Turbo Ramjet Research Articles

Design of thrust augmentation control schedule during mode transition for turbo-ramjet engine

A thrust augmentation control schedule was designed for turbo-ramjet engine during mode transition process to achieve high thrust performance. First, a mathematical model of turbo-ramjet engine was established, analyzing the reason for the existence of an insufficient thrust working range under conventional mode transition control schedule. On this basis, a thrust augmentation control schedule was devised for the thrust insufficient range considering the matching between the air inlet available airflow and the engine demand airflow. The control schedule could increase the flow capacity of the engine and reduce the overflow drag of air inlet by opening the ramjet bypass while keeping the turbofan components working at the maximum state, which could improve the maximum installation thrust performance of the engine and achieve a better air inlet/engine matching. By further designing the optimal regulation rule for the rear variable area bypass injector (RVABI), the installed thrust of the engine could be increased by more than 37.93% at most, while simultaneously reducing the fuel consumption by 0.16%. Finally, the turbo-ramjet engine was simulated along the isobaric flight path. The simulation results revealed that under conventional mode transition control schedule, the thrust insufficient occurred when Mach number is above 2.7. The designed thrust augmentation control schedule could effectively improve the maximum thrust of the thrust drop zone and provide sufficient thrust margin for engine during the whole working process. At the same time, considering the smooth thrust and thrust augmentation requirements, a compromise mode transition control schedule between the thrust augmentation control schedule and the conventional mode transition schedule is proposed. This control schedule could address the problem of insufficient thrust at high Mach numbers while the smoothness transition of engine total thrust could be guaranteed.

Overspeeding characteristics of turbomachinery for gas generator cycle air turbo ramjet engine

The main interest in the current study focuses on the possibility of overspeeding for the gas-generator cycle air turbo ramjet (GG-ATR) engine. The authors developed the air turbo ramjet engine and investigated its compressor performance. Based on those data, the authors developed the analytical code for the air turbo ramjet engine, which calculates the performances of turbomachinery, gas-generator, and ram combustor. The previous study described that the rotor overspeeding would not occur in the air turbo rocket engine. However, the current results show that degraded ram combustion can decrease the compressor pressure ratio and the compressor power. This reduced compressor power can cause overspeeding for the air turbo ramjet engine. The experimental results of compressor power and turbine inlet pressure support those analytical results.

Low toxic nitromethane based monopropellant for gas generator cycle air turbo ramjet engine

The present study proposes ethanol-nitromethane mixture as a safe, storable, and low toxic monopropellant for a gas generator cycle air turbo ramjet engine and evaluates its propulsion performance. The present study proposes that nitromethane is mixed with ethanol to adjust gas generator combustion temperature. The author developed the computational code for the present analysis and verified its accuracy. The specific thrust of ethanol-nitromethane is nearly identical to that of ethanol-liquified oxygen and is higher than methanol-hydroxylammonium nitrate aqueous solution and hydrazine. Hydrazine has the highest Isp among the propellants in the present analysis. However, Isp of ethanol-nitromethane is nearly equal to that of ethanol-liquified oxygen and higher than those of methanol-hydroxylammonium nitrate aqueous solution, and hydroxyl terminated polybutadiene-ammonium perchlorate. With the propellant-to-air ratio range in the present study, ethanol-nitromethane has the stoichiometric condition at the ram combustor around the propellant-to-air ratio range from 0.2 to 0.25 and can obtain high ram combustion temperature. This result is favorable for ethanol-nitromethane because it improves the specific thrust and the specific impulse. Therefore, ethanol-nitromethane can be a promising low-toxic liquid monopropellant for the air turbo ramjet engine.

Закономірності зміни оптимальних складу і параметрів силових установок з прямоточними контурами від швидкості надзвукового крейсерського польоту

