Tadashi Masuoka

Experimental Study on Transitional Phenomena of Extendible Nozzle

An extendible nozzle is considered to be a feasible device to improve the performance of booster engines because it has possibilities to provide altitude compensation and to achieve higher specific impulse. The booster engine with the extendible nozzle has to deploy its nozzle extension on engine firing. For the design of the extendible nozzle including its driving mechanics, it is required to clarify the transitional phenomena during the nozzle deployment. In order to investigate the transitional phenomena, firing tests using a sub-scale model were carried out on a high altitude test stand. In the firing test, the nozzle extension was sustained at a fixed position by supporting rods and the ambient pressure was varied to simulate an altitude change. In order to simulate the nozzle deployment, also the position of nozzle extension was adjusted by replacing the supporting rods. The nozzle extension was sustained at 57%, 67%, 77%, 87%, and 100% deployed positions. The axial load and side load acting on the nozzle extension were evaluated quantitatively by using small load cells installed in each supporting rod, and in addition, the distributions of wall pressure and heat flux along the inner wall of the nozzle extension were measured. Based on the experimental data, the transitional phenomena were investigated in both the mechanical and the thermal aspects. As results, the mechanical and the thermal conditions which can be critical in the design of the extendible nozzle were provided.

Basic Design and Fundamental Data Acquisition Test of Reentry Melting Promotion Type Satellite's Propellant Tank

Because the titanium tank is one of the greatest ground-damaging component when a low-Earth orbit satellite falls to ground, we plan to replace the general titanium propellant tank with a reentry melting promotion-type tank to reduce the ground hazard. We selected an aluminum-lined, carbon composite overwrapped tank with a lower melting temperature in a feasibility study and conducted fundamental tests including a liner material aluminum compatibility test with hydrazine propellant and an arc heating test. We also established a target specification for the tank and performed the basic design, trial manufacture, scale model PMD (Propellant Management Device) test under low gravity, CFRP outgassing property acquisition and CFRP resistance to radiation tests. This paper describes how the basic design and fundamental data acquisition test of the tank finished well. I. Introduction hen the satellite NROL-21 fell to Earth in 2008. the fall point contamination from the hydrazine propellant became problematic, while the falling rockets also caused considerable pollution. Although general tanks made from titanium alloy are used for satellite propulsion systems to save weight and ensure propellant compatibility, titanium alloy has a high melting point and is durable on reentry. However, because the titanium tank is one of the components most likely to damage the ground when a satellite in low-Earth orbit falls to ground, we plan to replace the general titanium propellant tank with one which will melt on reentry to reduce the ground hazard. We selected an aluminum-lined, carbon composite overwrapped tank with lower melting temperature in a feasibility study and conducted fundamental tests, including a liner material aluminum compatibility test with hydrazine propellant and an arc heating test. We also established a target specification for the tank, which included the mass, volume, MEOP (Maximum Expected Operating Pressure), propellant expulsion efficiency, propellant storage life, low cost, short manufacturing lead time etc. and performed the basic design, trial manufacture of key parts, scale model PMD (Propellant Management Device) test under low gravity, CFRP outgassing property acquisition test and CFRP resistance to radiation test. We set the volume 0.35m 3 at first. Because the volume was changed from 0.35m 3 to 0.43m 3 in 2013, we performed the basic design for 0.43m 3 volume.

Computer simulation of ultrasonic testing for aerospace vehicle

Non-destructive testing techniques are developed to secure reliability of aerospace vehicles used repetitively. In the case of cracks caused by thermal stress on walls in combustion chambers of liquid-fuel rockets, it is examined by ultrasonic waves visualization technique developed in AIST. The technique is composed with non-contact ultrasonic generation by pulsed-laser scanning, piezoelectric transducer for the ultrasonic detection, and image reconstruction processing. It enables detection of defects by visualization of ultrasonic waves scattered by the defects. In NIMS, the condition of the detection by the visualization is investigated using computer simulation for ultrasonic propagation that has capability of fast 3-D calculation. The simulation technique is based on finite-difference method and two-step elastic wave equations. It is reported about the investigation by the calculation, and shows availability of the simulation for the ultrasonic testing technique of the wall cracks.

Development of New Composite Propellant Tank for Satellites

Satellite components falling to the ground are a crucial problem, particularly titanium tanks in propulsion systems. Satellite tanks are generally made of titanium alloy because it is light and compatible with propellants used for the satellite. However, it also has a high melting point, which hinders melting on re-entry. To replace the titanium tanks, JAXA has studied as the tank for ground damage prevention on re-entry since 2010 and has started developing a new composite propellant tank since 2013. The tank is targeted; not only due to its demise properties but also its extended propellant storage, cost-effectiveness and a short delivery time. For the first development step, the minimum tank (DM: Development Model #1) which has the shortest cylindrical segment length and a volume of 250L were manufactured to confirm the strength proof of critical portions and the adequacy of production techniques. There are also plans to manufacture a PMD (Propellant Management Device) separately to the tank shell. This paper describes the design, trial productions, and data acquisition tests used when developing the new tank.

