Uncertainty Analysis of the LE-7A Liquid Oxygen-Hydrogen Rocket- Engine Hot Firing Tests

2005 ◽  
Vol 21 (6) ◽  
pp. 1126-1128
Author(s):  
David M. Lineberry ◽  
Hugh W. Coleman ◽  
Ryuichi Sekita
Keyword(s):  
Uncertainty Analysis ◽  
Rocket Engine ◽  
Liquid Oxygen ◽  
Firing Tests

Critical Suction Peformance Test of Turbopump Assembly for a Liquid-Propellant Rocket Engine

2018 ◽  
Author(s):  
Hang Gi Lee ◽  
Ju Hyun Shin ◽  
Suk Hwan Yoon ◽  
Dae Jin Kim ◽  
Jun Hwan Bae ◽  
...  
Keyword(s):  
Liquid Nitrogen ◽  
Performance Test ◽  
Rocket Engine ◽  
Lightweight Design ◽  
Cavitation Number ◽  
Inlet Pressure ◽  
Liquid Oxygen ◽  
Pump Failure ◽  
Critical Cavitation ◽  
Suction Performance

This study investigates the behavior of a turbopump assembly during critical cavitation of the propellant pumps in the upper rocket engine of the Korea Space Launch Vehicle-II. Turbopumps operate under conditions involving low pressure at the pump inlet and high rotational speeds to allow for a lightweight design. This severe environment can easily cause cavitation to occur in the pump. This cavitation can then cause the pump operation to fail. As the cavitation number in the pump decreases below the critical point, the pump fails to operate. There is concern regarding the behavior of the turbopump assembly arising from pump failure due to cavitation. It is necessary to verify the problems that may occur if the turbopump assembly operates under extreme conditions, such like the critical cavitation. This study performed tests to investigate the breakdown of pumps in the turbopump assembly. Tests were conducted with liquid nitrogen, water, and high-pressure air instead of the mediums used during actual operation of liquid oxygen, kerosene, and hot gas. The turbopump was tested at the design point of 27,000 rpm, while the inlet pressure of each pump was controlled to approach the critical cavitation number. The turbine power output was maintained during the tests. The results show that the breakdown point of the oxidizer pump using liquid nitrogen, which is a cryogenic medium, occurred at a lower cavitation number than during an individual component suction performance test using water. The fuel pump using water, meanwhile, experiences breakdown at similar cavitation numbers in both tests. As the breakdown of the pump occurs, the power required by that pump decreases, and the rotational speed of the turbopump increases. Compared with individual pump suction performance tests, this breakdown test can be used to determine the limit of the propellant inlet pressure of the turbopump and to characterize the behavior of the turbopump assembly when a breakdown occurs. Vibrations were also analyzed for tests at a high cavitation number and at the critical cavitation number. The vibration increased with breakdown and notable frequencies were analyzed.

scholarly journals Determination of receiver consumption characteristics using computer simulation

2020 ◽  
Vol 8 (2) ◽  
pp. 10-14
Author(s):  
S.S. Vasyliv ◽  
◽  
V.S. Zhdanov ◽  
M.V. Yevseyenko ◽  
◽  
...  
Keyword(s):  
Computer Simulation ◽  
Flow Rate ◽  
Rocket Engine ◽  
Flow Characteristics ◽  
Maximum Flow ◽  
Complex Configuration ◽  
Rocket Engines ◽  
Gas Leakage ◽  
Control Forces ◽  
Firing Tests

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.

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

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.

Infrared radiation characteristics of liquid oxygen/kerosene rocket engine plume with different number of nozzles

2019 ◽  
Vol 52 (3-4) ◽  
pp. 159-167
Author(s):  
Hong-Hua Cai ◽  
Wan-Sheng Nie ◽  
Ling-Yu Su ◽  
Tian-yi Shi ◽  
Kang-Kang Guo
Keyword(s):  
Infrared Radiation ◽  
Rocket Engine ◽  
Liquid Oxygen ◽  
Radiation Characteristics

Development of a Regeneratively Cooled Low Pressure Liquid Oxygen and Liquid Methane Rocket Engine

2018 ◽  
Author(s):  
Ethan Sichler ◽  
Eric Gonzalez ◽  
Jesus D. Montes ◽  
Richard Picard ◽  
Frank O. Chandler
Keyword(s):  
Rocket Engine ◽  
Low Pressure ◽  
Liquid Oxygen ◽  
Liquid Methane

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

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.

scholarly journals Analytical viscous flow model and test validation of a swirl injector for rocket engine application

2019 ◽  
Author(s):  
G. Fiore ◽  
C. Bach ◽  
J. Sieder ◽  
M. Tajmar
Keyword(s):  
Flow Model ◽  
Rocket Engine ◽  
High Viscosity ◽  
Liquid Oxygen ◽  
Fluid Viscosity ◽  
Prediction Errors ◽  
Viscous Effects ◽  
Swirl Injector ◽  
Close Form

The generally adopted flow model inside a swirl injector, widely used injection concept for propulsive applications, relies upon the hypothesis of ideal flow neglecting the fluid viscosity effects. This model showed significant prediction errors with relatively high viscosity propellants, often leading to the need of an experimental characterization of the injection elements. In this paper, an analytical approach is presented, which includes the effects of viscous diffusion on the injector performance leading to a close form flow solution. The built model is thus experimentally validated testing a liquid oxygen (LOx) and an ethanol injector: the good agreement between the model and the experimental results leads to the construction of the injectors operational maps describing the injector behavior even in the presence of viscous effects.

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