There are two aircraft concepts for transoceanic flights at supersonic speed. The first is the aircraft with moderate supersonic cruising speed M = 1.7...2.2, and the second is the aircraft with cruising speed M ≥ 3. Propulsion with a ramjet engine (ramjet) is preferable for high supersonic speeds. However, a ramjet has no starting thrust and is uneconomical at subsonic flight speeds. Combined propulsion with a turbojet duct for takeoff and a ramjet duct for supersonic cruise flight are used to overcome this contradiction. There are known propulsion with turboramjet engines, as well as various propulsion schemes with turbofan engines (afterburning and duct-burning ones), in which the outer contour together with the afterburner and nozzle can be considered the ramjet duct. When the turbojet duct is turned off, the operating process of such turbofan engines is practically the same as the ramjet, which allows using the advantages of a ramjet in supersonic cruise flight. Because flight at supersonic cruising speed can be provided by different propulsion compositions, the choosing problem of the composition and parameters of the propulsion for an aircraft with supersonic cruising speed is relevant. A calculation-analytical method was used to select the composition and parameters of the propulsion, which is based on the maximum relative mass of the payload criterion. Using the aircraft mass balance equation this criterion can be represented as a minimum conditions for the relative mass of the fuel and propulsion. Using this method, for an aircraft with a given mass, aerodynamic characteristics and a given flight profile, for the indicated propulsion compositions, the change patterns of the minimum relative mass of fuel and propulsion in the range of M = 2.5...4 are established. The established patterns allow choosing the propulsion composition and parameters depending on the speed of the supersonic cruising flight or choosing the flight speed of the aircraft at a given propulsion composition. The dependance of the propulsion working process optimal parameters for takeoff and cruising mode on the speed of supersonic cruising flight according to the criterion minimum of relative mass of fuel and propulsion were obtained.

Вибір складу, параметрів робочого процесу і режиму роботи силової установки літального апарата з надзвуковою крейсерською швидкістю польоту

The issues of choosing a composition, operating process parameters and operating modes of propulsion for supersonic cruising aircraft providing transoceanic flights are considered. A condition of maximum payload relative mass is used as a criterion for choosing the composition of the propulsion. This criterion can be transformed using the aircraft mass balance equation into a condition of the minimum relative mass of fuel and propulsion. A predictor-corrector method is used to solve the task. Choosing the composition, operating process parameters and operating modes of propulsion according to the cruising segment, taking into account the remaining flight segments takeoff, climb and descent by the empirical coefficients, is used as a predictor stage. Based on the solving results of this stage, the best competing variants of propulsions are selected for the purpose of subsequent more detailed analysis at a corrector stage by numerically solving the differential equations of the aircraft motion along an entire flight profile. At the same time, on each elementary section of the path, the control parameters of the propulsion are optimized according to the criterion of the minimum required fuel consumption to overcome this section. Propulsion with turbojet engine, turbofan and mixed flows turbofan engine, turbo-ramjet engine and turbofan engines that have ramjet modes are considered. Patterns of the change in the relative mass of the fuel and propulsion for the indicated compositions of the propulsions according to cruising segment of the flight for the cruising speeds corresponding to the M = 1.5...4 are established. These dependencies make it possible to select variants of propulsion competing in terms of payload for a cruising speed or select the most advantageous cruising speed for propulsion composition which is given. These dependencies were used to determine competing variants of the power plant, providing a minimum of the relative mass of fuel and propulsion for the airplane with a cruising speed corresponding to the M = 3. At the stage of the corrector, these variants of the propulsion were evaluated according to the criterion of the relative mass of fuel and propulsion by solving the equations of the aircraft motion along the entire flight profile with the optimization of the control parameters of the propulsion on each elementary segment of the profile according to the criterion of minimum fuel consumption to overcome it. In doing so, the operating process parameters of the propulsion were optimized near of those values that were obtained at the stage of the predictor. The analysis of the obtained results indicates that, in comparison with the predictor stage at the corrector stage, the parameters of the propulsion operating process that are optimal from the point of view of the relative mass of fuel and propulsion changed by about 12.5 %, and the relative mass of fuel and propulsion by about 3...4 %.

S4 - A demonstrator of HYPLANE, a single stage suborbital spaceplane and a hypersonic business jet

