Firing tests of a liquefied air cycle rocket engine (LACE)

The problem of implementing the detonation mode of fuel combustion in thermal propulsion systems has been widely studied last decade. There are many works on fundamental and applied research on pulsating detonation. Solid propellant detonation engines can develop significant forces for a short time at low structural masses, and therefore they are ideal for auxiliary systems for the removal of separated rocket parts. In addition, detonation processes can be used to create control forces for correcting the trajectory of aircraft. All these facts determine the relevance of the area of work. For studying detonation installations, it is necessary to create test stands, but the design of test installations is an urgent and complex optimization problem. It is advisable to solve this problem with the help of computer simulation. In the existing experimental methods, for designing, it is necessary to determine in advance the geometric parameters of receivers and pipelines that provide the necessary gas consumption for firing tests of detonation rocket engines. The work is devoted to the development of a method for determining the flow characteristics of a receiver with a pipeline of complex configuration based on the constructed model of the stand. Based on the initial data, a computer simulation of the air leakage process from the receiver was carried out, for which the Solid Works software package was used. The places of pressure drop, maximum flow rate, and air mass flow are determined. The low value of the flow rate factor is due to the complex configuration of the pipeline with numerous bends and two bellows. Comparison of calculation results with experimental data was held. The difference between the experimental and calculated values does not exceed 3.6%. The obtained information is used to select the required value of the oxidizer excess coefficient during firing tests of detonation rocket engine models. Keywords: flow rate, gas leakage, receiver, model.

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Design and Experimental Investigation on Gas Oxygen/Kerosene Ejector Rocket for RBCC Application

Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University ◽

10.1051/jnwpu/20183630558 ◽

2018 ◽

Vol 36 (3) ◽

pp. 558-564

Author(s):

Xianggeng Wei ◽

Fei Qin ◽

Lei Shi ◽

Baoqing Zhang ◽

Guoqiang He

Keyword(s):

Working Conditions ◽

Rocket Engine ◽

Combined Cycle ◽

Engine Test ◽

The Core ◽

Technical Requirements ◽

Core Components ◽

Design Requirements ◽

Working Parameters ◽

Firing Tests

The ejector rocket is one of the core components of the rocket based combined cycle propulsion system, and must be capable of variable working conditions. In order to meet technical requirements for RBCC application, the variable duty operating ejector rocket using the gas Oxygen/Kerosene was designed based on the gas pressurized propellant feed systems. Hot firing tests of four different working conditions had been completed. Experimental results show that the designed ejector rocket engine was stable and reliable, and the working parameters met the design requirements, and the working conditions were adjusted quickly. It lays a foundation for the study of the RBCC engine test and the engine technology of large adjustment ratio.

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Fault Detection of a Reusable Rocket Engine using Phase Plane Trajectory of Feature Vectors

Annual Conference of the PHM Society ◽

10.36001/phmconf.2019.v11i1.771 ◽

2019 ◽

Vol 11 (1) ◽

Author(s):

Seiji Tsutsumi ◽

Miki Hirabayashi ◽

Daiwa Sato ◽

Masaharu Abe ◽

Kaname Kawatsu ◽

...

Keyword(s):

Fault Detection ◽

Present Method ◽

Phase Plane ◽

Rocket Engine ◽

Principal Component ◽

Sensor Failure ◽

Plane Trajectory ◽

Feature Vectors ◽

Firing Test ◽

Firing Tests

A bivariate time-series analysis based on the phase plane trajectory of feature vectors extracted by principal component analysis is developed for fault detection in a reusable liquid-propellant rocket engine. Static-firing test results of the reusable rocket engine obtained at the Japan Aerospace Exploration Agency are employed for demonstration of the present method. The present method successfully detected temperature sensor failure from 19 firing tests of 62 sensors, even in the deviation of the engine operational sequence between the static-firing tests. The present method was also able to detect the system failure from 23 firing tests. Furthermore, the ability to distinguish the system and sensor failure was demonstrated.

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Hot Firing Tests of Liquid Rocket Engine Using LOX/LNG

40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit ◽

10.2514/6.2004-3528 ◽

2004 ◽

Cited By ~ 1

Author(s):

Yong-Ho Cho ◽

Haeng-Soo Chang

Keyword(s):

Rocket Engine ◽

Liquid Rocket Engine ◽

Liquid Rocket ◽

Firing Tests

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Evaluation of the high-frequency oscillation parameters of a liquid-propellant rocket engine with an annular combustion chamber

Technical mechanics ◽

10.15407/itm2021.01.016 ◽

2021 ◽

Vol 2021 (1) ◽

pp. 16-28

Author(s):

O.D. Nikolayev ◽

◽

I.D. Bashliy ◽

N.V. Khoriak ◽

S.I. Dolgopolov ◽

...

Keyword(s):

Finite Element ◽

Combustion Chamber ◽

Rocket Engine ◽

Element Model ◽

Numerical Approach ◽

Rocket Engines ◽

Liquid Propellant ◽

Liquid Propellant Rocket Engine ◽

Firing Tests ◽

Propellant Rocket

The high-frequency instability (HF instability) of a liquid-propellant rocket engine (LPRE) during static firing tests is often accompanied by a significant increase in dynamic loads on the combustion chamber structure, often leading to the chamber destruction. This dynamic phenomenon can also be extremely dangerous for the dynamic strength of a liquid-propellant rocket engine with an annular combustion chamber. Computation of the parameters of acoustic combustion product oscillations is important in the design and static firing tests of such rocket engines. The main aim of this paper is to develop a numerical approach to determining the parameters of acoustic oscillations of combustion products in annular combustion chambers of liquid-propellant rocket engines taking into account the features of the configuration of the combustion space and the variability of the physical properties of the gaseous medium depending on the axial length of the chamber. A numerical approach is proposed. The approach is based on mathematical modeling of natural oscillations of a “shell structure of an annular chamber – gas” coupled dynamic system by using the finite element method. Based on the developed finite-element model of coupled spatial vibrations of the structure of the annular combustion chamber and the combustion product oscillations, the oscillation parameters of the system under consideration (frequencies, modes, and effective masses) for its dominant acoustic modes, the vibration amplitudes of the combustion chamber casing, and the amplitudes of its vibration accelerations can be determined. The operating parameters of the liquid-propellant rocket engine potentially dangerous for the development of thermoacoustic instability of the working process in the annular combustion chamber can be identified. For the numerical computation of the dynamic gains (in pressure) of the combustion chamber, a source of harmonic pressure excitation is introduced to the finite element model of the dynamic system “shell structure of an annular configuration – gas” (to the elements at the start of the chamber fire space). The developed approach testing and further analysis of the results were carried out for an engine with an annular combustion chamber (with a ratio of the outer and inner diameters of 1.5) using liquid oxygen – methane as a propellant pair. The system shapes and frequencies of longitudinal, tangential and radial modes are determined. It is shown that the frequency of the first acoustic mode in the case of a relatively low stiffness of the combustion chamber casing walls can be reduced by 40 percent in comparison with the frequency determined for a casing with rigid walls.

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