共查询到20条相似文献,搜索用时 875 毫秒
1.
2.
The flame quenching process in combustors was observed by high speed camera and Schlieren system, at the inflow conditions of Ma = 2.64, T 0 = 1 483 K, P 0 = 1.65 MPa, T = 724 K and P = 76.3 kPa. Changing process of the flame and shock structure in the combustor was clearly observed. The results revealed that the precombustion shock disappeared accompanied with the process in which the flame was blown out and withdrawed from the mainflow into the cavity and vanished after a short while. The time of quenching process was extended by the cavity flame holder, and the ability of flame holding was enhanced by arranging more cavities in the downstream as well. The flame was blown from the upstream to the downstream, so the flame in the downstream of the cavity was quenched out later than that in the upstream. 相似文献
3.
Jong-Ryul KimGyung-Min Choi Duck-Jool Kim 《Experimental Thermal and Fluid Science》2011,35(1):165-171
The influence of varying combustor pressure on flame oscillation and emission characteristics in the partially premixed turbulent flame were investigated. In order to investigate combustion characteristics in the partially premixed turbulent flame, the combustor pressure was controlled in the range of −30 to 30 kPa for each equivalence ratio (Φ = 0.8-1.2). The r.m.s. of the pressure fluctuations increased with decreasing combustor pressure for the lean condition. The combustor pressure had a sizeable influence on combustion oscillation, whose dominant frequency varied with the combustor pressure. Combustion instabilities could be controlled by increasing the turbulent intensity of the unburned mixture under the lean condition. An unstable flame was caused by incomplete combustion; hence, EICO greatly increased. Furthermore, EINOx simply reduced with decreasing combustor pressure at a rate of 0.035 g/10 kPa. The possibility of combustion control on the combusting mode and exhaust gas emission was demonstrated. 相似文献
4.
Mikaël Orain Frdric Grisch Eric Jourdanneau Bjorn Rossow Christian Guin Brigitte Trtout 《Comptes Rendus Mecanique》2009,337(6-7):373-384
Simultaneous measurements of PLIF-kerosene and PLIF-OH have been successfully performed in a multipoint injection system for various overall equivalence ratio, air inlet temperature between 480 and 730 K and pressure up to 2.2 MPa. Single shot 2D-maps of the spatial distribution of kerosene vapour and OH radical in the combustor have been recorded with good signal-to-noise ratio. Results show that depending on the split between the pilot and the main injectors, the flame front exhibits a single or a double structure. Good spatial correlation between the repartition of the kerosene vapour and the position of the flame front was observed; in particular, no “dark zone” is observed between the fuel and the flame front. As temperature and pressure increase, fuel evaporation improves and the spatial distribution of OH radical becomes more homogeneous in the combustor, suggesting a partially-distributed combustion. To cite this article: M. Orain et al., C. R. Mecanique 337 (2009). 相似文献
5.
Within the framework of the ideal, i.e., inviscid and non-heat conducting, gas model we consider the problem of designing
the supersonic section of a two-dimensional or axisymmetric nozzle realizing a uniform supersonic flow limitingly similar
with a sonic flow when the choked flow involves a curvilinear sonic line. Emphasis is placed on nozzles with abruptly or steeply
converging subsonic sections and a strongly curved sonic line formed by the C
−-characteristics of the expansion fan with the focus at the lower bend point of the vertical section of the subsonic contour.
In the two-dimensional case, the least possible greater-than-unity Mach number M
em at the nozzle exit corresponds to the flow in which the first intersection of the C
+-characteristics originated at the closing C
−-characteristic of the expansion fan falls on the unknown contour of its supersonic part. For a uniform flow with M
e
< M
em the intersection of C
+-characteristics beneath the unknown contour make impossible its construction. A part of the contour realizing a uniform flow
with M
em > 1 ensures a limitingly rapid flow acceleration and forms the initial region of the supersonic generator of a maximum-thrust
nozzle. For this reason, in the case of a curvilinear sonic line the supersonic generators of these nozzles have two, rather
than one, bends, which, however, is interesting only for the theory. At least, in the calculated examples the thrusts of the
nozzles with one and two bends differ only by a hundredth or even thousandth fractions of per cent. 相似文献
6.