The mid-to-long term perspectives of suborbital hypersonic spaceplanes require the availability of ultra-fast, winged vehicles, characterized by low wing loading, streamlined fuselages, sharp nose and wing leading edges, able to manoeuvre along flight trajectories at small angles of attack. From another point of view, the cost associated with sub-orbital space access flight is strongly conditioned by the still small dimension and uncertainties of its market.Therefore, hypersonic technologies and vehicle design suitable for the suborbital market as well as for point-to-point fast transportation can facilitate the endeavour. A high-speed business-jet is widely recognized to be more technologically affordable than larger aircraft and even more feasible from an economic point of view. Along this wave, HYPLANE is proposed as a 6-pax hypersonic business jet designed for P2P stratospheric transportation and suborbital flight at reduced ticket price. It will be based on more readily state-of-art available technologies like titanium alloys for its structure which are suitable to manage the levels of heat produced on external surfaces at those speeds and turboramjet engines, instead of less proven scramjets required for faster aircrafts. No disruptive or currently low TRL technology is needed, rather enhancing already available innovations. Attention is given to maintaining high level of environmental protection, especially in terms of sonic boom effects on ground, global and local pollution and noise emissions. To flight validate the system and related technologies, a multi-mission Single Stage Suborbital Spaceplane (S4) demonstrator is proposed. In parallel, the evolution of the Campanian Grazzanise airport to become a spaceport is being supported in the frame of a Suborbital TEst Polygon (STEP) proposed to be established between Campania and Sardinia regions. It will make available the perfect operational scenario where to test and make use of the S4 demonstrator.

Alcohol and alkane fuel performance for Gas Generator cycle Air Turbo Ramjet Engine

The present study investigates the feasibility of alcohol fuels for a Gas Generator cycle Air Turbo Ramjet (GG-ATR) Engine in comparison with alkane fuels. The alkane and alcohol fuels in the present study range from C1 species (CH4 and CH3OH) to C4 species (n-C4H10 and 1-C4H9OH).The present study analytically evaluates Specific thrust and Specific impulse (Isp) of GG-ATR engine as functions of gas generator combustion temperature. For both alcohols and alkanes, the peak Gas Generator (GG) combustion temperatures exist for specific thrust and ram combustor temperature. Comparing alcohol to alkane in the lower GG combustion temperature, specific thrusts and Isp of alcohol tend to be higher than those of C2H6, C3H8, and 1-C4H10. GG combustion gases for alcohol fuels contain higher mole fractions of H2O and CO2 than those for alkanes. Because of high mole fractions of H2O and CO2, the ratio of GG combustion gas for alcohols to air flow rate is fortunately close to the stoichiometric ratio, resulting in higher ram combustion temperature. The high ram combustion temperature can contribute to the higher specific thrust and Isp. Lower GG combustion temperature is preferable because of a thermal limit of a turbine blade. Thus, alcohols have superiorities to alkanes for the GG-ATR engine application.

Advantage of ethanol fuel for gas generator cycle air turbo ramjet engine

Performance analyses of Gas Generator cycle Air Turbo Ramjet (GG-ATR) engine are conducted to investigate the feasibility of ethanol fuel, as comparing Liquefied Hydrogen (LH2), Liquefied Natural Gas (LNG), and n-C12H26. Liquefied Oxygen (LOX) is considered as an oxidizer. Gas generator combustion temperature and compressor pressure ratio are selected as the analytical parameters, which are varied from 900 to 1700 K and from 2.0 to 4.0, respectively.In general, specific impulses of light molecule weight fuels, such as LH2 and LNG, are higher than those of large molecule weight fuel. For storable fuel, Isp of ethanol is larger than that of n-C12H26 when gas generator temperature is less than 1300 K, although specific heat release of ethanol is only two thirds of n-C12H26 heat release. The main chemical species of n-C12H26 gas generator (GG) combustion gas are CH4 and CO. On the other hand, ethanol GG combustion gas is mainly composed of H2 and CO, resulting in smaller molecular weight than n-C12H26. These characteristics contribute to larger Isp of ethanol than n-C12H26 and indicate the feasibility of ethanol for GG-ATR engine fuel.

NUMERICAL SIMULATION OF INLET FOR TURBINE ASSISTED RAMJET

The ramjet engine performs faster than a Turbojet Engine at higher Mach numbers, but the engine can’t start with the own power itself. The starting thrust must be attained by additional thrust producing devices like solid rocket boosters. The more reasonable solution to overcome the starting problem is the integration of ramjet engine with turbojet engine, which leads to the Turboramjet. This work describes some challenges in the design of an inlet for turbo-ramjet. Since it is a critical component, its design has important effects on the overall performance of the engine. Here this work deals with a design of a turbo ramjet inlet and an analysis for its optimization. Our aim is to reduce total pressure loss and entropy rise. A design of turbo-ramjet intake is made with CATIA as well as a complete engine for better understanding. Gambit is used to mesh the geometry while ANSYS 14 is used to perform the analysis for design optimization.

