Combustion Chamber Of Liquid Rocket Engine Research Articles

Chemical modeling for methane oxy-combustion in Liquid Rocket Engines

Methane–oxygen burning is considered for many future rocket engines for practicality and cost reasons. As this combustion is slower than hydrogen–oxygen, flame ignition and stability may be more difficult to obtain. To address these questions, numerical simulation with realistic chemistry is appropriate. However the high pressure and turbulence intensity encountered in rocket engines enhance drastically the stiffness of methane oxy-combustion. In this work, Analytically Reduced Chemistry (ARC) is proposed for accurate chemistry description at a reasonable computational cost. An ARC scheme is specifically derived for typical rocket engine conditions. It is validated by comparison with its parent skeletal mechanism on a series of laminar flames. Then the numerical stiffness of chemistry is overcome with an original approach for time integration, allowing to run simulations close to the acoustic time step whatever the chemical stiffness. It is demonstrated on laminar cases that the flame structure is well preserved, and that numerical stability is ensured while decreasing significantly the computational cost. The performance of ARC with the fast time integration method is finally demonstrated in a 3D Large-Eddy Simulation of a lab-scale Liquid Rocket Engine combustion chamber, where a detailed flame analysis is conducted.

Математическое моделирование течения охладителя в тракте охлаждения камеры ЖРД с предельно высокой степенью оребрения

The paper presents a computational analysis of coolant distribution in the cooling channel of a liquid rocket engine combustion chamber, performed in order to develop a set of practical guidelines towards increasing efficiency of a cooling system featuring an extremely high degree of ribbing. We created a three-dimensional mathematical model comprising a closed system of hydrodynamic equations as well as initial and boundary conditions for an element of the liquid rocket engine chamber we modelled, the chamber featuring longitudinal cooling channel arrangement manufactured via additive technology. We computed velocity and pressure fields in characteristic cooling channel regions for various levels of coolant mass flow rate, which confirmed the feasibility of the layout proposed in terms of uniform coolant distribution in the cooling channel of the liquid rocket engine modelled. We obtained the friction loss ξ as a function of coolant mass flow rate and particle size of the powder used in the additive technology to manufacture the combustion chamber wall and cooling channel.

Direct Current Forcing of an Atmospheric Multiburner Flame for Rocket Combustor Emulation

This study investigated the effect of direct current electric fields on the response of an atmospheric, premixed methane-air flame using electrode geometries intended to emulate a simplified liquid rocket engine combustion chamber. The burners consisted of single- and multielement configurations to determine the difference between element radial location and flame-to-flame interaction as an atmospheric analogy to the functionality of a multiport liquid-propellant rocket engine injector. The results showed that electric field–induced ionic winds were capable of improving flame stability by extension of the lean flammability limit and increased blowoff velocity. Single-burner configurations closer to the anode wall had more significant effects on the flame response due to stronger electric fields generated at those locations. The cylindrical anode was more effective in changing the flame response due to larger surface area, leading to higher field strengths. It has been concluded that a system with multiple flames enhances the effects of applied electric fields, which seems to indicate the potential to effectively modify the flame response in rocket engines with multiple injection ports.

Coupled Lagrangian impingement spray model for doublet impinging injectors under liquid rocket engine operating conditions

To predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray, the conventional uncoupled spray model for impinging injectors is extended by considering the coupling of the jet impingement process and the ambient gas field. The new coupled model consists of the plain-orifice sub-model, the jet-jet impingement sub-model and the droplet collision sub-model. The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions. The overall model is benchmarked under various impingement angles, jet momentum and off-center ratios. Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics, such as the mass flux and mixture ratio distributions in quiescent air. Besides, impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions. First, a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile. The minimum average droplet diameter is achieved when the orifices work in cavitation state, and is about 30% smaller than the steady single phase state. Second, the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°. The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines.

Анализ особенностей рабочего процесса в камерах сгорания ЖРД со струйно-центробежными и центробежно-центробежными форсунками

Анализ особенностей рабочего процесса в камерах сгорания ЖРД со струйно-центробежными и центробежно-центробежными форсунками

極限マルチフィジクス環境における液体ロケットエンジン燃焼室の破損メカニズムの解明と寿命評価

極限マルチフィジクス環境における液体ロケットエンジン燃焼室の破損メカニズムの解明と寿命評価

179 熱-流体-構造連成解析による液体ロケットエンジン燃焼室の高温非弾性挙動評価

179 熱-流体-構造連成解析による液体ロケットエンジン燃焼室の高温非弾性挙動評価

G030094 マイクロチャンネル冷却構造体のECTを用いた非破壊検査

Nondestructive evaluation (NDE) method was investigated for life prediction of a liquid rocket engine combustion chamber. As a thickness of a cooling channel ligament of a combustion chamber is about 1 mm, high resolution is required for the NDE method. In this study, feasibility study of ECT (Eddy Current Testing) method was conducted by using the high-resolution probe composed of a small excitation coil (diameter of 3 mm and 20 turn) and a high-sensitive AMR (Anisotropic Magneto Resistance) sensor. A plate type specimen with a thickness of 1 mm and a chamber shape specimen with coolant channels made of OFHC (Oxygen Free High-conductivity Copper) were prepared. Artificial defects were machined to the specimens and NDE tests were conducted. As the results of the tests, a small defect with a length of 2 mm and a half depth of the material (0.5-0.6 mm) could be detected with proper excitation frequencies.

