Combustion instability in a small liquid rocket motor

1999 ◽  
Vol 103 (1023) ◽  
pp. 245-252 ◽  
Author(s):  
A. Osherov ◽  
B. Natan
Keyword(s):  
Combustion Chamber ◽  
High Efficiency ◽  
Dynamic Pressure ◽  
Rocket Engine ◽  
Combustion Instability ◽  
Acoustic Resonator ◽  
Vibration Data ◽  
Engine Design ◽  
Data Processing Method ◽  
Liquid Rocket

Abstract An experimental investigation of high frequency combustion instability in a liquid rocket engine of 3kN thrust was conducted. The diagnostic method of detecting combustion instability was based on the measurement of the dynamic pressure following an artificial disturbance in the combustion chamber. Test data spectral analysis was performed by using the stochastic vibration data processing method. Although the engine has demonstrated an absence of tendency to spontaneous instability, insufficient stability of the original engine design was evident during tests with artificial triggers. A Helmholtz type resonator in form of partitioned cavities tuned to a few different, close frequencies was designed and installed in the combustion chamber wall to avoid spontaneous or triggered combustion instability. The experimental results from hot tests with artificial triggers confirmed the high efficiency of the applied acoustic resonator.

scholarly journals Effect of excitation on gas centered swirl coaxial injectors

2017 ◽  
Author(s):  
Youngbin Yoon ◽  
Gujeong Park ◽  
Sukil Oh ◽  
Jinhyun Bae
Keyword(s):  
Momentum Flux ◽  
Rocket Engine ◽  
Combustion Instability ◽  
Electrical Conductance ◽  
Flux Ratio ◽  
Xenon Lamp ◽  
Momentum Flux Ratio ◽  
Conductance Method ◽  
Liquid Rocket ◽  
Coaxial Injector

Studies on combustion instability in liquid rocket engines are important in improving combustion efficiency andpreventing combustion chamber losses. To prevent combustion instability, methods such as baffles and cavities are used. The injector is located in the middle of the perturbation-propagation process in the rocket engine, so it is important to study the suppression of combustion instability using the design of the injector. Much research has been focused on the study of liquid excitation in a single injector; however, the actual injector used in a liquid rocket engine is a coaxial injector. In this study, the dynamic characteristics of a gas-centred swirl coaxial injector were investigated by varying the gap thickness and momentum-flux ratio. Spray photographs were captured by synchronizing a stroboscope and digital camera, and a high-speed camera and Xenon lamp were also used. To measure the liquid film, a measurement system was implemented using the electrical conductance method. For excitation of the gas, an acoustic speaker was used to impart a frequency to the gas. The gGas velocity and effect of excitation were measured by hot-wire anemometry. A mechanical pulsator was used for liquid flow excitation. Liquid fluctuation was measured by a dynamic pressure sensor. In both gas and liquid excitation cases, the gain increased as the gap thickness decreased and the momentum-flux ratio increased. From these results, it can be concluded that gap thickness and momentum-flux ratio are major factors in suppressing combustioninstability. DOI: http://dx.doi.org/10.4995/ILASS2017.2017.4653

scholarly journals Analysis of Tangential Combustion Instability Modes in a LOX/Kerosene Liquid Rocket Engine Based on OpenFOAM

2022 ◽  
Vol 9 ◽  
Author(s):  
Kangkang Guo ◽  
Boqi Xu ◽  
Yongjie Ren ◽  
Yiheng Tong ◽  
Wansheng Nie
Keyword(s):  
Heat Release ◽  
Standing Wave ◽  
Rocket Engine ◽  
Combustion Instability ◽  
Wave Mode ◽  
Acoustic Mode ◽  
Chamber Pressure ◽  
Liquid Rocket Engine ◽  
Operation Condition ◽  
Liquid Rocket

