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The Korea Aerospace Research Institute runs the ground-based Altitude Engine Test Facility (AETF) for simulating high-altitude environments. This facility is widely used to test the performance of gas turbine engines at ground level to high altitudes. When engine performance is tested in the AETF, total pressure and total temperature of airflow can be measured using rakes in a duct directly connected to the engine inlet. The total pressure and total temperature of airflow into an engine are the main factors affecting the thrust and specific fuel consumption rate of the engine. Furthermore, the total pressure and total temperature at the engine inlet are used as the main reference values when operating test equipment to determine the airflow rate and temperature corresponding to the performance test conditions of the engine. The performance evaluation of the rake used to measure this is crucial. Accordingly, in this study, to improve the accuracy of the total pressure and total temperature values measured using rakes, their total pressure coefficient and total temperature recovery factor are measured, and the measurement uncertainty of the total temperature recovery factor is evaluated.

A turbo-compression system design and its performance analysis procedure for a high altitude long endurance UAV (HALE UAV), of which cruising altitude is within the stratosphere, is presented. To fly at a relatively low speed for a long time and to make engine performance less sensitive to flight altitude, a hydrogen fueled internal combustion engine was chosen for a propulsion system. To utilize an internal combustion engine as a propulsion system at a high altitude, a proper inlet pressure boost system such as a series of turbochargers is required. Hydrogen is highly reactive gas and sometimes backfiring or preignition may occur due to its low ignition energy at stoichiometric ratio. Therefore, fuel to air ratio should be reduced as low as 0.6 to avoid such phenomena. Then rarefied ambient intake air pressure should be boosted up to 1.7 bar to produce required power from the lean burn engine. To gain high pressure ratio from the turbo compression system, at least three stage serial turbocharger with proper intercooler system at each compressor exhaust is required. To analyze multi-stage turbocharger performance at the cruising altitude, an explicit one-dimensional analysis method has been established mainly by matching required power between compressors and turbines. Each compressor performances were corrected according to Reynolds number at a given flight altitude. Compressor efficiency and surge margin deteriorate as the operating altitude increases. Then compressor efficiencies were reflected as functions of flight altitude and corresponding Reynolds number. Once operating points of each turbocharger was determined, then adequate turbochargers were searched for from commercially available models based on performance analysis results. Also, adequate water to air intercoolers were chosen for the turbo-compression system to secure flexibility of placing main components inside the engine bay as well as to obtain high heat exchange efficiency of the heat exchangers. Based on the designed turbo-compression system, technical demonstration test of the turbo-compression system inside altitude environment test chamber in Korea Aerospace Research Institute is planned. Altitude condition in stratosphere is simulated mainly with two stage centrifugal compressor and additional fan will be used to fine control the flight altitude. The turbo compression system will be controlled with a single waste gate located just downstream of the engine to secure simple controllability of the turbo compression system. The test results will validate main components as well as system layout design methods and give more reliable control schedule of the turbo compressions system according to the flight altitude.

S INCE the concept of a supersonic-combustion ramjet, i.e. scramjet, was introduced, it has always been considered to be a candidate power plant for a hypersonic vehicle for use as a civil transport or a portion of a space launch system. The classical concept of a scramjet engine envisions it to use hydrogen as the fuel and to operate highly efficiently in Mach numbers of 6 and above [1–3], producing a large flight range as well as reaching the destination fast. In practice, many problems have been found with the hydrogenburning scramjet idea: to achieve high fuel efficiency, the combustor must operate at a high supersonic Mach number. Such a high combustor Mach number prevents fuel from mixing with air. The volume of the tank holding hydrogen becomes very large. The skin temperature becomes so high that no existingmaterial could sustain a cruising condition. Gaseous hydrocarbon fuels, such as ethylene or methane, have also been investigated [4,5]. These fuels need tanks that are much smaller than those needed for hydrogen. But these fuels must still be cryogenically stored. For this reason, kerosene has been studied extensively as an alternate fuel for the past several decades [6,7]. Kerosene is a liquid at room temperature and requires a much smaller tank volume than that required for hydrogen. Although ignition and flame holding become difficult for a liquid fuel, the problems are solved by applying advanced fuel atomizing techniques [7] or using additives [8]. Kerosene, as well as other hydrocarbon fuels, can be used for vehicles having a flight Mach number of only 3.5 to 5. Flying at such a relatively lowMach number, the combustor entranceMach number becomes low supersonic. At such a low supersonic Mach number, fuel–air mixing will improve greatly over a vehicle flying at Mach 6 or higher. Besides, the skin temperature becomes so much lower. Heat flux is proportional to the square of flight speed, and therefore there is a large difference between theMach 4 andMach 6 vehicles in the skin temperature. In this study, a kerosene-burning scramjet engine model is designed and tested to obtain an overall insight into such an engine. A blowdown, free-jet-type continuous facility is used. A scramjet engine, in an integrated form of an intake, isolator, and combustor, is used as the test model. Ignition and fuel atomization parameters as well as the fuel equivalence ratio and the geometry parameter such as combustor length are changed for respective experiments. Through this experiment, some aspects of the relationship between the scramjet engine configuration and the operational characteristics such as pressure rise, thermal choking, supersonic-to-subsonic transition, and intake unstarts are discovered.

