Pre-cooled Turbojet Research Articles

Effect of Number and Spacing of Pre-cooled Turbojet Engines on Takeoff Noise of Hypersonic Transport

Effect of Number and Spacing of Pre-cooled Turbojet Engines on Takeoff Noise of Hypersonic Transport

Effect of Equivalence Ratio on Acoustic Characteristics in Pre-cooled Turbojet Engine Afterburner

Effect of Equivalence Ratio on Acoustic Characteristics in Pre-cooled Turbojet Engine Afterburner

Acoustic Characteristics of Rectangular Hypersonic Mixer Ejector

JAXA is currently studying the pre-cooled turbojet engine (PCTJ) to be installed in a hypersonic transport. As a device to reduce exhaust noise at takeoff, a mixer-ejector that is applicable to a rectangular hypersonic nozzle having a complicated geometry with an external ramp has been proposed and tested. The ejector introduces the low-speed secondary flow into the high-speed main jet and reduces the exhaust velocity by mixing. As a result, the mean jet velocity decreases and the jet noise decreases. A 1%-scale mixer-ejector was fabricated, and the mixing process between the main jet and the secondary flow was visualized using Mie scattering of water droplets. The direct relationship between the degree of mixing and the amount of noise reduction was investigated.

Performance analysis of precooled turbojet engine with a low-temperature endothermic fuel

For hydrocarbon-fueled TBCC engines, precooling is a crucial technology to solve “thrust gap” by expanding the flight Mach number of the turbojet engine, which can bring a sharp reduction of the specific impulse due to insufficient heat sink of hydrocarbon fuel. A low-temperature endothermic fuel precooling scheme is a great potential way to solve this problem. Heat sink experiments and thermodynamic cycle analysis are used to verify the effectiveness of the low-temperature endothermic fuel precooling scheme in improving specific impulse performance. The experiment results indicated that low-temperature endothermic fuel has a more considerable heat sink at lower temperatures than traditional endothermic hydrocarbon fuel. The results of thermodynamic cycle analysis show that the low-temperature endothermic fuel precooling scheme can effectively improve the specific impulse. Maximum specific impulse can reach above 800s with the low-temperature endothermic fuel precooling scheme, which is below 400s with hydrocarbon fuel precooling at Ma0 = 3.0. When cracking rate equals 1.0, and precooling efficiency equals 0.85, maximum specific impulse can reach 600s, increased by approximately 1.4 times at Ma0 = 4.0. The feasibility of precooled turbojet with the low-temperature endothermic fuel precooling scheme was preliminarily verified through this work.

Mach 4 Simulating Experiment of Pre-Cooled Turbojet Engine Using Liquid Hydrogen

This study investigated a pre-cooled turbojet engine for a Mach 5 class hypersonic transport aircraft. The engine was demonstrated under takeoff and Mach 2 flight conditions, and a Mach 5 propulsion wind tunnel test is planned. The engine is composed of a pre-cooler, a core engine, and an afterburner. The engine was tested under simulated Mach 4 conditions using an air supply facility. High-temperature air under high pressure was supplied to the engine components through an airflow control valve and an orifice flow meter, and liquid hydrogen was supplied to the pre-cooler and the core engine. The results confirmed that the starting sequence of the engine components was effective under simulated Mach 4 conditions using liquid hydrogen fuel. The pre-cooling effect caused no damage to the rotating parts of the core engine in the experiment.

3-Dimensional Numerical Simulation of Hypersonic Intake for Pre-Cooled Turbo Jet Engine

3-Dimensional Numerical Simulation of Hypersonic Intake for Pre-Cooled Turbo Jet Engine

ノズル形態が極超音速航空機の離陸騒音に及ぼす影響

Effect of nozzle configuration on the takeoff noise of a hypersonic transport is investigated based on small-scale nozzle experiments. In the Japan Aerospace Exploration Agency, the precooled turbojet (PCTJ) engine aimed at the hypersonic flight is under development. The takeoff noise of the full-scale vehicle is assessed before it is put into the operation using an acoustic simulation method. By matching the jet velocity and jet Mach number with the realistic PCTJ engine condition, the acoustic characteristics of the hypersonic nozzle are duplicated with a small-scale nozzle model. The jet noise database around the hypersonic nozzle is established through a number of experiments, and by taking into account 1) the nozzle scale, 2) the number of nozzles, 3) the direction of installation, 4) the measurement distance and angle based on the flight trajectory, and 5) the atmospheric absorption, the effective perceived noise level (EPNL) at the flyover point is calculated. It is shown that the nozzle configuration has a large impact on the takeoff noise, and that such an assessment could contribute to the improvement of the integration design of the vehicle.