Hydrazine decomposition phenomena observed by Neutron radiography at a catalyst bed

Most mono-propellant thruster technologies were developed in the 1960s and the basic principles and fundamental structures, such as the catalyst and propellant, have remained in use without major technical innovation. Conversely, much remains to be identified in terms of the concrete mechanisms and quantitative limitations of the phenomena inside the monopropellant thruster. One of our studies to improve the reliability of propulsion systems involved visualization, facilitating direct observation of the physical and chemical phenomena occurring within the catalyst bed of the mono-propellant thruster. In this paper, we introduce the visualization test results of the mono-propellant thruster utilizing Neutron Radiography. We have also succeeded in shooting movies of hydrazine decomposition phenomena using Neutron Radiography at the Kyoto University Research Reactor.

Feasibility study on a neutron diffraction method for non-destructive evaluation of residual strain distribution of a combustion chamber

A regeneratively cooled combustion chamber of a cryogenic liquid rocket engine is exposed to a large temperature difference between a combustion gas and a liquid hydrogen coolant. The large temperature difference induces large thermal stress in the chamber wall, and the strain is accumulated throughout the cyclic firing tests. Evaluation of the stress and the strain distribution in a chamber wall is essential for chamber life prediction because the chamber life is related to such stress and strain. A feasibility study on a neutron diffraction method for evaluation of residual strain distribution in a chamber wall was conducted. A method of positioning cooling channels inside a combustion chamber was established. The measurement time was adjusted to improve the fluctuation of the diffraction data. The measured data were compared with the numerical data by finite element analysis.

Effects of Deep Throttling on Rocket Engine Systems

Dependencies of injection pressure, injection velocity ratio, pump flow rate, and coolant physical condition at a cooling jacket on the engine thrust level were investigated quantitatively for typical rocket engine cycles, an expander-bleed cycle and a staged combustion cycle using a dynamic simulator for a rocket engine. For low injection pressure of oxygen in the expander-bleed cycle engine during deep throttling, the effect of gasification of the oxygen using a heat exchanger was examined numerically. We found that placement of the heat exchanger upstream of the oxygen injector would be effective to increase the injection pressure difference, even though some oscillation would occur.

Experimental and Numerical Study on Performance of Extendible Nozzle for Altitude Compensation

An extendible nozzle for altitude compensation is considered to be a feasible device to improve the performance of booster engines because it can provide higher thrust at sea level and higher specific impulse in vacuum. The booster engine with the extendible nozzle has to deploy its nozzle extension on engine firing. For the design of the extendible nozzle including its driving mechanics, it is required to clarify the transitional phenomena during the nozzle deployment. Based on the experimental results of the firing tests using a sub-scale model, the characteristics of thrust coefficient, the nozzle axial load, and the nozzle side load affected by the nozzle flow transition are examined. The backflow of combustion gas through the gap between the fixed nozzle and the nozzle extension was also examined with use of CFD analysis. As results, disadvantages in the nozzle performance and the nozzle loads in case of improper nozzle deployment condition are clarified.

Wall Thickness Measurements by Ultrasonic Probes in a Liquid Rocket Combustion Chamber

The regenerative cooling combustion chamber of a liquid rocket engine is exposed to a large temperature difference between the combustion gas and the liquid fuel. To estimate the inner wall damage, an evaluation of wall deformations is very important since chamber life is usually related to such deformations. In this study, the ultrasonic thickness measurement technique based on the pulse reflection method was chosen and conducted the measuring tests of ligament thicknesses and the analyses of ultrasonic wave propagations. Two types probe were manufactured which have a polystyrene and a local immersion delay. From the test and the analysis results with a polystyrene delay, it was shown that the b1 echo could not stably obtained in an A-scan graph and measurement accuracy was not sufficient in measuring the thicknesses of a convex and a concave ligaments. From the test results with a local immersion delay, the measurement accuracy within ±50 μm was attained to measure the thicknesses of a convex and a concave ligaments. Nomenclature S1 = First surface echo B1 = First bottom echo B2 = Second bottom echo Vl = Acoustic velocity of longitudinal wave Vs = Acoustic velocity of shear wave

Evaluation of Spark Plasma Sintering (SPS) Forming Method for Liquid Rocket Combustion Chambers

The objective of this study was to investigate new fabrication methods for combustion chambers to potentially reduce fabrication time and cost. For this purpose, the SPS (Spark Plasma Sintering) forming method was applied for bonding the inner cylinder and the outer jacket of combustion chambers. With the SPS forming method, very complicated bonding can be easily achieved while sustaining sufficient bonding strength between the inner cylinder and the outer jacket and providing perfect sealing of coolant channels. The primary task of the present study was to investigate the applicability of SPS bonding to the inner cylinder and the outer jacket in combustion chambers. To confirm the applicability of SPS bonding to combustion chambers, tensile test specimens and simulated combustion chamber specimens were manufactured. For these types of specimens, pressure proof tests using water were conducted to confirm the bonding strength at coolant channel pressure of up to 50 MPa. Furthermore, ultrasonic inspections were conducted before and after the pressure proof tests using an ultrasonic imaging device. From the test results, the applicability of SPS bonding for the inner cylinder and the outer jacket in combustion chambers was confirmed.