We analyze the propagation of nonlinear waves in homogenized periodic nonlinear hexagonal networks, considering successively 1D and 2D situations. Wave analysis is performed on the basis of the construction of the effective strain energy density of periodic hexagonal lattices in the nonlinear regime. The obtained second order gradient nonlinear continuum has two propagation modes: an evanescent subsonic mode that disappears after a certain wavenumber and a supersonic mode characterized by an increase of the frequency with the wavenumber. For a weak nonlinearity, a supersonic mode occurs and the dispersion curves lie above the linear dispersion curve (vp =vp0). For a higher nonlinearity, the wave changes from a supersonic to an evanescent subsonic mode at s=0.7 and the dispersion curves drops below the linear case and vanish for certain values of the wavenumber. An important decrease in the frequency occurs for both subsonic and supersonic modes when the lattice becomes auxetic, and the longitudinal and shear modes become very close to each other. The influence of the lattice geometrical parameters of the lattice on the dispersion relations is analyzed. 相似文献
7.
Experimental Characterization of Premixed Flame Instabilities of a Model Gas Turbine Burner 总被引:1,自引:0,他引:1
Kai-Uwe Schildmacher Rainer Koch Hans-Jörg Bauer 《Flow, Turbulence and Combustion》2006,76(2):177-197
In recent years, the NO
x
emissions of heavy duty gas turbine burners have been significantly reduced by introducing premixed combustion. These highly premixed burners are known to be prone to combustion oscillations. In this paper, investigations of a single model gas turbine burner are reported focusing on thermo-acoustic instabilities and their interaction with the periodic fluctuations of the velocity and pressure. Phase-locked optical measurement techniques such as LDA and LIF gave insight into the mechanisms.Detailed investigations of a gas turbine combustor rig revealed that the combustor as well as the air plenum oscillate in Helmholtz modes. These instabilities could be attributed to the phase lag of the pressure oscillations between the air plenum and the combustor, which causes an acceleration and deceleration of the air flow through the burner and, therefore, alternating patterns of fuel rich and lean bubbles. When these bubbles reach the reaction zone, density fluctuations are generated which in turn lead to velocity fluctuations and, hence, keep up the pressure oscillations.With increasing the equivalence ratio strong combustion oscillations could be identified at the same frequency. Similarly as with weak oscillations, Helmholtz mode pressure fluctuations are present but the resulting velocity fluctuations in the combustor can be described as a pumping motion of the flow. By the velocity fluctuations the swirl stabilization of the flame is disturbed. At the same time, the oscillating pressure inside the combustor reaches its minimum value. Shortly after the flame expands again, the pressure increases inside the combustor. This phenomenon which is triggered by the pressure oscillations inside the air plenum seems to be the basic mechanism of the flame instability and leads to a significant increase of the pressure amplitudes. 相似文献
8.
We establish the existence and stability of multidimensional steady transonic flows with transonic shocks through an infinite
nozzle of arbitrary cross-sections, including a slowly varying de Laval nozzle. The transonic flow is governed by the inviscid
potential flow equation with supersonic upstream flow at the entrance, uniform subsonic downstream flow at the exit at infinity,
and the slip boundary condition on the nozzle boundary. Our results indicate that, if the supersonic upstream flow at the
entrance is sufficiently close to a uniform flow, there exists a solution that consists of a C
1,α subsonic flow in the unbounded downstream region, converging to a uniform velocity state at infinity, and a C
1,α multidimensional transonic shock separating the subsonic flow from the supersonic upstream flow; the uniform velocity state
at the exit at infinity in the downstream direction is uniquely determined by the supersonic upstream flow; and the shock
is orthogonal to the nozzle boundary at every point of their intersection. In order to construct such a transonic flow, we
reformulate the multidimensional transonic nozzle problem into a free boundary problem for the subsonic phase, in which the
equation is elliptic and the free boundary is a transonic shock. The free boundary conditions are determined by the Rankine–Hugoniot
conditions along the shock. We further develop a nonlinear iteration approach and employ its advantages to deal with such
a free boundary problem in the unbounded domain. We also prove that the transonic flow with a transonic shock is unique and
stable with respect to the nozzle boundary and the smooth supersonic upstream flow at the entrance. 相似文献
9.
Two-fluid model and divisional computation techniques were used. The multi-species gas fully N-S equations were solved by upwind TVD scheme. Liquid phase equations were solved by NND scheme. The phases-interaction ODE equations were solved by 2nd Runge-Kutta approach. The favorable agreement is obtained between computational results and PLIF experimental results of iodized air injected into a supersonic flow. Then, the numerical studies were carried out on the mixing of CH
4
and kerosene injected into a supersonic flow with H
2
pilot injection. The results indicate that the penetration of kerosene approaches maximum when it is injected from the second injector. But the kerosene is less diffused compared with the gas fuels. The free droplet region appears in the flow field. The mixing mechanism of CH
4
with H
2
pilot injection is different from that of kerosene. In the staged duct, H
2
can be entrained into both recirculation zones produced by the step and injectors. But CH
4
can only be carried into the recirculation between the injectors. Therefore, initiations of H
2
and CH
4
can occur in those regions. The staged duct is better in enhancing mixing and initiation with H
2
pilot flame. 相似文献
10.