予冷ターボジェットエンジンのアフターバーナにおける燃焼効率の計測

予冷ターボジェットエンジンのアフターバーナにおける燃焼効率の計測

スリット噴射が矩形極超音速ノズルからの排気騒音に及ぼす影響

Effects of a slit injection (SI) on exhaust noise from a rectangular hypersonic nozzle were investigated experimentally. In the Japan Aerospace Exploration Agency (JAXA), the pre-cooled turbojet (PCTJ) engine specified for a hypersonic transport is under development. In this study, a 1/100-scaled model of the rectangular hypersonic nozzle for the PCTJ is manufactured. The main jet (MJ) is injected to the atmosphere through the nozzle. The total pressure of the MJ was set at the takeoff condition of the vehicle. From slit shaped converging nozzles mounted at the end of the ramp, air is injected to the MJ. The sound pressure level (SPL) around the nozzle was measured with microphones. Using the SI, especially in the ramp direction, the SPL was decreased by 1 to 2dB at wide frequencies up to 50kHz. At the same time, the Overall SPL (OASPL) decreased monotonously with the increase in the momentum ratio of the SI and the MJ. It is considered that the governing parameter of the SI is the momentum ratio. The growth rate of the shear layer was estimated from Pitot pressure distribution. It is implied that the shear layer growth rate is linked to the acoustic radiation.

Thrust Measurement of a Rectangular Hypersonic Nozzle Using an Inclined Baffle Plate

Thrust measurement of a rectangular hypersonic nozzle employed in the precooled turbojet engine under development in the JapanAerospace Exploration Agency is carried out using an inclined baffle plate. A 1.0%-scaled model of the nozzle was manufactured, and its gross thrust was investigated. The hypersonic nozzle has a variable throat and an external ramp to compensate the drastic change in the nozzle pressure ratio through the operation. At takeoff, an intensive jet noise is expected due to the high-speed exhaust jet, and in the present study, the impact of aerodynamic-tab jet noise suppressors on the thrust vector of the nozzle is investigated. Because of the simplicity, a baffle plate, which basically provides an on-axis forcemeasurement, is applied to a simultaneousmeasurement of the magnitude andangle of the thrust vector by giving anangle to the impact surface relative to the jet. It is shown that the magnitude of the thrust vector is kept constant regardless of the aerodynamic-tabmass fraction, which implies that a lossless mixing between the main jet and aerodynamic tab can be assumed. It is also shown that the thrust angle changes with the increase in the aerodynamic-tab mass fraction because of the deformation of the jet cross section.

Feasibility of Aerodynamic-Tab Jet Noise Suppressors in a Hypersonic Nozzle at Takeoff

I N THE Japan Aerospace Exploration Agency (JAXA), a precooled turbojet engine specified for hypersonic propulsion is under development [1,2]. This engine uses a precooling cycle using liquid hydrogen fuel, and it is capable of accelerating a vehicle from takeoff toMach 5. It is planned that the vehicle takes off horizontally using existing runways and that the exhaust velocity is supersonic throughout the operation. Thus, jet noise suppression at takeoff is one of the most important issues in the development. At takeoff, the nozzle pressure ratio (NPR) is limited to 2.7 since no ram pressure can be expected in such an airbreather, while at cruise, NPR reaches 200 due to the increase in flight Mach number and altitude [2]. To compensate such drastic change in NPR, a rectangular hypersonic nozzlewith a variable throat realizing the change in throat area by 2.5 times will be employed [2]. However, it is still insufficient to satisfy all operation conditions of the engine, and the nozzle will inevitably be operating under offdesign conditions (overexpansion mode) at takeoff. Because of the complicated geometry and the offdesign operation, it seems to be challenging to apply jet noise suppressors to such an actual hypersonic nozzle. Tabs attached at the nozzle exit are expected to be a promising device for screech and broadband noise reduction in a supersonic jet, and their noise reduction performance and mechanism have been investigated in detail [3–5]. Recently, a new concept of jet noise suppressors using auxiliary gas injection has been proposed [6–8], which achieves similar effect to tabs aerodynamically. Both in mechanical-tab and aerodynamic-tab (AT) concepts, much research is carried out using round/rectangular nozzles that are farther idealized when compared with actual hypersonic nozzles. Furthermore, in theAT concept, bleed air will be used as its working gas, and the total pressure and temperature available in real engines should be taken into account. In the present study, noise reduction performance and feasibility of AT jet noise suppressors in an actual hypersonic nozzle operating under takeoff conditions are investigated experimentally.