Self-excited high frequency combustion instability (HFCI) of first-order tangential (1T) mode was observed in a staged-combustion LOX/Kerosene liquid rocket engine numerically. Two different kinds of 1T patterns, standing wave mode and traveling wave mode, were captured in the present work. In the nominal operation condition, the ratio of oxygen-to-fuel (O/F) was 2.5. Propellant was evenly distributed in all injectors and no HFCI occurred. The chamber pressure obtained from the numerical simulation and experiment showed a good agreement, which validated the numerical model. When the mass flow of fuel for two injectors was modified, severe HFCI occurred. The pressure wave node was located at a fixed diameter, showing a 1T standing wave mode. As the O/F was set 4.4 and the propellant distribution was completely uniform, the numerical result yielded a 1T wave node featured a spinning behavior, which was a traveling 1T wave mode. Once the HFCI arose, no matter what standing mode or spinning mode, the pressure and heat release oscillated totally in phase temporally and coupled spatially. The heat release from combustion was fed into the resonant acoustic mode. This was the thermoacoustic coupling process that maintained the HFCI.

Radial Inlet and Exit Design for a 10 MWe sCO2 Axial Turbine

2019 ◽  
Author(s):  
Stefan D. Cich ◽  
J. Jeffrey Moore ◽  
Michael Marshall ◽  
Kevin Hoopes ◽  
Jason Mortzheim ◽  
...  
Keyword(s):  
Power Density ◽  
High Efficiency ◽  
Environmental Changes ◽  
Space Shuttle ◽  
Rocket Engine ◽  
Thermal Mass ◽  
Steam Turbines ◽  
Axial Turbine ◽  
Mechanical Constraints ◽  
Liquid Rocket

Abstract An enabling technology for a successful deployment of the sCO2 closed-loop recompression Brayton cycle is the development of a high temperature turbine not currently available in the marketplace. This turbine was developed under DOE funding for the STEP Pilot Plant development and represents a second generation design of the Sunshot turbine (Moore, et al., 2018). The lower thermal mass and increased power density of the sCO2 cycle, as compared to steam-based systems, enables the development of compact, high-efficiency power blocks that can respond quickly to transient environmental changes and frequent start-up/shut-down operations. The power density of the turbine is significantly greater than traditional steam turbines and is rivaled only by liquid rocket engine turbo pumps, such as those used on the Space Shuttle Main Engines. One key area that presents a design challenge is the radial inlet and exit collector to the axial turbine. Due to the high power density and overall small size of the machine, the available space for this inlet, collectors and transition regions is limited. This paper will take a detailed look at the space constraints and also the balance of aero performance and mechanical constraints in designing optimal flow paths that will improve the overall efficiency of the cycle.

Combustion Test of Oxidizer-Rich Single Triplex Injector Preburner for Staged Combustion Cycle Rocket Engine

2015 ◽  
Author(s):  
Su-Ji Lee ◽  
In-Sang Moon ◽  
Il-Yoon Moon ◽  
Seong-Up Ha
Keyword(s):  
Combustion Chamber ◽  
Dynamic Pressure ◽  
Rocket Engine ◽  
Rocket Engines ◽  
Fuel Injector ◽  
Staged Combustion ◽  
Combustion Pressure ◽  
The Republic ◽  
Combustion Tests ◽  
Combustion Cycle

In the Republic of Korea, research on staged-combustion cycle liquid propellant rocket engines (LPRE) is proceeding to improve efficiency of rocket engines. Recently oxidizer-rich preburner using single triplex injector is developed in relation to the main injector development and combustion tests have been performed. This preburner is designed to operate in nominal conditions with the combustion pressure of 10 MPa, OF ratio of 60. For a stable ignition, LOx is fed in two steps. Triethylaluminum-Triethylborane (TEAB) is used as hypergolic fuel for ignition, supplied through a fuel injector. Despite the small amount of fuel flow rate and high pressure condition, the combustion pressure was stably maintained around 10 MPa as designed. As a result of Fast Fourier Transform (FFT) of the combustion chamber dynamic pressure, 1L mode frequencies related to the acoustic instability and hydraulic resistance exist in the combustion chamber. But their amplitudes are less than 1% of the combustion pressure and it does not affect the combustion. Therefore combustion test is stably completed. In the near future, coupled tests with uni-element triplex injector preburner and uni-element gas/liquid injector main combustion chamber will be carried out.

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