continuous phase simulation and particle-trajectory model was adopted in the dispersed phase simulation. The results demonstrates: air bleeding can improve the flow field after the strut and the stability of trapped vortex in the cavity; change of bleeding temperature has little effect on the total pressure recovery coefficient and significant effect on combustion efficiency; When fuel-air ratio changes, the combustor performs better in a lean oil state.

OPTIMIZATION OF A PULSE COMBUSTOR FOR A DISPERSE DRYING SYSTEM

The paper summarises the development and optimisation of a valveless pulse combus-tor designed as a heating source for drying in disperse systems. The basic research dealt with changing of the pulse combustor geometry, e. g. the volume of a combustion chamber and tail pipe length to obtain smooth trajectory of pressure oscillations and low emission of toxic substances. The pulse combustion mechanism and existing designs of pulse combustors applied in drying have been presented and analysed. Examples of specific applications of this technique are delivered.

Numerical investigations on effects of bluff body in flat plate micro thermo photovoltaic combustor with sudden expansion

In order to reveal combustion characteristics of H_2/air mixture in a micro-combustor with and without bluff body, the effects of inlet velocities, equivalence ratios and bluff body's blockage ratios on the temperature field, pressure of the combustor wall, combustion efficiency and blow-off limit were investigated. The numerical results indicate that the sudden expansion plate micro combustor with bluff body could enhance the turbulent disturbance of the mixed gas in the combustion chamber and the combustion condition is improved. Moreover, a low-speed and high temperature recirculation region was formed between the sudden expansion step and the bluff body so that the high and uniform wall temperature(>1000 K) could be gotten. As a result, it could strengthen the mixing process, prolong the residence time of gas, control the flame position effectively and widen the operation range by the synergistic effect of the bluff body and steps. When the blockage ratio ranged from 0.3 to 0.6, it could be found that the bluff body could play a stabilizing effect and expand combustion blow burning limit, and combustion efficiency firstly was increased with the inlet velocity and equivalence ratio, and then was decreased.