Goal Programming for Stable Mode Transition in Tandem Turbo-ramjet Engines

This article, in order to guarantee the stable mode transition in tandem turbo-ramjet engines, investigates the multi-objective and multi-variable goal programming algorithm. First, it introduces the structural features of the variable cycle turbo-ramjet engines, the principles of selecting the mode transition operation point and the design parameters, and the characteristics of the turbofan mode and the ramjet mode. Second, a component-based variable cycle turbo-ramjet engine model is developed to simulate the mode transition process. Third, the Newton-Raphson algorithm is used to solve the multi-variable and multi-objective optimization problem. The results show that with the maximum residua of only 0.06%, this algorithm has an acceptable convergence that meets the predetermined goals. Finally, the simulation shows that the stable turbo-ramjet mode transition could be realized with the mode transition control law developed by the algorithm.

4029 エルボ下流に配置した千鳥配列円管群を通過する流れ(G05-6 物体まわりの流れ(1),G05 流体工学)

The Air Turbo Ramjet (ATR) engine is considered to be one of the propulsion systems for space planes. The precooler is one of the heat exchangers and a key technology for the ATR engine. It produces, however, the inlet distortions of total pressure. The tube banks with staggered arrangement are located just downstream of an elbow. The velocity distributions are measured in the downstream of the tube banks. For comparison, the tube banks with straight duct are also examined. The results show that the low spot of an arranging in A-shape in the elbow duct with tube banks. Moreover, a low speed area appears by backward side in the straight duct with yawed tube banks.

Carbon-carbon composite turbine disk for the air turbo ramjet engine (ATREX)

A feasibility study of three-dimensionally fiber-reinforced carbon-carbon composites (3DC/Cs) for application to a turbine disk of ATREX (Air turbo ramjet engine with expander cycle) was carried out. Spin burst tests at room temperature were conducted using 3D-C/C disks, and the fracture behaviors were characterized. A 3D-C/C disk was totally fractured at a peripheral speed of 516 m/s (r = 150 mm), which is sufficient for the ATREX application. However, fiber bundles at the disk periphery prematurely suffered micro-scale damage, and fragments of the fiber bundle unit flew out before total fracture occurred. In order to prevent the fly-out behavior, the disk was impregnated with Si only near its periphery. Although this treatment increased the initiation speed of the flyout behavior, this improvement was considered insufficient for purposes of the ATREX application. Next, a simplified analysis was conducted to characterize the fly-out behavior. Based on this analysis, the following three measures were discussed: (1) decreasing bundle thickness (i.e. using fine fiber texture), (2) increasing toughness of the fiber bundle interface, and (3) minimizing local curvature in waviness of the fiber bundles in the circumferential direction.

タンデム翼列によるファンの高負荷化

For axial flow compressors and fans in the aircraft engines higher pressure ratio is required in order to attain the high thrust engines. In this study, the fan with the tandem cascades was introduced to increase the fan pressure ratio. The use of tandem cascades in the fan allows savings in length and weight and therefore a compact fan could be built. The design of fan with tandem cascades and the fan testing were carried out to develop the high pressure ratio fan for the Air Turbo Ramjet (ATR) propulsion system. The ATR is a combined cycle engine which performs like a turbojet engine at subsonic speeds and a ramjet at supersonic speeds. In particular, high fan pressure ratio contributes to increase the engine thrust during subsonic flight at which the engine does not make use of ram effect. The results of the fan testing indicate that the pressure ratio of 2.2 is achieved in single stage fan.

Integrated Vehicle Comparison of Turbo-Ramjet Engine and Pulsed Detonation Engine

The pulsed detonation engine (PDE) is a unique propulsion system that uses the pressure rise associated with detonations to efficiently provide thrust. A study was conducted under the direction of the NASA Langley Research Center to identify the flight applications that provide the greatest potential benefits when incorporating a PDE propulsion system. The study was conducted in three phases. The first two phases progressively screened a large matrix of possible applications down to three applications for a more in-depth, advanced design analysis. The three applications best suited to the PDE were (1) a supersonic tactical aircraft, (2) a supersonic strike missile, and (3) a hypersonic single-stage-to-orbit (SSTO) vehicle. The supersonic tactical aircraft is the focus of this paper. The supersonic, tactical aircraft is envisioned as a Mach 3.5 high-altitude reconnaissance aircraft with possible strike capability. The high speed was selected based on the perceived high-speed fuel efficiency benefits of the PDE. Relative to a turbo-ramjet powered vehicle, the study identified an 11% to 21% takeoff gross weight (TOGW) benefit to the PDE on the baseline 700 n.mi. radius mission depending on the assumptions used for PDE performance and mission requirements. The TOGW benefits predicted were a result of the PDE lower cruise specific fuel consumption (SFC) and lower vehicle supersonic drag. The lower vehicle drag resulted from better aft vehicle shaping, which was a result of better distribution of the PDE cross-sectional area. The reduction in TOGW and fuel usage produced an estimated 4% reduction in life cycle cost for the PDE vehicle. The study also showed that the simplicity of the PDE enables concurrent engineering development of the vehicle and engine.

