小林 弘明

In this paper are presented the advanced concept of air inlet applying for the air breathing propulsion systems of hypersonic flight vehicle like a space plane. It is designated to be the Multi-Row Disk (MRD) inlet It was devised to give the better inlet performance (mass capture ratio (MCR) and total pressure recovery (TPR)) over the wide flight conditions up to around Mach 6 by improving its off-design performance. The MRD inlet is a kind of axisymmetric air inlet composed of the center conical envelope spike and the cowl. As shown in Fig. 1, the unique structure is employed on the conical envelope spike that is composed of a tip cone and several round disks arranged so as to shape the conical envelope. The space in between tìie disks can be changed mechanically and thus the overall spike length of the MRD inlet is adjustable to meet the shock on lip condition at the given flight speed independent of the throat area control. The performance of MCR and TPR are governed respectively by the shock on lip condition to reduce a spillage flow and the throat area control and thus the MRD inlet is easier to improve them independent from each other under the off design conditions. It was made clear from the result of wind tunnel tests that the MRD inlet achieves the same on-design performance as the conventional inlet with the solid surface conical spike and improves TPR by 10% comparing with the conventional ones in case of flow matching with the turbo fan. It was revealed from the result of wind tunnel tests and numerical simulations that the cavities formed in between the disks gave little effect on a boundary layer growth. © 2004 by the American Institute of Aeronautics and Astronautics. Inc.

This paper proposes a new supersonic inlet, which is called Multi-Row Disk (MRD) inlet The MRD inlet has a centerbody being composed of a tip cone and following disks. Because centerbody geometry is variable by changing space between disks, air compression ratio and mass capture ratio of this inlet can be controlled independently of each other. However, conical cavity flow formed between disks may affect the boundary layer profile negatively resulting in the deteriorated performance of inlets. We made a basic research on conical cavity flow and validation tests of the MRD inlet in 2002 and 2003. There was significant improvement in the off-design point performance of the MRD inlet relative to general axisymmetnc inlets.

The way to demonstrate the ATREX engine in

A nozzle of the ATREX Engine must have variable geometry to change pressure ratio from 3 (SLS) to 550 (Mach 6). An axisymmetric plug nozzle will be employed for the ATREX Engine. Throat area of the axisymmetric plug nozzle can be controlled by moving plug back and forth in the axial direction. The plug nozzle shows relatively higher performance than conventional C-D nozzle under any ambient pressure. The objective of this study is to estimate thrust efficiency and boat tail drag of the plug nozzle. Several types of the plug nozzle were tested in a supersonic wind tunnel. Injecting secondary flow thorough the cowl was devised and shown to be effective in the reduction of boat tail drag.

Axisymmetric and three dimensional CFD simulations were performed to predict and improve internal and external performance of supersonic inlet of ATREX engine. Aerodynamic performances such as total pressure recovery, mass capture ratio and intake external drag were in good agreement with the experimental results over a Mach number range from 1.5 to 5.5 and over a angle of attack from 0 deg to 4 deg. However, because the CFD result estimates the flow separation caused by the subsonic diffuser, the static pressure of the CFD result was lower than the experimental data. Furthermore, some ideas to improve the performances are shown, that is to increase the bleed flow rate of the spike and decrease the area gradient of the subsonic diffuser.

This paper describes the study of frost formation on the precooler tube surfaces. Frosting rates are quantitatively shown by experimental and numerical methods, which depends on the temperature and humidity of the air flow as well as the cooling wall temperature. The effectiveness of a method to improve the precooler performance under frosting condition was investigated by experiments using a sub-scale heat exchanger model. Addition of a methanol proved to be most effective compared with other possible substances in both cases of lower and higher cooling wall temperature. Then the effectiveness of the methanol addition was ascertained for the practical condition that means the same tube configuration and flow velocity as the precooler designed for the ATREX engine firing test model. The result showed that the addition of the same quantity as the water vapor could restrain the frost layer from choking the flow in the duration of 300 seconds, which is sufficient time for precooler operation. The required methanol mass along the ATREX engine flight path was estimated to be less than 3 % of fuel hydrogen on board. Accordingly, the method came to be promising candidate for practical application.

