Takeshi Kanda
National Aerospace Laboratory of Japan
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43rd AIAA Aerospace Sciences Meeting and Exhibit | 2005
Kouichiro Tani; Takeshi Kanda; Shin-ichiro Tokutome
Aerodynamic performances of the combined cycle engine model was tested in a subsonic to transonic flow region. In this low speed, the engine works in an ejector jet mode. To simulate ejector eect, gas nitrogen was exhausted from the nozzle which located downstream of the inlet. Wall pressure distributions and Pitot pressure at the exit of the model were measured to grasp basic features of the engine flow field. The fluctuating Pitot pressure was also measured to estimate the mass flow oscillation. Two dimensional CFD was carried out to compare the experimental results and the flow structure inside the inlet was examined. With the current configuration, high pressure region on the ramp which caused the increment of drag, was formed and well predicted by CFD.
34th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit | 1998
Sadatake Tomioka; Takeshi Kanda; Kouichiro Tani; Tohru Mitani; Takashi Shimura; Nobuo Chinzei
To attain intensive combustion in M8 flight conditions at which a basic strutless engine failed to initiate intensive combustion, a strut for the inlet contraction ratio of 5 with fuel injectors was installed and the engine was tested in M8 flight conditions, with a total temperature of 2550 K, a total pressure of 10 MPa and an airflow Mach number of 6.7. In the small fuel flow rate regime, pressure increments in the diverging combustor were observed due to the boundary layer combustion. This pressure rise was enhanced with the auxaliry injection from the strut which enhanced the mixing of the main fuel. The intensive combustion within the constant area combustor was attained at higher fuel flow rates, which resulted in transitions to an engine unstart condition. A prediction method was adopted to estimate the limiting fuel flow rate for attainment of the intensive combustion within the constant area combustor and for the transition to the unstart. Introduction Supersonic combustion ramjet (scramjet) is expected to be the most effective propulsion system for Single Stage To Orbit (SSTO) transportation system and the hypersonic transportation system of next generation. Many studies on components of the scramjet engines such as inlet, combustor and nozzle have been carried out. On the other hand, intensive interactions between these components are expected in real engines [1]. Thus, it is necessary to conduct testing of complete engine model to investigate the interactions and the engine performances. However, only limited data on these complete engine models and their performances have been published [2]. Using a blow-down type wind tunnel facility (denoted as RamJet engine Test Facility; RJTF) at NALKRC, we have conducted tests of a scramjet engine at various flight conditions from M4 to M8 [3-8]. This engine had a sidewall compression type inlet section with a contraction ratio of 2.9. The ratio was limited to assure starting capability in the M4 flight conditions. However, this low contraction ratio resulted in low pressure levels at the combustor entrance in the M6 and M8 flight conditions, and no intensive combustion within the engine was observed at both flight conditions [3, 6]. To attain intensive combustion in M6 flight conditions, it was necessary to place a strut within the engine for further compression and low velocity region enlargement [4]. To attain intensive combustion also in M8 flight conditions, a strut for the inlet contraction ratio of 5 with fuel injectors was fabricated and installed. Present paper reports the results of the engine tests in the M8 flight conditions, with a total temperature of 2550 K, a total pressure of 10 MPa and an airflow Mach number of 6.7. Copyright
10th AIAA/NAL-NASDA-ISAS International Space Planes and Hypersonic Systems and Technologies Conference | 2001
Takeshi Kanda; Nobuo Chinzei; T Kenji Kudo; Atsuo Murakami
The dual-mode operation was studied experimentally in a Mach 2.5 wind tunnel, to which a scramjet combustor model was connected directly. The total temperature and the total pressure of the air were 2000 K and 1.0 MPa, respectively. The air was heated in a vitiation heater. The mole fraction of oxygen was 21 %. Two kinds of the ramjet-mode operation and a scramjet-mode operation were attained. In a ramjetmode operation, the air was decelerated in the straight duct section. Subsonic combustion and the acceleration around the entrance of the divergent section followed. In another ramjet-mode, the air was decelerated in the divergent section, subsonic combustion followed, and the gas was choked at the exit of the combustor. These operating modes were attained in the same combustor and at the same fuel flow rate by changing the fuel injection position. There was no significant difference in the thrust increment among the three operating modes. NOMENCLATURE
36th AIAA Aerospace Sciences Meeting and Exhibit | 1998
Masatoshi Kodera; Kazuhiro Nakahashi; Tetsuo Hiraiwa; Takeshi Kanda; Tohru Mitani
Computations of internal viscous flowfields of scramjet models were conducted at inflow Mach number of 5.4. An unstructured hybrid grid method was used to compute complex geometries such as scramjet models with a short strut. The numerical method to solve the Navier-Stokes equations on the hybrid grid was developed using a finite volume cell vertex scheme and the LU-SGS implicit time integration algorithm. The computational results using one-equation turbulence models showed good agreement with the experimental data. The flow features and the changes of flowfields due to the short strut located in the upper passage were discussed. It was revealed that a thick subsonic region did not exist in the combustor near the top wall at Mach number 5.4. It was favorable features to avoid the engine unstart. With the strut, relatively low velocity regions became larger and the down wash flow toward the cowl behind the step became strong. The overconcentration of the fuel toward the top wall during the weak combustion was found in the experiment. From the computational results, the reason of this overconcentration was realized that the airflow near the injector was turned to the top wall due to the small influence by the combustion in the experiment. The computational time and the accuracy of the present method were the same level as the conventional structured grid methods. Thus the present method seemed to be engineeringly very useful for analysis and design of the high-speed propulsion engines.