Computational Analysis of the Flow Field Near the Boat-tail Region of Annular Plug Nozzles.

Flow fields over annular plug nozzles are computationally simulated and the boat-tail drag characteristics are clarified and discussed based on the simulation results. The plug nozzle configuration is taken from the ATREX (Air Turbo Ramjet) engine under development at the ISAS. The simulations are carried out for the axi-symmetric nozzle configuration using compressible Navier-Stokes equations. The computed result shows that there exists a low pressure region downstream of the Prandtl-Meyer expansion on the boat-tail region and the separation shock wave appears due to the interaction of the main flow and the nozzle plume downstream. The result also shows that the primary cause of the boat-tail drag turns out to be the low pressure region downstream of the Prandtl-Meyer expansion. The flow field with the secondary flow injection is then simulated and the computed result shows that the pressure in the region in front of the separation shock wave increases and the secondary flow injection successfully reduces the boat-tail drag. The effect of the turbulence model used in the computation is discussed to ensure the reliability of the solutions.

小型・高推力エアーターボラムエンジン

The Air Turbo Ramjet (ATR) is a combined cycle engine which performs like a turbojet engine at subsonic speeds and a ramjet at supersonic speeds and therefore the ATR is an attractive propulsion system for the wide operation range (e.g. Mach 0 to Mach 4). The ATR can provide a higher specific impulse than a solid fuel rocket engine and a higher thrust per frontal area than a turbojet engine. The major ATR components are the inlet, fan (compressor), turbine, gas generator, combustor and exhaust nozzle. In the ATR, the turbine drive gas is generated by a decomposed liquid or solid fuel gas generator. In order to carry heavier payloads and to attain shorter flight time, the compact and high thrust engine is required. In this study, the ram combustor with the double-staged flameholders and the fan with tandem blade were introduced to shorten the engine length and to increase the fan pressure ratio, respectively. Furthermore, the engine testing was carried out on sea level static condition to confirm the engine component integration technologies for the ATR propulsion system.

Development Study on ATREX Engine

This is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Institute of Space and Astronautical Science (ISAS) cooperation with the Ishikawajima Harima Heavy Industries (IHI), the Kawasaki Heavy Industries (KHI), the Mitsubishi Heavy Industries (MHI). ATREX engine will be applied for the propulsion system of fly-back booster of TSTO space plane. ATREX is the combined cycle (a fan-boosted ramjet) engine providing the effective thrust from sea level static to flight Mach number 6. ATREX is worked on the expander cycle with precooling the incoming air as shown in Fig. 1 . ATREX employs the tip turbine configuration which allows the compactness and the light weight of turbo machinery and the variable geometry airintake and plugnozzle which allow the wide range operation conditions.From 1990 to 1992, “ ATREX-500“ has been tested at the sea level static conditions. ATREX-500 is the 1/4-scale model of which fan inlet diameter is 300mm and overall length 2,200mm. From 1992 have been performed the wind tunnel tests on the primary components of ATREX, the axisymmetric variable geometry airintakes, the precoolers and the variable geometry plug nozzles. In parallel to the windtunnel tests, the ram combusters have been tested simulating the hypersonic flight conditions and the application studies on advanced carbon-carbon composite for the tip-turbine and fan assembly has been proceeded.In 1994 initiated the flight test plan in which ATREX will be verified in the practical flight conditions by using an unmanned flying test bench.In 1995 will be tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing on the precooler.The present paper addresses the high loading ram combuster experiment using the mixer with skewed lobes to generate swirl flow and the analytical studies and the designs on the precooler and the precooled ATREX engine and the flight test plan.

