• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.269-273
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• 2005

A practical, 30-tonf-class fullscale thrust chamber has been combustion tested using real propellants for the first time in the domestic technology scene. The very first combustion test was conducted at a low mass flow rate condition for the preliminary assessment of any problems associated with its function and performance while reducing risks from a high chamber pressure never achieved before. A test for the design condition achieved through a low-pressure stage shows stable characteristics of all the static pressures and thrust. Dynamic pressures measured in the manifolds and the chamber did not reveal any distinct wave coupled to a specific frequency and their intensities reside in the allowable range. Moreover, it is encouraging to find no physical failures with a thrust chamber hardware.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.40-46
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• 2006

A 30-tonf-class fullscale thrust chamber for the application to a Low-Earth-Orbit Space Launch Vehicle has been combustion tested over the wide ranges of a mixture ratio and a chamber pressure. The thrust chamber designed for a pump-fed open cycle engine was tested with an ablative chamber instead of a regenerative one for the initial evaluation of its performance and function. The test results revealed stable combustion characteristics. The hardware survived the harsh environment and showed very sound functional characteristics. The measured combustion efficiency turned out to be 95% and a specific impulse at sea level was estimated as 254sec, which are comparable to or above the predetermined design values.

• 유체기계공업학회:학술대회논문집
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• 2006.08a
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• pp.451-454
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• 2006

This paper describes the results of cold flow test and hot firing tests of an uni-element coaxial swirl injector and hot firing tests of a subscale combustor, as to the development effort of coaxial swirl injector for high performance liquid rocket engine combustor. A major design parameter for coaxial swirl injector is the recess number of a bi-swirl injector. The results of hot firing tests of the uni-element injector combustor and the sub-scale combustor are analyzed to investigate the effect of the recess number influencing on the combustion performance and pressure fluctuation. The test results of a cold flow test of the unielement combustor shows that it was shown that the change in recess number has significant effect on mixing characteristics and efficiency, while the effect of recess number on atomization characteristic is not The results of a series of firing tests using unielement and subscale combustor show that the recess length significantly affects the hydraulic characteristics, the combustion efficiency, and the dynamics of the liquid oxygen/kerosene bi-swirl injector. As a point of combustion performance, combustion efficiencies are 90% for unielement combustor and 95% for subscale combustor. The difference in the characteristic velocities between the unielement combustor and the subscale combustor may be caused by the difference in thermal loss to the combustor wall and the relative lengths of the combustion chamber. For a mixed type coaxial swirl combustor, the pressure drop across the injector increases as recess number becomes larger. The low frequency pressure fluctuation observed in unielement combustor can be related to the propellant mixing characteristics of the coaxial bi-swirl injector. The effect of the recess number on the pressure fluctuation inside the combustion chamber is more significant in un i-element combustor than the subscale combustor, of which the phenomena are also observed in time domain and frequency domain.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.1
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• pp.1-8
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• 2013

A liquid-monopropellant hydrazine thruster has several outstanding advantages such as relatively-simple structure, long/stable propellant storability, clean exhaust products, and so on. Therefore hydrazine thruster has such a wide application as orbit and attitude control system (ACS) for space vehicles. A hydrazine thruster with the medium-level thrust to be used in the ACS of space launch vehicles (SLV) has been developed, and its ground firing test result is presented in terms of thrust, impulse bit, temperature, and chamber pressure. It is verified through the performance test that the response and repeatability of thrust are very excellent, and the thrust efficiencies compared to its ideal requirement are larger than 93%.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.2
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• pp.60-71
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• 2022

The reusable, low-cost launch vehicle development trend in the recent launch vehicle market is being subdivided into several ways, and the throttleable engine is one of them. Plus, several nations have selected methane as a next-generation propellant due to its cleanness. In this research, a throttleable pintle injector using gas methane and liquid oxygen as propellants was developed, followed by its spray and combustion characteristics analysis, including high pressure cold and hot tests. The designed throttleable pintle injector has a double sleeve structure, and its tightness and functionality are confirmed through repetitive atmospheric, high-pressure cold tests, and hot tests. Though some design errors were discovered and a low throttling level was unable to be achieved in the combustion test.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.99-103
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• 2017

