Akinaga Kumakawa

LE-X -Japanese Next Liquid Booster Engine-

The LE-X engine is under study for Japan’s next flagship expendable launcher (post H2A) to be operated in the next decade with enhanced reliability and reduced cost. The goal of LE-X development is to meet the requirements from the vehicle for higher reliability, lower production cost and appropriate performance. Technology development itself is also a purpose of this investigation and will be applied to other forthcoming engines to be developed in Japan. The early-stage feasibility study of the LE-X engine was completed in 2005 through primary studies on system design, engine component design, cost reduction, reliability prediction, subscale testing, and computational simulation. In 2006, engine system analysis and fundamental studies on LE-X components by means of element tests were successfully conducted. In 2007, we have optimized the engine baseline configuration from aspect of cost reduction activities. Significant cost reduction will be achieved by drastic simplification of the engine system, and the innovation of the manufacturing process. Technology development will be ongoingly conducted to mitigate development risks, such as precise life prediction analysis of combustion chamber, prediction of combustion instability, and high-fidelity simulation.

High-Frequency Flame Oscillation Observed at a Coaxial LOX/LH2 Injector Element

To study the mechanism of the initiation of combustion instability as hydrogen injection temperature decreases, a hydrogen temperature ramping test was conducted with a single coaxial injection element with LOX/LH2 at a chamber pressure of 8.0 MPa. Two types of injectors were used in the tests. One employed a straight bore LOX post and the other employed a taper-reamed LOX post for better atomization of LOX. The combustion flame was visualized with a high-speed video camera at a rate of 6,000 frames per second. Results showed that unstable combustion was initiated when the hydrogen injection temperature decreased to less than a certain cryogenic temperature. By observing the movement of the prominent pattern of OH emission on the flame, the flame was found to propagate downstream at a constant speed with the flame angle remaining constant during stable combustion. On the other hand, injection pressure peaks appeared during unstable combustion. In this case, a block of flame with strong OH emission was occasionally observed. A block of flame caught up with an anterior block and coalesced into a large block with strong OH emission. This coalesced block of the flame is herein termed “flame burst”.

Effect of Liquid Disintegration on Flow Instability in a Recessed Region of a Shear Coaxial Injector

A liquid jet disintegration phenomenon in the recessed region of a shear coaxial injector was examined to find the relation between the difference of the recess depth and flow instability which may induce combustion instability of liquid rocket engines. An injector with a rectangular cross section and a recess for a central post, which modeled the shear coaxial injector element employed in liquid rocket engines, was constructed. The injector was made of transparent acrylic glass to allow observation of the disintegration phenomenon in the recess. Cold-flow tests with water and nitrogen gas with ambient pressures of 0.2, 0.3 and 0.4 MPa were conducted. Results showed that a condition arose in which the flow in the recess was choked by two-phase flow. The choked flow was accompanied by vibration of the central post which caused a significant change of the disintegration pattern from moderate disintegration to violent disintegration. A similar transition from a fiber-type flow to a super-pulsating disintegration flow reported by Chigier and Reitz was also observed under non-choked conditions for a coaxial injector without a recessed region. The boundary of the transition was found to depend on certain values of ReL/(WeG) 0.5 for each recess depth, including the two-phase choked flow condition. This means that the transition from a fibertype flow to super-pulsating disintegration leads to the transition from a non-choked flow to a choked flow.

Hot-gas-side Heat Transfer Characteristics of a Ribbed Combustor

For expander cycle liquid rocket engines, the enhancement of the heat transfer between combustion gas and regenerative coolant is one of the key issues. The adoption of a ribbed combustor is one possible method of enhancing heat transfer. Hot firing tests of sub-scale combustors with hot gas side wall ribs were conducted to evaluate enhancement of hot gas side wall heat extraction. Based on the firing test results, the influence of combustion pressure and mixture ratio on heat transfer enhancement was evaluated as the basic heat transfer characteristics with ribs and the scale effect was also examined. Nomenclature Pc = chamber pressure MR = mixture ratio ηC* = efficiency of characteristic exhaust velocity φ = rib efficiency

Numerical Investigations of Heat Transfer Enhancement in a Thrust Chamber with Hot Gas Side Wall Ribs

Heat transfer enhancement due to hot gas side wall ribs in a thrust chamber is discussed in this paper. Three-dimensional Favre-averaged Navier-Stokes simulations conjugating heat conduction simulation were performed for LOX/GH2 calorimeter chambers considering the finite rate chemical reactions. Atomization process of a LOX jet was modeled as LOX droplets by Discrete Phase Model. The computed results reveal the threedimensional combustion flow mechanism inside the chambers, and predict heat transfer enhancement due to hot gas side wall ribs. The computed heat flux agreed well with the experimental values, although some discrepancies were observed. Maximum heat flux and temperature on the hot gas side ribbed cylinder as well as the throat section were predicted.

Life prediction of CIP formed thrust chamber

Abstract The authors previously proposed a new fabrication method of closeout for a regeneratively cooled thrust chamber, the CIP (Cold Isostatic Pressing) forming method, by which a low stiffness closeout is easily obtained. In this study, sintered aluminum alloy was chosen as a porous closeout material which had relatively a lower Youngs modulus and a lighter weight. The forming conditions of the porous closeout and its elastic and plastic behaviors were investigated. Then the fatigue life of the high pressure chamber cooled by liquid hydrogen with the sintered aluminum alloy closeout was analyzed by means of nonlinear FEM (Finite Element Method). The results showed that the optimum design condition for a long-life rocket thrust chamber could be achieved with a low-stiffness closeout consisting of a CIP formed aluminum alloy layer with a porosity of 17% and a TFE (polytetrafluoroethylene) insulation layer with a thickness of 20 mm. The chamber with this low-stiffness closeout had a prolonged fatigue life three times longer than that of a conventional chamber with an electroformed nickel closeout.

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.

Experimental and Numerical Study on Characteristics of Fuel Mixers for a Liquid Rocket Engine

Some rocket engines have a fuel mixer upstream of the injector to mix two hydrogen flows of different temperatures. In the mixing process, this fuel mixer may generate large fluctuations of flow properties, which in turn may increase combustion pressure fluctuations. Therefore, fuel mixers must be designed carefully to prevent such large fluctuations. In addition, fuel mixers must have good mixing characteristics and be free of large flow property fluctuations even at off-design points when rocket engines require deep-throttling capability. In this study, we experimentally and numerically investigated the effects of fuel mixer configuration in a rocket engine on the downstream flow properties. In the experiments, we used three different mixer models with different cryogenic hydrogen injection hole configurations (small holes, large holes, and a mixture of both size holes), and conducted experiments using cryogenic hydrogen and gaseous hydrogen under different flow conditions corresponding to engine throttling. The mixers with large injection holes showed better mixing characteristics than the mixer with smaller holes even under conditions of throttling.

ZrO2/Ni composite plating for high pressure thrust chambers

The chamber wall of a water-cooled calorimetric rocket combustor was electroplated with ZrO2 dispersed nickel. Combustion tests were conducted using liquid oxygen/gaseous hydrogen at a chamber pressure of 7 MPa. The structural integrity and performance of the composite plating as a thermal barrier were examined.