33rd Joint Propulsion Conference and Exhibit | 1997
Kouichiro Tani; Takeshi Kanda; Tetsuji Sunami; Tetsuo Hiraiwa; Sadatake Tomioka
A side wall compression type scramjet model was tested at Mach 4, 6 and 8 flight conditions. The effects of the incoming Mach number on the aerodynamic performance were investigated with a simple model, and it was found that the compression ratio was not high enough in M6 and M8 cases. Additional structures like a strut were applied to the model to enhance the pressure ratio. The results showed that it would achieve higher pressure in some regions, but always associated with the local pressure decrease. The aerodynamic coefficients study of every configuration tested here showed the modifications with these structures would increase the drag coefficient by the factor of up to 1.3. The modification should be made with a proper configuration , at proper location, considering with the combustion requirement.
38th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit | 2002
Tetsuo Hiraiwa; Takeshi Kanda; Tohru Mitani; Yoshinari Enomoto
National Aerospace Laboratory of Japan (NAL) has been testing sidewall-compression type scramjet engines under Mach 8 flight condition since 1995. Over 100 tests under the condition have been conducted at NALs RamJet- engine Test Facility (RJTF). However, no significant achievement has been taken from these experiments. Several type/configuration of engines were delivered and tested, but they could not produce thrust larger than their internal drag. This report presents recent modification to one of the engines and its experimental achievements: To reduce the drag caused by a strut, which has been generally used as a flame holder and a part of air compressing system, a ramp-compression wall (ramp block) is installed on the topwall instead of the strut. Then, this engines fundamental performance and characteristics are displayed and discussed. Additionally, we show some experimental results of heat flux on the ramp. Comparing these results with our former experimental results, we discuss advantages of the ramp block model.
44th AIAA Aerospace Sciences Meeting and Exhibit | 2006
Kouichiro Tani; Takeshi Kanda; Shin-ichiro Tokudome
Modifled combined cycle engine models were tested in subsonic and transonic regions to improve the aerodynamic performance and overcome the problems which were found in the previous tests with the ejector jet prototype model. Three inlet geometries and two cowl options were tested and all models successfully reduced the inlet drag comparing to that of the prototype. In a high subsonic or transonic ∞ow, the current models showed the possibility of the thrust augmentation by the ejector-jet cycle. One dimensional aerodynamic estimations were also carried out and the results were compared with the data evaluated from the pitot pressure measurements. The mass ∞ow was in good agreement and, with a slight adjustment of the parameter, the total pressure prediction method showed reasonable match with the experimental data.
43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit | 2007
Sadatake Tomioka; Tetsuo Hiraiwa; Shuuichi Ueda; Kouichiro Tani; Noboru Sakuranakai; Takeshi Kanda
A rocket-ramjet combined cycle engine model, embedding twin rocket chamber on top wall side of a scramjet flow pass, was fabricated and was tested under sea-level, static conditions. The rocket chamber was driven with gaseous hydrogen and oxygen at nominal operation condition of 3 MPa in chamber pressure and 7.5 in mixture ratio. Gaseous hydrogen was also injected through secondary injector orifices to pressurize the ramjet combustor. Contraction in the inlet section was changed to investigate air ingestion performance, while a mechanical contraction mechanism with variable contraction ratio was installed near the exit of the engine to enhance pressure recovery within the diverging portion of the flow pass. Choking of the ingested airflow was not attained, and the airflow rate was 2/3 of the design value. Pressure-rise within the diverging portion of the flow pass was not intensive to attain choked condition at the engine exit. The mechanical contraction enhanced penetration of shock system associated with thermal and mechanical throttling at the engine exit, which in turn, enhanced mixing between the airflow and the rocket plume. However, the shock system penetrated further upstream on the airflow side with lesser total pressure, increased the back ‐pressure to the inlet section, resulting in reduced airflow rate. An excess injection of secondary fuel generated a shock system-like pressure rise, however, enhanced mixing was not sufficient to sustain pressure-rise through secondary combustion within the ramjet combustor with the excessive injection turned-off. Thus, enhancement of mixing between the airflow and rocket plume without help of the shock system is necessary for further thrust augmentation.
AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies Conference | 2005
Kanenori Kato; Takeshi Kanda; Kenji Kudo; Atsuo Murakami
A rocket based combined cycle engine was tested in scramjet-mode. The combustor model had two rockets in the combustor section. These rockets were used not only for thrust production unit, but also for fuel supply unit. The test facility consisted of an air supply section, a combustor section, a straight duct section, and a divergent duct section. Total pressure of the supplied air was 1 MPa, and the total temperature of the air was 2400 K. In the tests, in order examine the effect of the length of the straight duct section, four kinds of length of straight duct were connected downstream to the rockets. Two combustion patterns were observed for different lengths of straight duct. Pitot pressure and gas sampling were measured on the exit plane of the combustor model. Better combustion conditions were attained in the case of a longer duct. Nomenclature A = cross section Fc = impulse function of combustion gas L = length of straight duct section (including the straight part of combustor downstream of rocket) Pw = wall pressure Pwi = wall pressure at the entrance of the combustor W = width of the straight duct section x = distance from the position of the rocket nozzle exit along the airflow direction y = vertical distance from the bottom z = transverse distance from sidewall surface (rocket side) (O/F)r = mass flow ratio of oxygen to fuel in the rocket φ = equivalence ratio to airflow ηc = combustion efficiency
12th AIAA International Space Planes and Hypersonic Systems and Technologies | 2003
Tohru Mitani; Sadatake Tomioka; Takeshi Kanda; Nobuo Chinzei; Toshinori Kouchi
Meeting Information: 12th AIAA International Space Planes and Hypersonic Systems and Technologies, Norfolk, Virginia, Dec. 15-19, 2003