As a development test of the 75-tonf LOx/Kerosene liquid rocket engine for KSLV-II first Stage Engine, hot firing test of 75-tonf engine are performed. The current status of development test on first stage 75-tonf engine system including combustion chamber, turbopump, gas generator, propellant supply system are presented. During the 75tonf engine test campaign, the development of startup sequence of LOx-Kerosene engine system, engine startup using pyrostarter, ignition of gas generator, steady operation and engine shutdown is successfully performed. As a passenger test during engine hot firing tests, Thrust Vector Control system (TVC) of the engine are also evaluated during engine hot firing test. The results of hot firing test of 75-tonf thrust engine system will be used for the design confirmation and performance evaluation of 75 tonf engine system for KSLV-II first Stage.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.869-872
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• 2011

Ignition tests of an oxidizer rich preburner for a staged combustion cycle liquid rocket engine were performed to evaluate combustion performance. Design operation conditions of the tested oxidizer rich preburner are about 60 of OF ratio and 20 MPa of combustion pressure. The entire kerosene and some LOx injected into the mixing head is burned in combustion chamber and the remaining LOx injected through center holes of combustion chamber is vaporized. Full flow ignition method with hypergolic fuel was used. Each propellant was supplied in two stages for soft ignition. Test results, low frequency oscillation was occurred in low flow rate conditions under 45% of design flow rate. Stable ignition in the course of design combustion pressure was able to induce by minimization of low flow rate ignition region to escape low frequency oscillation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.149-152
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• 2008

Effects of chamber configuration on combustion characteristic velocity of full-scale combustion chamber for 30-tonf-class liquid rocket engine were studied. The configurations of combustion chamber are ablative and channel cooling chamber (${\varepsilon}$=3.2) which have detachable mixing head, and single body regenerative cooling chamber which has nozzle expansion ratio of 3.5 and 12, respectively. The combustion chambers have chamber pressure of 53${\sim}$60 bar and propellant mass flow rate of 89 kg/s, and the injectors of all combustion chamber have recess number 1.0 and double-swirl characteristics. The hot firing test results at design point show that the combustion characteristic velocity of the regenerative cooling chamber which has nozzle expansion ratio of 12 is higher than that of other combustion chambers. The reasons for the above result are the increases of combustion pressure and enthalpy of kerosene which is heated due to cooling of the chamber wall before injection into the combustion field.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.5
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• pp.8-14
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• 2010

For the development of a liquid rocket engine, hot-firing tests of a regeneratively cooled thrust chamber were performed at chamber pressures of approximately 30 and 60 bars. In the paper, pressure fluctuation data, which were obtained from the dynamic pressure transducers installed in propellant manifolds and combustion chamber, were analyzed. Compared to the data at chamber pressure of 60 bar, the results at chamber pressure of 30 bar showed low-frequency oscillations around 150 Hz in the combustion chamber. The low-frequency waves in the combustion chamber were coupled with those in the manifolds. However, the RMS values of the chamber pressure fluctuations at chamber pressure of 30 bar were only 0.8% of the chamber pressures. Thus, it can be inferred that the thrust chamber operates in the stability boundary even at low chamber pressure.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.7
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• pp.91-97
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• 2004

The first Korean liquid propellant rocket KSR-III successfully finished its flight test on Nov. 28, 2002. Herein, we summarize the results of range safety system operation which is employed for the first time in flight tests of rockets developed by Korea Aerospace Research Institute(KARI). During the flight, safety-critical flight data including instantaneous impact points are monitored in realtime by range safety officers utilizing Range Safety Display Systems. The recorded screen of the display system is presented for the explanation of safety operation. In addition, comparisons are made between onboard navigation system based and radar based results in calculating instantaneous impact points, and also errors from the finally recorded impact point are described.