Development study on ATREX engine

This is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Instutute of Space and Astronautical Science (ISAS) in cooperation with the Ishikawajima Harima Heavy Industries (IHI), the Kawasaki Heavy Industries (KHI), the Mitsubishi Heavy Industries (MHI). ATREX engine will be applied for the propulsion system of fly-back booster of TSTO space plane. ATREX is the fan-boosted ramjet producing the effective thrust from sea level static to flight Mach number 6. ATREX is worked on the expander cycle with precooling incoming air as shown in Fig.1. ATREX employs the tip turbine configuration which allows compactness and light weight of turbo machinery and the variable geometry airintake and plug nozzle which allows the wide range flight conditions. ATREX development study has been conducted with the sea level static tests since 1990. ATREX engine for tests is the scaled model designated by “ATREX-500” of which fan inlet diameter is 300 mm and overall length 2,200 mm. From 1992 have been performed the wind tunnel tests on the primary aerodynamic components such as the axisymmetric variable geometry air intakes, the precoolers and the variable geometry plug nozzles. The application study on advanced carbon-carbon composite for the tip-turbine and fan assembly has been conducted. This study status is presented in the another session of this IAF congress 1. The flight test of ATREX is now planning to verify the engine performance and functions in the practical flight conditions by using an unmanned flying test bench. In 1995 was tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing problem on the precooler. The present paper addresses primarily the test results of precooled ATREX engine.

96/04309 The basis for a method to integrate work environment in life cycle assessments

This is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Institute of Space and Astronautical Science (ISAS) cooperation with the Ishikawajima Harima Heavy Industries (IHI), the Kawasaki Heavy Industries (KHI), the Mitsubishi Heavy Industries (MHI). ATREX engine will be applied for the propulsion system of fly-back booster of TSTO space plane. ATREX is the combined cycle (a fan-boosted ramjet) engine providing the effective thrust from sea level static to flight Mach number 6. ATREX is worked on the expander cycle with precooling the incoming air as shown in Fig. 1. ATREX employs the tip turbine configuration which allows the compactness and the light weight of turbo machinery and the variable geometry airintake and plugnozzle which allow the wide range operation conditions.From 1990 to 1992, “ ATREX-500“ has been tested at the sea level static conditions. ATREX-500 is the 1/4-scale model of which fan inlet diameter is 300 mm and overall length 2,200 mm. From 1992 have been performed the wind tunnel tests on the primary components of ATREX, the axisymmetric variable geometry airintakes, the precoolers and the variable geometry plug nozzles. In parallel to the windtunnel tests, the ram combusters have been tested simulating the hypersonic flight conditions and the application studies on advanced carbon-carbon composite for the tip-turbine and fan assembly has been proceeded.In 1994 initiated the flight test plan in which ATREX will be verified in the practical flight conditions by using an unmanned flying test bench.In 1995 will be tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing on the precooler.The present paper addresses the high loading ram combuster experiment using the mixer with skewed lobes to generate swirl flow and the analytical studies and the designs on the precooler and the precooled ATREX engine and the flight test plan.

96/04311 Biological mine drainage treatment

This is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Institute of Space and Astronautical Science (ISAS) cooperation with the Ishikawajima Harima Heavy Industries (IHI), the Kawasaki Heavy Industries (KHI), the Mitsubishi Heavy Industries (MHI). ATREX engine will be applied for the propulsion system of fly-back booster of TSTO space plane. ATREX is the combined cycle (a fan-boosted ramjet) engine providing the effective thrust from sea level static to flight Mach number 6. ATREX is worked on the expander cycle with precooling the incoming air as shown in Fig. 1. ATREX employs the tip turbine configuration which allows the compactness and the light weight of turbo machinery and the variable geometry airintake and plugnozzle which allow the wide range operation conditions.From 1990 to 1992, “ ATREX-500“ has been tested at the sea level static conditions. ATREX-500 is the 1/4-scale model of which fan inlet diameter is 300 mm and overall length 2,200 mm. From 1992 have been performed the wind tunnel tests on the primary components of ATREX, the axisymmetric variable geometry airintakes, the precoolers and the variable geometry plug nozzles. In parallel to the windtunnel tests, the ram combusters have been tested simulating the hypersonic flight conditions and the application studies on advanced carbon-carbon composite for the tip-turbine and fan assembly has been proceeded.In 1994 initiated the flight test plan in which ATREX will be verified in the practical flight conditions by using an unmanned flying test bench.In 1995 will be tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing on the precooler.The present paper addresses the high loading ram combuster experiment using the mixer with skewed lobes to generate swirl flow and the analytical studies and the designs on the precooler and the precooled ATREX engine and the flight test plan.