Svetoslav Marinov Matthias Kern Nikolaos Zarzalis Peter Habisreuther Antonio Peschiulli Fabio Turrini O. Nuri Sara 《Flow, Turbulence and Combustion》2012,89(1):73-95
One of the most promising methods for reducing NO x emissions of jet engines is the lean combustion process. For realization of this concept the percentage of air flowing through the combustor dome has to be drastically increased, which implies high volume fluxes in the primary zone of the combustion chamber and represents a substantial challenge in regard to the flame stabilization. Swirl motion is thus applied to the air flux by the swirl generator and decisively contributes to the flame stabilization. The current paper reviews an atmospheric investigation of a burner configuration in regard to the weak extinction limit, comprising a confined non-premixed swirl-stabilized flame. The burner can be supplied with either kerosene or after a small adaption with natural gas (methane). Therefore, a comparison of a kerosene-fuelled flame (spray flame) to a natural gas fuelled one (methane flame) can be performed. Both are realized by almost identical burner configuration and at identical conditions. The main idea of this work is to align the stability characteristics of both flames by means of similarity. However, fundamental differences regarding the flame structures of the flames are detected through in-flame measurements. This determines the limits of the current approach and motivates an appropriate choice of flame modeling. 相似文献
11.
Experimental investigation of ram accelerator propulsion modes 总被引:1,自引:0,他引:1
Experimental investigations on the propulsive modes of the ram accelerator are reviewed in this paper. The ram accelerator is a ramjet-in-tube projectile accelerator whose principle of operation is similar to that of a supersonic air-breathing ramjet. The projectile resembles the centerbody of a ramjet and travels through a stationary tube filled with a premixed gaseous fuel and oxidizer mixture. The combustion process travels with the projectile, generating a pressure distribution which produces forward thrust on the projectile. Several modes of ram accelerator operation are possible which are distinguished by their operating velocity range and the manner in which the combustion process is initiated and stabilized. Propulsive cycles utilizing subsonic, thermally choked combustion theoretically allow projectiles to be accelerated to the Chapman-Jouguet(C-J) detonation speed of a gaseous propellant mixture. In the superdetonative velocity range, the projectile is accelerated while always traveling faster than the C-J speed, and in the transdetonative regime (85–115 % of C-J speed) the projectile makes a smooth transition from a subdetonative to a superdetonative propulsive mode. This paper examines operation in these three regimes of flow using methane and ethylene based propellant mixtures in a 16 m long, 38 mm bore ram accelerator using 45–90 g projectiles at velocities up to 2500 m/s.This article was processed using Springer-Verlag TEX Shock Waves macro package 1990. 相似文献
12.
Three-dimensional unsteady Euler simulations are presented for the interaction of a streamwise vortex with an oblique shock
of angle β = 23.3° at Mach 3 and 5. The flowfield features are analyzed for weak, moderate and strong interaction regimes. The details
of the free recirculation zone at conditions of subsonic and supersonic flow on the vortex axis are considered. The vortex
breakdown under conditions of a subsonic vortex core is characterized by a continuous growth and gradual degeneration of the
region, unlike the supersonic core condition wherein a steady recirculation zone is achieved. The possibility of using a localized
steady and pulsed periodic energy deposition on the vortex axis for stimulating the breakdown is demonstrated for various
interaction regimes. It is shown that the formation of a subsonic wake downstream of an energy source lying on the vortex
axis contributes to a more significant growth of the dimensions of the recirculation zone compared to the case when the vortex
core remains supersonic. The possibility of achieving the effects similar to the steady case is demonstrated by the effect
of a pulsed periodic energy source on the flow under consideration for corresponding equivalence parameters.
相似文献
13.
