Conceptual Investigations for a Methane-Fueled Expander Rocket Engine

Brown, Corey D.

doi:10.2514/6.2004-4210

Cited by 21 publications

(12 citation statements)

References 1 publication

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“…[1][2][3][4][5] Methane is known to be a low-density hydrocarbon having advantages that are similar to those of both kerosene and hydrogen. It is also a cryogenic fuel having a density about six times higher than that of hydrogen, and it requires less insulation and fewer handling concerns than hydrogen.…”

Section: Introductionmentioning

confidence: 99%

Flowfield and Heat Transfer Characteristics of Cooling Channel Flows in a Methane-Cooled Thrust Chamber

Negishi¹,

Daimon²,

Kawashima³

et al. 2012

48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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In recent years, methane has attracted attention as a propellant for liquid rocket engines because of its various advantages compared to typical propellants such as hydrogen. When methane is used as a coolant for a regenerative cooling system, its near-critical thermodynamic and transport properties experience large variations because its critical pressure is higher than that of typical propellants; this significantly influences the flowfield and heat transfer characteristics. Therefore, adequate understanding of the flowfield and heat transfer characteristics of methane in regenerative cooling channels is a prerequisite for future engine development. In this study, conjugated coolant and heat transfer simulations were performed to investigate the flowfield and heat transfer characteristics of transcritical methane flows in a sub-scale methane-cooled thrust chamber. The computed results were validated against experimental data measured in hot firing tests. They compared well with the measured pressures and temperatures in cooling channels, and wall temperatures were within the permitted levels. Detailed flow analysis revealed peculiar flow structures in the cooling channel: a strong secondary flow induced in the concave-heated part in the channel throat section and the coexistence of two different gas phases-ideal and real-in a single cross-section in the cylindrical region. A high wall temperature appeared in the cylindrical region of the thrust chamber under the considered conditions; this was due to the heat transfer deterioration induced by an M-shaped velocity profile and a turbulentheat flux reduction. Nomenclature a, b,  = coefficients in cubic-type equation of state C p = constant pressure heat capacity G = mass flow rate H = channel height h = heat transfer coefficient Pr = Prandtl number p = pressure q = heat flux R = gas constant Re = Reynolds number R u = universal gas constant s = normal distance T = temperature u = streamwise velocity W = molecular weight x, y, z = coordinate system Z = compressibility factor y + = normalized wall distance 2  = thermal conductivity  = viscosity  = density Subscripts b = bulk c = coolant cr = critical cw = coolant-side wall gw = hot-gas-side wall in = inlet out = outlet t = turbulent throat = throat pcr = pseudo-critical w = wall

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Section: Introductionmentioning

confidence: 99%

Flowfield and Heat Transfer Characteristics of Cooling Channel Flows in a Methane-Cooled Thrust Chamber

Negishi¹,

Daimon²,

Kawashima³

et al. 2012

48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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“…22 Institutions in the USA, Russia, China, and Europe have already done some work on the research of LOX/methane LRE, including its thermodynamic cycle, flame stabilization, comparison of kerosene and methane as rocket fuels, and so on. [23][24][25] In addition to the balancing characteristics in thermodynamic properties and specific impulse, researches indicate three more positive properties of methane as an aerospace fuel. First, methane is a ''clean'' fuel which means that it has less soot deposition as its coking point is roughly twice as high as kerosene.…”

Section: Analysis Of Lox/methane Liquid-rocket Engine (Lre)mentioning

confidence: 99%

Conceptual study of a dual-rocket-based-combined-cycle powered two-stage-to-orbit launch vehicle

Zhang

Wang

2017

Proceedings of the Institution of Mechanical Engineers, Part G:

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The liquid oxygen/methane staged cycle liquid-rocket engine is one of the most potential rocket engines in the future for its higher performance, higher fuel density and reusable capacity. Two working states of this liquid-rocket engine named as full-load state and half-load state are defined in this paper. Based on this liquid-rocket engine, a dual-rocket-basedcombined-cycle propulsion system with liquid oxygen /air/methane as propellants is therefore proposed. The dualrocket-based-combined-cycle system has then five working modes: the hybrid mode, pure ejector mode, ramjet mode, scramjet mode and pure rocket mode. In hybrid mode, the booster and ejector rockets driven by the full-load liquidrocket engine work together with the purpose of reducing thrust demand on ejector rocket. In scramjet mode, the fuelrich burned hot gas generated by the half-load liquid-rocket engine is used as fuel, which is helpful to reduce the technical difficulty of scramjet in hypersonic speed. The five working modes of dual-rocket-based-combined-cycle are highly integrated based on the full-or half-load state of the liquid oxygen/methane staged cycle liquid-rocket engine, and the unified single type fuel of liquid methane is adopted for the whole modes. Then a preliminary design of a horizontal takeoff two-stage-to-orbit launch vehicle is conducted based on the dual-rocket-based-combined-cycle propulsion system. Under an averaged baseline thrust and specific impulse, the launch trajectory to reach a low Earth orbit at 100 km is optimized via the pseudo-spectral method subject to maximizing the payload mass. It is shown that the two-stage-to-orbit vehicle based on the dual-rocket-based-combined-cycle can achieve the payload mass fraction of 0.0469 and 0.0576 for polar mission and equatorial mission, respectively. Conclusively, insights gained in this paper can be usefully applied to a more detailed design of the dual-rocket-based-combined-cycle powered two-stage-to-orbit launch vehicle.

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“…Liquid rocket engines (LREs) have to sustain impressive thermal loads and mechanical stresses since combustion gas temperature (up to about 3600 K), and high pressure reign in the thrust chamber [2]. The operation of the modern LREs is enabled by the use of advanced cooling jackets.…”

Section: Mtp Breadboardmentioning

confidence: 99%

“…In the HYPROB final demonstrator, the refrigerant is foreseen to be injected into the cooling system as a compressed liquid, with temperature and pressure values lower and higher than the critical ones (Tcr = 190.56 K and Pcr = 4.59 MPa), respectively. Combustion gases warm the fluid, which exits the cooling jacket as a supercritical gas [2]. For these reasons, the MTP working conditions include inlet pressure and temperature values ranging from 8.0 to 15.0 MPa and 130-140 K, respectively.…”

Section: Mtp Breadboardmentioning

confidence: 99%

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Thermal analyses supporting the development of a liquid rocket engine

Ricci¹,

Natale²,

Battista³

et al. 2016

IJHT

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CIRA manages the HYPROB Program, supported by the Italian Ministry of University and Research, with the aim of improving the national capabilities to develop rocket engines for future applications. A system line, named HYPROB-BREAD, focusing on LOX/LCH4 technology, is included. The final goal is the design and test of a regenerative LOX/LCH4 demonstrator. A step-by-step approach has been adopted to validate critical design aspects by simplified technological breadboards. The reliable operation of an engine is ensured by thermally efficient cooling jackets, which require the in-depth comprehension of the coolant behaviour. For this purpose, the MTP-BB has been tested. Another important issue is the evaluation of the thermal loads, transferred by the combustion hot gases to the thrust chamber walls. In this view, a Subscale Calorimetric Breadboard has been designed; 13 disks surround the chamber: they are fed up by water and provide the cooling and the measurement of the exchanged thermal power. The final article is a 3-ton-class LOX/LCH4 regenerative demonstrator, whose coolant is liquid methane, flowing in a cooling system, made by several axial channels. This paper aims at describing the thermal investigations, conducted in the design and the verification phase for the aforementioned breadboards and demonstrator cooling jacket.

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Conceptual Investigations for a Methane-Fueled Expander Rocket Engine

Cited by 21 publications

References 1 publication

Flowfield and Heat Transfer Characteristics of Cooling Channel Flows in a Methane-Cooled Thrust Chamber

Flowfield and Heat Transfer Characteristics of Cooling Channel Flows in a Methane-Cooled Thrust Chamber

Conceptual study of a dual-rocket-based-combined-cycle powered two-stage-to-orbit launch vehicle

Thermal analyses supporting the development of a liquid rocket engine

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