Soo Seok Yang

Cowl and cavity effects on mixing and combustion in scramjet engines

To investigate the supersonic combustion patterns in scramjet engines, a model scramjet engine was tested in the T4 free-piston shock tunnel. The test model had a rectangular intake, which compressed the freestream flow through a series of four shock waves upstream of the combustor entrance. A cavity flame holder was installed in the supersonic combustor to improve ignition. The freestream test condition was fixed at Mach 7.6, at an altitude of 31 km. This experimental study investigated the effects of varying fuel equivalence ratios, the influence of the cavity flame holder, and the effects of cowl shape. As a result, supersonic combustion was observed at equivalence ratios between 0.11 and 0.18. Measurements indicated that the engine thermally choked at a fuel equivalence ratio of 0.40. Furthermore, the cavity flame holder and the W-shaped cowl showed improved pressure distribution due to greater reaction intensity. With the aid of numerical analysis, the cavity and the W-shaped cowl are shown to be effective in fuel–air mixing.

Effects of Flameholder Configurations on Combustion in Scramjet Engines

The effects of flameholder configurations on combustion within a scramjet combustor are investigated through experiments, numerical simulation, andquasi-one-dimensional analysis. For the experimental test, a blowdownwind tunnel with a vitiated air heater is used. The test model is composed of an isolator, fuel injectors, and a diverging combustor. To investigate the effects of different flameholder configurations, three kinds of exchangeable fuelinjection plates are fabricated: no cavity, plain cavity, and zigzag cavity. The test results indicate that the zigzag cavity cases obtain the highest values of combustion-induced pressure among the three. The results of the numerical analysis show that the zigzag cavity generates a transverse-directional flow and enhances combustion. The zigzag cavity case also shows the highest combustion efficiency in the quasi-one-dimensional analysis. With a zigzag cavity installed, a scramjet engine combustor can be shortened, and its friction drag can be lowered.

Development of the Scramjet Engine Test Facility in Korea Aerospace Research Institute

A hypersonic air-breathing engine test facility was constructed at Korea Aerospace Research Institute for a hypersonic engine research. This facility, so called Scramjet engine test facility(SeTF) is a blow-down type, high enthalpy wind tunnel which has the pressurized air supply system, air heating system, free-jet type test chamber, supply system, facility control, and measurement system, and exhaust system. In this paper, Details of the specifications and configuration of the Scramjet engine test facility are described, and the performance data of the SeTF are also included.

Improvement of a Model Scramjet Engine Design Using Various Types of Ground Tests

The scramjet, one of the most promising candidates for future transport systems, has many advantages such as a simple configuration and a high specific impulse. However, various characteristics of scramjet engines in a hypersonic environment are difficult to predict from theory. In the present study, we investigated the characteristics and the performance of the scramjet engine with various types of ground tests. An engine model (S1) was tested in the free-piston shock tunnel T4. (S1) showed active supersonic combustion and good starting characteristics, but its operating range is rather limited. To improve the performance of the model scramjet engine, we also investigated major components, such as the combustor and the intake by experiments with different model configurations. In the combustor test results, the zigzag cavity flame holder showed the best combustion performance among the three cases. By testing various intake configurations, the most efficient intake configuration was also determined. A revised model (S2), which was redesigned based on the component test results, showed enhanced thrust performance and a wider operating range in the free jet test using the free piston shock tunnel HIEST.

Scramjet Engine Combustor Tests in a Supersonic Wind Tunnel with a Vitiated Air Heater

Effects of flame holder configurations on the combustion within the scramjet combustor are investigated by experiments and quasi-one-dimensional analysis. For the test, a blowdown windtunnel with vitiated air heater is used. The test model is composed of an isolator, fuel injectors and a diverging combustor. In order to investigate the effects of flame holder configurations, three kinds of fuel injection plates, “No cavity”, “Plain Cavity” and “zigzag cavity” are made for exchange. As a result, zigzag cavity showed the highest value of combustion induced pressure and estimated combustion efficiency among others. The zigzag cavity is expected to generate transverse directional pressure nonuniformity and enhance the fuel-air mixing due to irregularity of flame front. With application of zigzag cavity, scramjet engine combustor can be shortened and attain lower friction drag.

Study for Starting Characteristics of Scramjet Engine Test Facility (SETF)

Korea Aerospace Research Institute started on design and development of a hypersonic air-breathing engine test facility from 2000 and completed the test facility installation in July 2009. This facility, designated as the Scramjet engine test facility (SETF), is a blow-down type high enthalpy wind tunnel which has a pressurized air supply system, air heater system, free-jet type test chamber, fuel supply system, facility control/measurement system, and exhaust system with an air ejection. Unlike most aerodynamic wind-tunnel, SETF should simulate the enthalpy condition at a flight condition. To attain a flight condition, a highly stagnated air comes into the test cell through a supersonic nozzle. Also, an air ejector of the SETF is used for simulating altitude conditions of the engine, and facility starting. SETF has a storage air heater (SAH) type heating system. This SAH can supply a hot air with a maximum temperature of 1300K. Using the SAH, SETF can achieve the Mach 5.0 flight at an altitude of 20 km condition. SETF has a free-jet type test cell and this free-jet type test cell can simulate a boundary layer effect between an airplane and engine using the facility nozzle, but it is too difficult to predict the nature of the facility. Therefore it is required to understand the starting characteristics of the facility by experiments. In 2009, a Mach 3.5 test of SETF was done for acceptance testing which is a maximum air supply condition of 20 kg/s. SETF showed the facility efficiency of a 100% without a test model at the Mach 3.5 condition. In 2010, a Mach 6.7 aerodynamic test campaign with a scramjet engine intake. But SETF could not start at the Mach 6.7 condition with the existing ejector system at that time. To get a facility starting, we modified the ejector system. After modification of the ejector system, SETF started at the Mach 6.7 condition with a facility efficiency of 58%. In this paper, the starting characteristics of the SETF with various flight conditions, and modifications of the ejector system will be described.Copyright

Multi Stage Axial Compressor Design and Performance Evaluation

Recently, needs for Unmanned Air Vehicle (UAV) and small aircraft are increasing and demands for small turbo jet or turbo fan engines are also increasing. Then, size and weight are the two main restrictions in UAV or small aircraft propulsion system applications. One method for resolving such a problem is to increase the pressure rise per stage and to reduce the number of stages. Nowadays, matured compressor aerodynamic design techniques enable us to design highly loaded axial compressors. This paper covers from the design step of a highly loaded transonic axial compressor to the performance test result and its analysis. At the fore part of the paper, aerodynamic process of a multi stage axial compressor is introduced. To satisfy both of the mass flow and pressure rise, the compressor should rotate at a high rotational speed. Therefore the transonic flow field forms in the rotor stages and it is designed with a relatively high pressure rise per stage to satisfy its design target. Basically, one dimensional and quasi three dimensional compressor design were carried with compressor design codes. The compressor stage consists of 3 stages, and the bulk pressure ratio is 2.5. The first stage is burdened with the highest pressure ratio and less pressure rises occur in the following stages. Also it is designed that tip Mach number of the first rotor row does not exceed 1.3. The final design was confirmed by iterating three dimensional CFD calculations to satisfy design target and some design intentions. In the latter part of the paper, its performance test processes are briefly introduced. The performance test result showed that the overall compressor performance targets; pressure ratio and efficiency are well achieved. From the test results, we found some clues for further improvement and optimization of the compressor aerodynamic performance.Copyright