A series of experiments were conducted to characterize the self-ignition and combustion of thermally cracked kerosene in both a Mach 2.5 model combustor with a combustor entrance height of 51 mm and a Mach 3.0 model combustor with an entrance height of 70 mm. A unique kerosene heating and delivery system was developed, which can prepare heated kerosene up to 950 K at a pressure of 5.5 MPa with negligible fuel coking. The extent of China no. 3 kerosene conversion under supercritical conditions was measured using a specially designed system. The compositions of gaseous products as a result of thermal cracking were analyzed using gas chromatography. The mass flow rates of cracked kerosene were also calibrated and measured using sonic nozzles. With the injection of thermally cracked kerosene, the ability to achieve enhanced combustion performance was demonstrated under a variety of airflow and fuel conditions. Furthermore, self-ignition tests of cracked kerosene in a Mach 2.5 model combustor over a range of fuel injection conditions and with the help of different amounts of pilot hydrogen were conducted and discussed.
In order to obtain a minimized weight design for a hypersonic vehicle, a design system that involves integrating aeroheating analysis with thermal protection system (TPS) sizing has been developed. The design system has the capability to analyse aeroheating and TPS thermal response and to provide an optimization capability to determine the minimum thickness of the TPS materials required to prevent the underlying structure from exceeding its operating temperature limit. Features of this system are the extended capability to analyse thermal response of metallic multi-layer TPS with non-grey insulation and the smart TPS for next generation or future TPS design. The application of the above design system to a model spacecraft is presented, showing good agreement with previously obtained results, and proving its reliability for future design use. The application of the extended capabilities also shows that the above design system has strong applicability to the following generations of TPS materials. and 'dynamic' TPS sizing [3,4]. Based on the concept of dynamic TPS sizing, Cowart and Olds [4] developed a tool for automated TPS design, in which the aeroheating was predicted by the fast engineering code MINIVER (MINIature VERsion aero-thermal analysis program) because dynamic TPS sizing calls for continuous updating of TPS parameters as the design process proceeds. Dynamic TPS sizing is regarded as the preferred method of the two cited above [3]. The design system developed in this paper is based on the concept of dynamic TPS sizing of various TPS models, such as metallic multi-layer TPS with non-grey insulation and smart TPS. TPS DESIGN AND SIZING SYSTEMThe present system structure is shown in Fig. 1. The ACAT provides the aerothermal environments required for the TPS design. The TPS-MDPT provides parameters of thermal protection materials and material mapping on the surface by the heat rate and temperature obtained from the ACAT. The TPS-TAT provides a means to calculate the transient in-depth conduction from the surface of the TPS. The OCT JAERO125
A pulsed schlieren system using a nanosecond-flashlamp has been developed to study the flow-field of supersonic combustion. A high speed camera was used to capture the instantaneous flow structures. The camera and the flash-lamp source were synchronized using a pulse generator.Due to the extremely high luminous density of the spark, the background illumination has been suppressed effectively with an exposure time of order of microsecond. The instantaneous schlieren images of injection and combustion of pilot hydrogen and vaporized kerosene in a Ma 2.5 combustor have been captured and analyzed.
In this research, a thermal analysis method has been developed to analyze the heat transfer process associated with endothermic hydrocarbon regenerative-cooled structure of a combustor. The complex heat transfer processes relevant to such cooling structure exposed in severe heat environment are modeled by three coupled processes: hot side boundary condition specification, flow and convective heat transfer of fuel within cooling channels and heat transfer in combustor structure with cooling channels embeded. To speed up the simulation process while achieving good accuracy, efforts are made in several aspects: first, hot side heat environment of combustor is obtained either by measurement results from upgraded heat flux sensors developed based on the principle of Gardon heat-flux gauge or by quasi 1-D analysis of combustor using static pressure distribution as input; second, to quantitatively describe flow and heat transfer behavior of hydrocarbon coolant when heavy cracking happens, a five-component surrogate model is developed and together used with a thermal cracking model consisting of 18 species and 24 reactions; third, to account for multiple effects happening in redistribution channels, a special method for flow rate redistribution prediction is developed based on characteristic time scale analysis. The validness of this integrated analysis method is tested by comparing simulation results with measurement data from lab tests of a supersonic model combustor. The overall correctness of exit coolant fuel temperature and wall temperature distribution prediction is within 5% and 10% separately. The fast speed and decent accuracy of this method developed make it very promising to be put into use for hydrocarbon regenerative-cooling analysis. NomenclatureA = area S = perimeter q = heat flux ρ = density h = enthalpy htc = heat transfer coefficient w T = temperature of wall , f T = temperature of fuel w τ = friction of wall f C = friction coefficient t S = Stanton number u N = Nusselt number r P = Prandtl number
In this paper, a numerical analysis coupling heat transfers of the combustor internal flow, the coolant flow and the strut wall is developed and applied for the optimization of strut cooling using aviation kerosene as coolant at flow conditions corresponding to the combustor entrance condition for Mach 6 scramjet flight. The coupling procedure is tested and proven to be an efficient method of being capable to obtain the converged temperature and heat transfer solutions of the cooled strut within a few iteration steps. Four cooling designs with varied diameter, length and position of the cooling channels are investigated and their improvements on fuel injection and mixing are also verified compared to the wall injection. The kerosene-cooled strut (Strut4) is tested in a Mach 2.5 supersonic tunnel with inlet total temperature and total pressure of 1900K and 1.45MPa respectively for 60 seconds. The damaged part in the upper leading edge of the strut is observed, which is consistent with the result obtained by the numerical analysis.
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