Observation of Flame Stabilized at a Hydrogen-Turbojet-Engine Injector Installed into a Lab-Scale Combustion Wind Tunnel

A lab-scale combustion wind tunnel was developed for investigation of low-pressure ignition and flame holding in a sub-scale pre-cooled turbojet engine with hydrogen fuel in order to make engine start at high altitudes sure. The combustion wind tunnel is a blow-down type. A fuel injector of the sub-scale pre-cooled turbojet engine was installed into the combustion wind tunnel. Conditions in which a flame can be stabilized at the fuel injector were examined. The combustor pressure and equivalence ratio were varied from 10 to 40 kPa and from 0.4 to 0.8, respectively. The mean inlet air velocity was varied from 2 to 48 m/s. Flames stabilized at 20 kPa in pressure and 0.6 in equivalence ratio were observed. It was found that the decrease in the combustor pressure narrows the mean inlet air velocity range for successful flame holdings. Flame holding at lower combustor pressures is realized at the equivalence ratio of 0.4 in the low mean inlet air velocity range, and at the equivalence ratio of 0.6 in the high mean inlet air velocity range. Flame luminosity is the largest near the fuel injector. The flame luminosity distribution becomes flatter as the increase in the mean inlet air velocity.

Research on hypersonic aircraft using pre-cooled turbojet engines

Systems analysis of a Mach 5 class hypersonic aircraft is performed. The aircraft can fly across the Pacific Ocean in 2h. A multidisciplinary optimization program for aerodynamics, structure, propulsion, and trajectory is used in the analysis. The result of each element model is improved using higher accuracy analysis tools. The aerodynamic performance of the hypersonic aircraft is examined through hypersonic wind tunnel tests. A thermal management system based on the data of the wind tunnel tests is proposed. A pre-cooled turbojet engine is adopted as the propulsion system for the hypersonic aircraft. The engine can be operated continuously from take-off to Mach 5. This engine uses a pre-cooling cycle using cryogenic liquid hydrogen. The high temperature inlet air of hypersonic flight would be cooled by the same liquid hydrogen used as fuel. The engine is tested under sea level static conditions. The engine is installed on a flight test vehicle. Both liquid hydrogen fuel and gaseous hydrogen fuel are supplied to the engine from a tank and cylinders installed within the vehicle. The designed operation of major components of the engine is confirmed. A large amount of liquid hydrogen is supplied to the pre-cooler in order to make its performance sufficient for Mach 5 flight. Thus, fuel rich combustion is adopted at the afterburner. The experiments are carried out under the conditions that the engine is mounted upon an experimental airframe with both set up either horizontally or vertically. As a result, the operating procedure of the pre-cooled turbojet engine is demonstrated.

Development study of a precooled turbojet engine

A precooled turbojet engine has been developed by JAXA used for the hypersonic airplane and spaceplane. The subscale engine named “S-engine” whose thrust and weight are about 1.2 kN and 100 kg was designed, fabricated and tested. The components and the system firing tests under the sea-level-static condition were successfully conducted.In the next phase, a flight test of the S-engine is planned using a stratospheric balloon in 2010 called balloon-based operation vehicle (BOV). The vehicle is dropped from an altitude of 40 km by a high altitude balloon. After 40 s free-fall, the vehicle is pulled up and the S-engine operates for 30 s at about Mach 2. High-altitude tests of the core-engine verified the performance and healthiness of the engine under the condition corresponding to the BOV flight trajectory.

Conceptual Study on Hypersonic Turbojet Experimental Vehicle (HYTEX)

Pre-cooled turbojet engines have been investigated aiming at realization of reusable space transportation systems and hypersonic airplanes. Evaluation methods of these engine performances have been established based on ground tests. There are some plans on the demonstration of hypersonic propulsion systems. JAXA focused on hypersonic propulsion systems as a key technology of hypersonic transport airplane. Demonstrations of Mach 5 class hypersonic technologies are stated as a development target at 2025 in the long term vision. In this study, systems analyses of hypersonic turbojet experiment (HYTEX) with Mach 5 flight capability is performed. Aerodynamic coefficients are obtained by CFD analyses and wind tunnel tests. Small Pre-cooled turbojet is fabricated and tested using liquid hydrogen as fuel. As a result, characteristics of the baseline vehicle shape is clarified, . and effects of pre-cooling are confirmed at the firing test.

S0505-3-2 空力タブノズル形状が矩形プラグノズルからのジェット騒音に及ぼす騒音低減効果(車両・機体・流体機械に関連した騒音問題へのアプローチと対策(3))

Effects of aerodynamic tab geometry on exhaust noise from a rectangular plug nozzle were investigated experimentally. A 1/100-scaled model of the rectangular plug nozzle for the pre-cooled turbojet engine planed in JAXA is manufactured and the exhaust noise characteristics were investigated. Take off condition of the vehicle were simulated. The nozzle pressure ratio is 2.7. Aerodynamic tabs were used as a noise reduction device. There are three types of aerodynamic tab geometries, namely, 4 kinds of round nozzles, 2 kinds of wedge nozzles and a slit nozzle. Sound pressure measurements were carried out. It was shown that by use of wedge and slit nozzles, the aerodynamic tab mass flow rate, necessary to gain 2.0 dB reduction in OASPL, decreases by 16 % when compared with round nozzles.

