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.
We conducted a sensitivity analysis of engine performance on measurements of gas sampling, pitot pressure and static pressure. It indicates that the heat loss in engine and the insufficient quenching of reactions in the gas sampling probes results in the mismatch in the mass flux through the engines. An analytical study of the reactive flow in probes showed that the introduction of a small heat of reaction with radical recombination to the energy equation yielded a saddle-type singularity in a phase plane for temperature and radical. The separatrix running from the initial point to the saddle point gave the criterion dividing the thermal runaway and the chemical quenching. A one-dimensional reaction code with a full-kinetics confirmed this analytical results.
Quenching of reactions in probes is the prerequirement in measurements of combustion performance of engines by the gas sampling. In order to quantify degrees of the quenching in the probes used for scramjet testing, we tested four kinds of gas-sampling probes, three freezing-oriented and a reaction-oriented probes using a supersonic combustor operating with an M 2.5 air flow of the temperature up to 2250K and compared the combustion efficiency (ηc) measured. The freezingoriented probes yielded gas compositions consisting of partially burned H2 and O2 even at an air temperature of 2200K. The pitot pressure indicated that the combustion took place in the supersonic combustor not in the probes. The freezing-oriented probes with the tip diameter of 0.3mm indicated the correct ηc in the scramjet combustor.
A large antenna reflector with a continuous mesh surface has the advantage of higher electrical gain and lower side-lobe level characteristics. We designed a hybrid cable network for a continuous and highly accurate mesh surface to realize a surface accuracy of 1.0mmRMS for a 10m class antenna reflector. This cable network consists of both a triangle-shaped network and a cobweb-shaped network. One segment model with a diameter of 2.8m was manufactured to confirm the surface adjustment capabilities. A partial model consisting of seven segments and with a diameter of 7m was also manufactured to verify the surface accuracy and the interface conditions between the mesh surface and the truss structure. The surface accuracy was confirmed to within 0.96mmRMS after three courses of adjustments and measurements. The surface accuracy repeatability was confirmed to within 1.0mmRMS. The new design concept of the highly accurate mesh surface was also verified by those experiments.
A Genetic Algorithm (GA) has been applied to optimize a wing aerodynamic shape for generic subsonic transportation aircraft. A three-dimensional compressible Navier-Stokes (N-S) solver is used to evaluate aerodynamic performance of a wing. Although a GA has proved its ability to solve hard optimization problems, a GA coupled with a three-dimensional N-S solver has not been done yet because it needs large computational time even on the latest supercomputer. To overcome this difficulty, Numerical Wind Tunnel at National Aerospace Laboratory, a parallel vector machine with 166 processing elements, was used. In the optimum design obtained from the present GA, the design principles for the wing developed by existing theory and experience are found to be realized. This indicates feasibility of the present approach for the aerodynamic optimization in advanced computational environment.
Various test methods have previously been adopted to test large deployable space structures. However, traditional methods are not applicable to a precise evaluation of three-dimensional deployment characteristics with quasistatic deployment motions. This paper describes the design and performance of a deployment test system using magnetically suspended sliders. The sliders support the deployable structure vertically, and each of them moves freely with negligible friction under the horizontal surface of a ceiling plate. All the sliders control the tension of their vertical cables independently using the feedback signals of the tension sensor to compensate for the gravity of the deployable structure. The test system using magnetically suspended sliders enables maximum drag with a horizontal motion of less than 0.25N and tension control precision from a suspension cable of within 0.25N. The adequacy of this test system is confirmed by a deployment test using a 7m test model.
Ceramic thermal barrier coating systems will be of growing importance for the reusable high performance engine. These are very attractive ways to achieve a high performance engine because of their potential for reducing the film-cooling requirements in the rocket chamber. However, due to the large difference in thermal coefficient of expansion between the coating materials and the metal wall and the low ductility of the ceramic coating, cracks occur in the ceramic coating layer or spalling occurs during repeated thermal cycles. One method of improving adhesion of the coating to the metal wall is to apply functionally graded materials (FGM). In this test series, high altitude performance tests (HAPT) of a regeneratively cooled 1, 200N thrust engine composed of ZrO2-8%Y2O3(PSZ)/NiCr FGM chamber were conducted with nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) bipropellant. To enhance the engine performance, we employed the high performance unlike quadlet element injector. The film cooling fraction was reduced to 0% of the total fuel flow rate to obtain high performance. The combustion chamber used in HAPT was composed of perfect PSZ/NiCr FGM, i. e., the chamber inner wall was made of PSZ, the cooling wall side was made of NiCr and intermediate materials were PSZ/NiCr FGM. A total of 50 firing tests including high altitude performance tests and sea level tests were performed to evaluate the engine performance in terms of vacuum specific impulse (ISPV), and also to obtain chamber thermal data. The high performance of the engine, i. e., ISPV=318s at PC=1.4MPa, was verified.
National Space Development Agency of Japan (NASDA) started to develop H-II launch vehicle in 1983 and succeeded in the 1st flight on February in 1994. Meanwhile, several troubles were experienced in the development of LE-7 which is the 1st stage engine of H-II. The LE-7 is a staged combustion cycle engine with high pressure and high Isp like SSME. Therefore, high delivery pressure and high performance were required for LH2 turbopump, which is the main key component, and caused several troubles in the development. This paper describes the development of LH2 turbopump containing the design description, the contents of some troubles and their counterplans. In addition, some activities to solve turbopump life and endurance will be introduced.