Here is presented an experimental and analytical study on a precooler for hypersonic air-breathing engines. Precooling of the incoming air breathed by an air-inlet gives extension of the flight envelope and improvement of the thrust and specific impulse. Three precooler models were installed into an air-turbo ramjet engine and tested under the sea level static condition. When the fan inlet temperature was down to 160K, the engine thrust and specific impulse increased by 2.6 and 1.3 times respectively. parametric studies on the precooler design values and a sizing analysis were also performed, Decrease of tube outer diameter on the precooler is only way to increase heat exchange rates without increase of its weight and pressure loss.

The present paper addresses the development study of the air turbo ramjet engine with expander cycle (ATREX) being conducted since 1986 in the Institute of Space and Astronautical Science in cooperation with the industries (IHI, KHI, MHI and SHI). The ATREX is expected as one of the most promising candidates for the propulsion system of a future space plane. The ATREX is the combined cycle engine performing like a turbojet at lower flight speed and a fan-boosted ramjet at higher flight speed beyond Mach 3 to 6. The 1/4-scaled model of ATREX whose fan inlet diameter of 300mm was built for system verification under sea level static conditions. Carious components such as turbo-fan, precooler, mixer, regeneratively cooled combustor and heat exchanger as well as the expander cycle have been developed and verified in the firing test. Number of 63 tests with 3,300 sec of the total duration have been conducted step by step since 1990 at Noshiro Testing Center of ISAS. The recent study focused on the precooler. The latest model (Type-III) was designed taking into consideration the reduction of size and weight as well as the heat-dynamic performance aiming at the flight model/ Engine performance is improved such that the Thrust and specific impulse are increased 1.8 and 1.2 times respectively by reducing the fan inlet temperature to 180K. The heat exchange performance of precooler was 80-90 % of the design value and the large pressure losses of the air flow occurred due to the frost formation on the tube surfaces. This frost formation phenomenon has been made clear analytically and experimentally and several methods eliminating it have been devised and tested.

The flight of spaceplane is always under accelarating in the assent way and always under decelerating in the desent way and yet cruising in the return way. Besides, its flight envelope is considerably wider than that of airplane. Thus the integrated design method is required to build the best transportation system optimized taking into account the propulsion system and the airframe under the entire flight conditions. In this paper its shown an optimization method on TSTO spaceplane system. Genetic algorithm (GA) was applied to optimize design parameters of engine, airframe, and trajectory simultaneously. Several types of engine were quantitatively compared using payload ratio as an evaluating function. From a Viewpoint of the relation between performance and weight, it was concluded that the precooled turbojet is the most promising engine for TSTO among Turbine Based Combined Cycle (TBCC) engines.

A flight-type ATREX engine consists of many components; inlet, precooler, fan, turbine, combustor, nozzle, and so on. These components have been developed by a number of sea-level combustion tests and wind tunnel tests in ISAS. In this paper, a detailed characteristic model of each component is described. These characteristic models are integrated into a flight-type ATREX engine simulator.

In this paper, specification for TSTO spaceplane powered by airbreathing engines is presented. A Multi-criteria Trade-off Analysis (MTA) of Turbine Based Combined Cycle (TBCC) engines was made to estimate the performance of an flight-type engine under the current level of technology. The cycle analysis has been performed after all the component models are reviewed. The flight-type ATREX engine can meet all demands from the TSTO spaceplane but Isp. The following calculation results showed the possibility of ATREX engine providing higher Isp by improving the performance of some components.

The ATREX engine has been developed by ISAS for the propulsion system of a fly-back booster of a future TSTO space plane. Summary of the engine system, current R&D status and plans on the ATREX are presented in this paper, Several studies on system optimization, precooled expander cycle, thermal-fluid performance, control and composite materials have been performed by the system firing tests and wind tunnel tests, etc. As the next step, we propose a development plan in the next decade, in which a half-scale prototype engine with the flight weight will be produced for the demonstration flight test.