Numerical Simulation of the Spray Characteristics in Small Scale Liquid Rocket Engine Combustion Chamber

The mathematical and physical model of the liquid propellant spray in straight nozzle was proposed for studying the performance characteristics of the small-scale liquid rocket engine. With the Fluent software, the numerical simulation was carried out. Sauter mean diameter (SMD) of the HAN-based liquid propellant (LP1846) in the engine combustor changing with spray pressure, nozzle diameter and the liquid surface tension were analyzed. The results indicate that: in the spray pressure region of 1.8MPa~3.0MPa, at a fixed spray pressure, the smaller is the nozzle diameter, the smaller is the droplets’ SMD and the relationship between the SMD and the nozzle diameter is approximately linearity; for the same nozzle diameter and spray pressure, the larger is the surface tension, the larger is the liquid droplets’ SMD.

Subtranscritical and Supercritical Properties for LO2-CH4 at 15MPa

Future launchers will use rocket propulsion systems burning CH 4 /LO 2 at high chamber pressure, and so it is useful to analyze the thermophysical properties of these species and their combustion products at these conditions. High-pressure (real-gas) effects significantly modify combustion regimes, for instance, the propellants Re injected near the critical temperature, and so to simulate mixing and combustion processes in subtranscritical regimes high-pressure effects must be described. This paper analyzes the compressibility factors of CH 4 , O 2 , CO 2 , and H 2 O at a pressure of 15 MPa and calculates the difference between ideal- and real-gas thermophysical properties for these species in the range of temperature in which experimental data are available. Finally, the paper describes thermophysical properties at typical liquid rocket engine combustion chamber conditions (100 < T < 6000 K, 15 MPa) by polynomial fits. As there are no experimental data at high temperatures, theories are necessary to predict properties at temperatures different from the experimental ones: thus, low-temperature experimental data (National Institute of Standards and Technology tables) are used in conjunction with predictions obtained with the Lee-Kesler equation of state (for density and isobaric specific heat) and with the Chung et al. method (for viscosity and thermal conductivity). This paper will try to clarify the impact of subtrans-supercritical parameters on mixing and combustion in future liquid rocket engines.

Friction Stir Welding of GRCop-84 for Combustion Chamber Liners

GRCop-84 is a copper-chromium-niobium alloy developed by the Glenn Research Center for liquid rocket engine combustion chamber liners. GRCop-84 exhibits superior properties over conventional copper-base alloys in a liquid hydrogen -oxygen operating environment. The Next Generation Launch Technology program has funded a program to demonstrate scale-up production capabilities of GR -Cop 84 to levels suitable for main combustion chamber production for the prototype rocket engine. This paper describes a novel method of manufacturing the main combustion chamber liner. The process consists of several steps: extrude the GR -Cop 84 powder into billets, roll the billets into plates, bump form the plates into cylinder halves and friction stir weld the halves into a cylinder. The cylinder is then metal spun formed to near net liner dimensions followed by finish machining to the final configuration. This paper describes the friction stir weld process development including tooling and non-destructive inspection techniques, culminating in the successful production of a liner pre-form completed through spin forming.

액체로켓 연소기용 구리합금의 열/기계적 특성에 관한 실험적 연구

Mechanical and physical properties of a copper alloy for a liquid rocket engine(LRE) combustion chamber liner application were tested at various temperatures. All test specimens were heat treated with the condition they might experience during actual fabrication process of the LRE combustion chamber. Physical properties measured include thermal conductivity, specific heat and thermal expansion data. Uniaxial tension tests were preformed to get mechanical properties at several temperatures ranging from room temperature to 600℃. The result demonstrated that yield stress and ultimate tensile stress of the copper alloy decreases considerably and strain hardening increases as the result of the heat treatment . Since the LRE combustion chamber operates at higher temperature over 400℃, the copper alloy can exhibit time-dependent behavior. Strain rate, creep and stress relaxation tests were performed to check the time-dependent behavior of the copper alloy. Strain rate tests revealed that strain rate effect is negligible up to 400℃ while stress-strain curve is changed at 500℃ as the strain rate is changed. Creep tests were conducted at 250℃ and 500℃ and the secondary creep rate was found to be very small at both temperatures implying that creep effect is negligible for the combustion chamber liner because its operating time is quite short.

Bipropellant combustion in a liquid rocket combustion chamber

A comprehensive computer simulation code is developed for the combustion-flow analysis of a bipropellant liquid rocket engine combustor. A bipropellant group combustion model is built in the computer code to facilitate the determination of various combustion modes, including the combustion of oxidizer droplets, fuel droplets and gas mixture. Several physical submodels are also adopted to account for the interaction between bipropellant droplets and the gas-phase flow. An axisymmetric cylindrical liquid rocket engine combustion chamber installed with two annular-ring injectors is selected to demonstrate the capabilities of the present computer simulation code. Two types of injection processes are analyzed. One is characterized by a non-premixed spray process and the other is a premixed spray process. Selected numerical computations show that the premixed spray results in a better utilization of combustion chamber space and gives a higher combustion efficiency in general. But finer-droplet sprays do not necessarily lead to a higher combustion efficiency as traditionally speculated.