巡航マッハ数に応じた超音速ジェットエンジンの多目的最適化

The aim of this paper is to apply a multi-objective optimization generic algorithm (MOGA) to the conceptual design of the hypersonic/supersonic vehicles with different cruise Mach number. The pre-cooled turbojet engine is employed as a propulsion system and some engine parameters such as the precooler size, compressor size, compression ratio and fuel type are varied in the analysis. The result shows that the optimum cruise Mach number is about 4 if hydrogen fuel is used. Methane fuel instead of hydrogen reduces the vehicle gross weight by 33% in case of the Mach 2 vehicle.

予冷ターボジェットエンジンからのジェット騒音に及ぼすノズルスケール，全温およびアフターバーナの影響

Effects of nozzle scale, total temperature, and an afterburner on jet noise characteristics from a pre-cooled turbojet engine are investigated experimentally. In JAXA (Japan Aerospace Exploration Agency), a pre-cooled turbojet engine for an HST (Hypersonic transport) is under development. In the present study, 1.0%- and 2.4%-scaled models of the rectangular plug nozzle (Nozzles I and II) are manufactured, and the jet noise characteristics are investigated under a wide range of total temperatures. For Nozzle I, no air-heater is utilized and the total temperature is 290K. For Nozzle II, a pebble heater and an afterburner (AB) are utilized upstream of the nozzle model, and the total temperature is varied from 520K (pebble heater) to 1540K (pebble heater + AB). The total pressure is set at 0.27 and 0.30MPa(a) for both nozzle models. Jet noise is measured using a high-frequency microphone set at 135 deg from the engine inlet, and normalized jet noise spectra are obtained based on AUjn law and Helmholtz number. For cases without afterburner, the normalized spectra agrees well regardless of the nozzle scale and total temperature where the velocity index lies from n = 7.7 to 9.2, and the correlation factor between the two facilities is shown to be about 1dB. For the case with afterburner, the normalized spectrum does not agree with other conditions where the velocity index n seems to be about 4.

矩形ならびに軸対称プラグノズルから放出される超音速ジェット騒音の予測

Effects of nozzle geometry, and total temperature on supersonic jet noise radiated from rectangular and axi-symmetric plug nozzles are investigated, experimentally. In JAXA (Japan Aerospace Exploration Agency), a pre-cooled turbojet engine for an HST (Hypersonic transport) is under development. In the present study, three kinds of subscale nozzle models are employed, namely two kinds of rectangular plug nozzles (RPN1 and RPN2) and an axi-symmetric plug nozzle (APN), and the jet noise data are acquired at aft angles of the jets by use of 1/4 inch high frequency microphones. The total pressure is set at 0.3MPa(a), which corresponds to the take-off condition of the vehicle, and the total temperature is varied from 290K to 860K. The jet noise spectra obtained are reduced to normalized spectra by use of a scaling law of heated jets (AU n law). It is shown that the normalized spectra collapse onto two lines according to each nozzle geometry, regardless of the total temperature. For APN, the peak SPL is smaller by about 8 to 14 dB when compared with that for RPN1 and RPN2, which implies that the axi-symmetric plug nozzle could be much quieter than rectangular plug nozzle.

矩形プラグノズルからの排気騒音に及ぼす空力タブの効果

Effects of aerodynamic tabs on exhaust noise from a rectangular plug nozzle were investigated experimentally. In JAXA (Japan Aerospace Explanation Agency), a pre-cooled turbojet engine for the 1st stage propulsion of a TSTO (Two stage to orbit) is planned. In the present study, a 1/100-scaled model of the rectangular plug nozzle for the pre-cooled turbojet engine is manufactured and the exhaust noise characteristics were investigated. Compressed air is injected through the rectangular plug nozzle into the atmosphere. The nozzle pressure ratio was set at 2.7, which corresponds to the take-off condition of the vehicle. Aerodynamic tabs were installed at the ramp end (Upper AT), the cowl end (Lower AT) and the sidewall end (Side AT). The SPL (Sound pressure level) was measured with a high-frequency microphone. Without AT, the sound spectrum has a broadband peak at which the SPL is around 105dB. For Lower and Side ATs, the OASPL (Overall SPL) of the exhaust noise decreases, especially around ramp end. At the maximum, the OASPL was reduced by 2.4dB with about 2% loss of the main jet total pressure at the cowl exit. It is shown that the aerodynamic tabs are effective in noise reduction in a rectangular plug nozzle.