The flight of Spaceplane is always under accelarating in the assent way and always under decelarating in the desent way and yet cruising in the return way. Besides, its flight envelope is considerably wider than that of airplane. Thus the integrated design method is required to build the best transportation system optimized taking into account the propulsion system and the airframe under the entire flight conditions. In this paper it is shown an optimization method on TSTO spaceplane system. Genetic algorithm (GA) was applied to optimize design parameters of engine, airframe, and trajectory simultaneously. Several types of engine were quantitatively compared using payload ratio as an evaluating function. It was concluded that precooled turbojets is the most promising engine for TSTO among Turbine Based Combined Cycle (TBCC) engines.

A feasibility of ATREX (Air-Turbo-Ram Expander cycle) engine with conventional aft-turbine configuration has been studied to be developed in about 10 years, if the development project has started under enough resources. The novel tip-turbine of the original ATREX engine is replaced by a conventional aft-turbine, and the maximum turbine inlet temperature (TIT) is reduced to 1200K, to realize the engine by only using approved metal technologies of modern jet engines. The capability of the performance has been shown by parametric studies by changing components' design parameters. The study shows that the performance of the ATREX engine is not less than that of pre-cooled turbo jet. Some technical issues on developing the new ATREX engine have been addressed. The most important issue would come from the transient total temperature change due to the rapid acceleration from sea level static (SLS) condition (288K) to Mach 6 at 30km of altitude (1680K) in 6 minutes. The deformation due to transient thermal expansion has to be controlled. Especially, the change of the tip clearance and the clearance between rotors and stators are pointed out to be important design issues. The ATREX engine, which has shorter axial length and simpler rotor, has structural advantage over turbo jet. © 2002 International Astronautical Federation. Published by Elsevier Science Ltd. All rights reserved.

Here is presented an experimental and analytical study on a precooler for hypersonic air-breathing engines. Precooling of the incoming air breathed by an air-inlet gives extension of the flight envelope and improvement of the thrust and specific impulse. Three precooler models were installed into an air-turbo ramjet engine and tested under the sea level static condition. When the fan inlet temperature was down to 180K, the engine thrust and specific impulse increased by 2.0 and 1.2 times respectively. Thick frost formed on the tube surfaces at the entrance part of the precooler blocked the air-flow passage. On the other hand, very thin frost formed at the exit part because the water vapor included in the air was changed to mist particles due to the low temperature of the air in this part. Parametric studies on the precooler design values and a sizing analysis were also performed. Decrease of tube outer diameters on the precooler is only way to increase heat exchange rates without increase of its weight and pressure loss.

Air intake is one of the most important components for an airbreathing propulsion system of supersonic and hypersonic vehicles. Air intake can be evaluated by air mass capture ratio and total pressure recovery ratio. In higher Mach number flight condition, larger total pressure losses occurs in the compression processes of air intake and reduces the propulsion performance. By utilizing the precompression coming from oblique shocks generated underneath vehicle forebody, a part of functions loaded in air intake can be substituted by the forebody precompression, thereby overall propulsive performance is able to be improved effectively. In the present paper, the precompression effects given by nose shape of forebody and geometrical arrangement of air intake underneath fuselage were analyzed by CFD calculation using 3-dimensional compressible Navier-Stokes equations.

Mixed-compression type axisymmetric air intakes for ATREX engine have been tested in the supersonic wind tunnel from Mach 0.5 to 4 since 1993. The throat area of the intake can be variable with a translating center spike to accomplish starting and off-design operation since the ATREX intake must work well over the wide flight Mach number up to 6. Here are presented effects of the intake design Mach number, the air bleed from a center spike and/or a cowl around the throat, an angle of attack and blunt nose of the spike on the intake performance characteristics, that is total pressure recovery and mass capture ratio. It is found that bleeding from the center spike and the cowl influences mainly on total pressure recovery and mass capture ratio respectively. The advantage of rounding properly off the spike nose is confirmed. Small center spike cone angle and/or blunt nose is sensitive to the angle of attack.