Blowout Limits of Cavity-Stabilized Flame of Supercritical Kerosene in Supersonic Combustors

Zhang, Taichang; Wang, Jing; Li, Qi; Fan, Xuejun; Zhang, Peng

doi:10.2514/1.b35120

Cited by 25 publications

(8 citation statements)

References 22 publications

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“…The characteristic length scale is the height (H) of the wall-cavity. Cavity flame holders have been studied extensively [7][8][9][10][11][12]. RR F is the maximum reaction rate of the fuel (in 1/seconds) that is defined in Section 5 below.…”

Section: Approachmentioning

confidence: 99%

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A Method to Compute Flameout Limits of Scramjet-Powered Hypersonic Vehicles

Mbagwu

Driscoll

2015

51st AIAA/SAE/ASEE Joint Propulsion Conference

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As a hypersonic vehicle travels upward along an ascent trajectory, the static pressure (p 3 ) in the scramjet combustor will decrease, which can lead to engine flameout. At low pressures the chemical reactions between the fuel and air become excessively slow. However, during the ascent the flight Mach number is increasing; this increases the stagnation temperature and the static temperature (T 3 ) at the combustor entrance so it tends to prevent flameout. To investigate this tradeoff, a general method to understand how the flameout limit varies during ascent was developed. The method consists of two parts; first the static temperature and pressure at the entrance to the combustor (T 3 , p 3 ) are computed as a function of the vehicle altitude; this is done using a reduced-order propulsion model called MASIV. In the second part the values of (T 3 , p 3 ) are inserted into an empirical relation for the critical Damkohler number in order to determine if flameout occurs or not at each altitude. The empirical flameout relation is based on previous ground-based measurements made at AFRL and elsewhere.

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Section: Approachmentioning

confidence: 99%

“…Resulting profiles ofω H2 (x) appear in Figure 4b. This value is the react nrate that is required in equation (12). The final step is to compute the maximum fuel reaction rate (RR F ) that is defined to be:…”

Section: Masiv Combustor Modelmentioning

confidence: 99%

A Method to Compute Flameout Limits of Scramjet-Powered Hypersonic Vehicles

Mbagwu

Driscoll

2015

51st AIAA/SAE/ASEE Joint Propulsion Conference

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“…For the cases where both T 0 and are relatively large, the substantial amount of heat release cause a significant rise of pressure in the isolator so as to cause the the change of the inlet flow conditions: a case referred to as combustor "unstart". Such "rich" blowout and "unstart" pheomena have been discussed in detail in the recent study of the authors [13] .…”

Section: Regime Nomogram Of Flame Stabilization Modesmentioning

confidence: 99%

“…All the experiments were conducted in a Ma = 2.5 direct-connect test facility, which has been detailedly described in Zhang et al [12][13][14] . The facility consists of a vitiated air supply system and a multipurpose supersonic model combustor.…”

Section: Experimental Specifications and Analysismentioning

confidence: 99%

Characterization of flame stabilization modes in an ethylene-fueled supersonic combustor using time-resolved CH* chemiluminescence

Yuan

Zhang

Yao

et al. 2017

Proceedings of the Combustion Institute

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Flame stabilization in a Ma = 2.5 direct-connect supersonic combustor was experimentally characterized with a fixed stagnation pressure of 1.0 MPa and wide ranges of stagnation temperature ( T 0 ) from 1200 K to 1800 K and global equivalence ratio ( ) of ethylene/air from 0.1 to 0.8. Four typical flame stabilization modes were identified by using the time resolved CH * chemiluminescence and presented as a regime nomogram in the T 0 -parameter space. As increasing T 0 , the range of for the flame stabilization modes is widened and that for the oscillation mode is therefore narrowed. The widely used quasi-1D analysis, with the experimentally determined wall static pressure distribution as input parameters, suggests that the combustor operates in a scramjet mode when the flame is stabilized in the cavity shear layer and in a ramjet mode when the flame is in the jet-wake. The flame oscillation mode, observed only for a narrow range of , was found to correlate with the combustor transition between the scramjet and ramjet modes.

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“…At flight Mach number about 6-8, the total temperature of gas after combustion may exceed 2800 K and the average wall heat flux to the combustor inner wall ranges from 1.0 to beyond 10.0 MW/m 2 depending on the combustor design and the flight condition [1]. In order to deal with the challenge of combustor cooling and improve fuel mixing and combustion efficiency, liquid kerosene is heated to a supercritical state or even cracked gas by convective heat transfer along the cooling channels in the combustor walls before reaching the fuel injector [2]. The heat flux distribution might be changed with supercritical kerosene.…”

Section: Introductionmentioning

confidence: 99%

Measurement of heat flux distribution of supercritical kerosene fueled supersonic combustor

Cheng¹,

Wang²,

Gong³

et al. 2016

32nd AIAA Aerodynamic Measurement Technology and Ground Testing Conference

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Heat flux distribution of supercritical kerosene fueled, single-side expanded supersonic combustor with two dislocated cavities was experimentally studied. The effects of inlet Mach number, total temperature, mass flow rate and fuel equivalence ratio on the combustion and heat transfer characteristics in the supersonic combustor were analyzed. The isolator inlet Mach number is 2.0 and 2.5, the total temperature is 1305 K to 1701 K and the mass flow rate is 2.0 kg/s to 3.0 kg/s. Pilot hydrogen and liquid or supercritical (773±20 K) China No.3 kerosene were injected in front of the cavities with the equivalence ratio ranges from 0.52 to 0.88. Results show that heat flux increases with the inlet temperature and mass flow rate, however, the influence of equivalence ratio is non-monotonic in the range of this study. The two inlet Mach numbers also trigger different combustion modes, which further complicates the heat flux distribution. In the end, a three parameter correlation is proposed and fitted to normalize the experiment results for comparison and discussion. Nomenclaturė = Heat flux qm = Mass flow ϕ = Equivalence ratio K1, K2 = Sensitivity of heat flux and temperature difference of the heat flux sensor St = Stanton number Re = Reynolds number Pr = Prandtl number Ma = Isolator inlet Mach number E = Output signal of heat flux sensor Tb = Body temperature of heat flux sensor 1 Engineer, Th = Temperature at the head of the heat flux sensor D = Diameter of the heat flux sensor P = Heating power for calibration δ = Thickness of the cooling inner wall cp = Specific heat at constant pressure x = Axial distance from inlet Subscript 0 = Stagnation condition Superscript * = Reference state of Eckert's Reference Enthalpy Method

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Blowout Limits of Cavity-Stabilized Flame of Supercritical Kerosene in Supersonic Combustors

Cited by 25 publications

References 22 publications

A Method to Compute Flameout Limits of Scramjet-Powered Hypersonic Vehicles

A Method to Compute Flameout Limits of Scramjet-Powered Hypersonic Vehicles

Characterization of flame stabilization modes in an ethylene-fueled supersonic combustor using time-resolved CH* chemiluminescence

Measurement of heat flux distribution of supercritical kerosene fueled supersonic combustor

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