The main aim of this article is to provide a theoretical understanding of the physics of supersonic mixing and combustion. Research in advanced air-breathing propulsion systems able to push vehicles well beyond $M=4$ M = 4 is of interest around the world. In a scramjet, the air stream flow captured by the inlet is decelerated but still maintains supersonic conditions. As the residence time is very short $(\sim \!\!\mathrm{1ms})$ ( ~ 1 ms ) , the study of an efficient mixing and combustion is a key issue in the ongoing research on compressible flows. Due to experimental difficulties in measuring complex high-speed unsteady flowfields, the most convenient way to understand unsteady features of supersonic mixing and combustion is to use computational fluid dynamics. This work investigates supersonic combustion physics in the Hyshot II combustion chamber within the Large Eddy simulation framework. The resolution of this turbulent compressible reacting flow requires: (1) highly accurate non-dissipative numerical schemes to properly simulate strong gradients near shock waves and turbulent structures away from these discontinuities; (2) proper modelling of the small subgrid scales for supersonic combustion, including effects from compressibility on mixing and combustion; (3) highly detailed kinetic mechanisms (the Warnatz scheme including 9 species and 38 reactions is adopted) accounting for the formation and recombination of radicals to properly predict flame anchoring. Numerical results reveal the complex topology of the flow under investigation. The importance of baroclinic and dilatational effects on mixing and flame anchoring is evidenced. Moreover, their effects on turbulence-scale generation and the scaling law are analysed. 相似文献
14.
Supersonic H2-air combustions behind oblique shock waves 总被引:1,自引:0,他引:1
In order to study the mechanisms of initiation and stabilization of H2-Air combustions (stoechiometric mixture initially atT
0=293 K andp
0=0.5 bar) in supersonic flow conditions behind an oblique shock wave (OSW), an original technique is used where OSW is generated in this mixture by the lateral expansion of the burnt gas behind a normal CJ gaseous detonation propagating into a bounding reactive mixture. Four Mach numberM of propagation of OSW are considered in the study, namelyM=7.7-6.1-4.4 and 3. Depending on the Mach numberM and inclinaison angle of OSW different regimes of combustion may occur in the driven mixture. For high values ofM (6.1 and 7.7) delayed steady overdriven oblique detonation waves (SODW) were obtained with a near CJ detonation wave as the critical regime. It was found that SODW obtained correspond quite well to prediction of the polar method. When thermal conditions behind the OSW are lower, either for high Mach number 6.1 and 7.7 for smaller angle than the previous case, or for lower Mach number, 4.4 and 3, the flame initiated at the apex is stabilized as a turbulent oblique flame behind the OSW. With much lower conditions, no combustion appears in the H2-Air mixture. 相似文献
15.
16.
In this paper, we study the well-posedness problem on transonic shocks for steady ideal compressible flows through a two-dimensional
slowly varying nozzle with an appropriately given pressure at the exit of the nozzle. This is motivated by the following transonic
phenomena in a de Laval nozzle. Given an appropriately large receiver pressure P
r
, if the upstream flow remains supersonic behind the throat of the nozzle, then at a certain place in the diverging part of
the nozzle, a shock front intervenes and the flow is compressed and slowed down to subsonic speed, and the position and the
strength of the shock front are automatically adjusted so that the end pressure at exit becomes P
r
, as clearly stated by Courant and Friedrichs [Supersonic flow and shock waves, Interscience Publishers, New York, 1948 (see
section 143 and 147)]. The transonic shock front is a free boundary dividing two regions of C
2,α flow in the nozzle. The full Euler system is hyperbolic upstream where the flow is supersonic, and coupled hyperbolic-elliptic
in the downstream region Ω+ of the nozzle where the flow is subsonic. Based on Bernoulli’s law, we can reformulate the problem by decomposing the 3 ×
3 Euler system into a weakly coupled second order elliptic equation for the density ρ with mixed boundary conditions, a 2 × 2 first order system on u
2 with a value given at a point, and an algebraic equation on (ρ, u
1, u
2) along a streamline. In terms of this reformulation, we can show the uniqueness of such a transonic shock solution if it
exists and the shock front goes through a fixed point. Furthermore, we prove that there is no such transonic shock solution
for a class of nozzles with some large pressure given at the exit.
This research was supported in part by the Zheng Ge Ru Foundation when Yin Huicheng was visiting The Institute of Mathematical
Sciences, The Chinese University of Hong Kong. Xin is supported in part by Hong Kong RGC Earmarked Research Grants CUHK-4028/04P,
CUHK-4040/06P, and Central Allocation Grant CA05-06.SC01. Yin is supported in part by NNSF of China and Doctoral Program of
NEM of China. 相似文献
17.
This large eddy simulation (LES) study is applied to three different premixed turbulent flames under lean conditions at atmospheric
pressure. The hierarchy of complexity of these flames in ascending order are a simple Bunsen-like burner, a sudden-expansion
dump combustor, and a typical swirl-stabilized gas turbine burner–combustor. The purpose of this paper is to examine numerically
whether the chosen combination of the Smagorinsky turbulence model for sgs fluxes and a novel turbulent premixed reaction
closure is applicable over all the three combustion configurations with varied degree of flow and turbulence. A quality assessment
method for the LES calculations is applied. The cold flow data obtained with the Smagorinsky closure on the dump combustor
are in close proximity with the experiments. It moderately predicts the vortex breakdown and bubble shape, which control the
flame position on the double-cone burner. Here, the jet break-up at the root of the burner is premature and differs with the
experiments by as much as half the burner exit diameter, attributing the discrepancy to poor grid resolution. With the first
two combustion configurations, the applied subgrid reaction model is in good correspondence with the experiments. For the
third case, a complex swirl-stabilized burner–combustor configuration, although the flow field inside the burner is only modestly
numerically explored, the level of flame stabilization at the junction of the burner–combustor has been rather well captured.