In this study, the effects of the diameter and quantity of fuel nozzles combined with various cavity-type flame holders for supersonic combustion were investigated empirically and numerically. Although increasing the quantity of fuel nozzles yields a greater fuel-to-air interface area, which results in a higher fuel–air mixing rate, smaller nozzle diameters yield a lower penetration depth, which results in a lower combustion performance at the top wall region. This study adopts the cavity shapes of fuel injectors from previous research studies to design new injector configurations with different sizes and quantities of fuel nozzles. The diameter and quantity of nozzles are determined to have the same equivalence ratio and momentum ratio as were used in previous research under a given fuel pressure. A comparison of results from the present tests with previous research shows that the plain cavity and zigzag cavity have distinct characteristics. Configurations using these cavity types were investigated via t...

This paper introduces a failure analysis procedure that underpins real-time fault prognosis. In the previous study, we developed a systematic eventization procedure which makes it possible to reduce the original data size into a manageable one in the form of event logs and eventually to extract failure patterns efficiently from the reduced data. Failure patterns are then extracted in the form of event sequences by sequence-mining algorithms, (e.g. FP-Tree algorithm). Extracted patterns are stored in a failure pattern library, and eventually, we use the stored failure pattern information to predict potential failures. The two practical case studies (marine diesel engine and SIRIUS-II car engine) provide empirical support for the performance of the proposed failure analysis procedure. This procedure can be easily extended for wide application fields of failure analysis such as vehicle and machine diagnostics. Furthermore, it can be applied to human health monitoring & prognosis, so that human body signals could be efficiently analyzed.

A Mach 5 nozzle and a diffuser of the Scramjet Engine Test Facility (SETF) were made for a hydrocarbon-fueled scramjet engine. SETF, attached with a diffuser guide, started with a model of 60% blockage, though the model engine could not start by over expansion of the facility nozzle. The model was moved into the nozzle to escape the shock generated from the nozzle exit, both SETF and the engine could start. The pitot rake experiments (blockage of 2.3%) were done for measuring the core flow in the test section. From the pitot experiments, the core flow was expanded by an under expansion. It means that the core flow in the test section was related with a model blockage. SETF and the engine with a blockage of 33% work normally. From a series of experiments, SETF started with a normal shock efficiency of 58%, regardless of a blockage ratio.

Combustion tests were conducted for a hydrogen-fueled Mach 5 scramjet engine model using a blow-down facility. No fuel and two fuel flow rate cases were tested for two different model configurations. Time history of the wall static pressures inside the model and their time-averaged spatial distribution were used for the analysis of the flow and combustion characteristics. For shorter model, supersonic combustion was occurred for both of the fuel flow rate cases. For longer model, supersonic combustion was occurred for less fuel case, whereas thermal choking and subsonic combustion were occurred for more fuel case. Intake started even for this subsonic combustion case.

O PERATION of a scramjet engine involves many gas dynamic and combustion phenomena. The components of a scramjet engine, such as the intake, isolator, fuel injector, and flame holder, interact with each other in a complicated manner and may result in successful engine operation or fall into intake unstart, misfire, blowout, or thermal choking. Much of the previous experimental research on scramjets focuses on a single component, and examples of experimental research on the integrated intake, isolator, and combustor are fewer.Moreover, much research uses impact facilities, i.e., shock tubes or shock tunnels, and sometimes they cannot represent real operational conditions of the scramjet engine. For an example, some research shows that the scramjet intake starts more easily in impact facilities than in continuous facilities [1]. In this study, the effect of the combustor configuration on the flow and combustion in a scramjet engine is explored experimentally. A blowdown freejet-type continuous facility is used.A scramjet engine, in an integrated form of an intake, isolator, and combustor, is used as the test model. Fuel injector, cavity flame holder, and combustor configurations are changed for respective experiments. Through this experiment, some aspects of the relationships between the scramjet engine injector, cavity, and combustor configuration and the operational characteristics such as intake starts and unstarts, supersonic combustion, and combustion-driven unstarts are brought to light.

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 re...

This paper aims to develop an event-driven failure analysis and prognosis system that is able to monitor ship status in real time, and efficiently react unforeseen system failures. In general, huge amount of recorded sensor data must be effectively interpreted for failure analysis, but unfortunately noise and redundant information in the gathered sensor data are obstacles to a successful analysis. This paper therefore applies `Equal-frequency binning` and `Entropy` techniques to extract only important information from the raw sensor data while minimizing information loss. The efficiency of the developed failure analysis system is demonstrated with the collected sensor data from a marine diesel engine.

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.

High performance sensors and modern data logging technology with real-time telemetry facilitate system fault diagnosis in a very precise manner. Fault detection, isolation and identification in fault diagnosis systems are typical steps to analyze the root cause of failures. This systematic failure analysis provides not only useful clues to rectify the abnormal behaviors of a system, but also key information to redesign the current system for retrofit. The main barriers to effective failure analysis are: (i) the gathered data (event) logs are too large in general, and further (ii) they usually contain noise and redundant data that make precise analysis difficult. This paper therefore applies suitable pre-processing techniques to data reduction and feature extraction, and then converts the reduced data log into a new format of event sequence information. Finally the event sequence information is decoded to investigate the correlation between specific event patterns and various system faults. The efficiency of the developed pre-filtering procedure is examined with a terminal box data log of a marine diesel engine.

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.

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 © 2012 by ASME

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.