Furthermore, the critical flame drift from the combustor into the burner was possible to capture in the LES context (which
was not possible with the RANS plus k–ɛ model), however, requiring tuning of a prefactor in the reaction closure. 相似文献
18.
Shock tube studies of thermal radiation of diesel-spray combustion under a range of spray conditions
A tailored interface shock tube and an over-tailored interface shock tube were used to measure the thermal energy radiated
during diesel-spray combustion of light oil, α-methylnaphthalene and cetane by changing the injection pressure. The ignition
delay of methanol and the thermal radiation were also measured. Experiments were performed in a steel shock tube with a 7
m low-pressure section filled with air and a 6 m high-pressure section. Pre-compressed fuel was injected through a throttle
nozzle into air behind a reflected shock wave. Monochromatic emissive power and the power emitted across all infrared wavelengths
were measured with IR-detectors set along the central axis of the tube. Time-dependent radii where soot particles radiated
were also determined, and the results were as follows. For diesel spray combustion with high injection pressures (from 10
to 80 MPa), the thermal radiation energy of light oil per injection increased with injection pressure from 10 to 30 MPa. The
energy was about 2% of the heat of combustion of light oil at P
inj = about 30 MPa. At injection pressure above 30 MPa the thermal radiation decreased with increasing injection pressure. This
profile agreed well with the combustion duration, the flame length, the maximum amount of soot in the flame, the time-integrated
soot volume and the time-integrated flame volume. The ignition delay of light oil was observed to decrease monotonically with
increasing fuel injection pressure. For diesel spray combustion of methanol, the thermal radiation including that due to the
gas phase was 1% of the combustion heat at maximum, and usually lower than 1%. The thermal radiation due to soot was lower
than 0.05% of the combustion heat. The ignition delays were larger (about 50%) than those of light oil. However, these differences
were within experimental error.
An abridged version of this paper was presented at the 18th Int. Symposium on Shock Waves at Sendai, Japan during July 21 to 26, 1991 and at the 19th Int. Symposium on Shock Waves at Marseille, France during July 26 to 30, 1993. 相似文献
An abridged version of this paper was presented at the 18th Int. Symposium on Shock Waves at Sendai, Japan during July 21 to 26, 1991 and at the 19th Int. Symposium on Shock Waves at Marseille, France during July 26 to 30, 1993. 相似文献
19.
Detailed numerical and experimental investigations of pseudo-shock systems in a Laval nozzle with parallel side walls are
carried out. The location of the pseudo-shock system is defined in this system of two choked Laval nozzles by the ratio of
the critical cross sections A2*/A1*{{A}_{2}^*/{A}_{1}^*} , the stagnation pressure loss across the shock system and viscous losses. The wall pressure distributions and high-speed
schlieren videos recorded in the experiments are compared to the results of a steady and an unsteady numerical simulation.
For the steady case, good agreement is found between the calculated and measured shock structure and pressure distribution
along the primary nozzle wall, except for a remaining slight deviation in the shock position. For the unsteady case, in which
asymmetric shock configurations are observed, deviations of the results with respect to the stochastic wall attachment of
the shock system are given which indicate the necessity of further investigations on that topic. 相似文献
20.
模型超燃冲压发动机内着火过程分析 总被引:26,自引:0,他引:26
在燃烧室入口来流为Ma=2.64、T0=1483K、P0=1.65MPa、T=724K、P=76.3kPa条件下,采用高速摄影和连续激光高速纹影对等截面型开窗燃烧室内氢气射流自燃过程、火花塞点燃氢气过程和引导氢气火焰点燃煤油过程进行了观测,获得了燃烧室内着火过程中火焰和流场波系结构的动态演化过程;观察到了初始火焰区首先起始于燃烧室下游,并逆流传播实现发动机着火的过程;分析表明燃料能否着火、以及着火位置与燃料着火时间、燃烧室流速和火焰稳定器安装情况相关,多火焰稳定区延长了燃料驻留时间,使燃料更容易着火。关键词 超燃冲压发动机,点火过程,火焰传播,火焰稳定器 相似文献