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資料番号: ARDS06125000

condition was calculated by incorporating results of a chemical equilibrium numerical code. The fuel rates causing the flow choking were successfully reproduced and matched with the limit of fuel rates observed in our tests. Inlet unstart of the engine caused the loss of the ejector-pumping effect in the diffuser system, which triggered flow choking. Choking was also promoted by the drag of the gas sampling rakes. The choking in diffuser flow and the engine unstart may be coupled, resulting in hysteresis in the diffuser breakdown, which was also experienced in our tests. As the rocket-based, combined-cycle engine is scheduled to be tested under the Mach 4 condition in 2007, the safety of the experiment is now a primary issue. The current analysis showed that the new engine would easily cause choking of diffuser because of the larger propellant supply rates and the relatively low specific impulse. Reasonable operation of the wind tunnel to control the flow choking in future tests was examined.

Ramp-compression-type inlets for the combined cycle engine were investigated in a Mach 4 wind tunnel. Geometries of the inlet models were changed to investigate their effects on the aerodynamic performances and the starting characteristics. A bent cowl improved the starting characteristic by weakening the shock from the cowl leading edge to suppress the shock-induced separation and by reducing the internal contraction ratio of the cowl duct, thus, avoiding choking at the inlet throat. With the current geometries, the inlet started with a slightly larger internal contraction ratio than that restricted by the Kantrowitz-Donaldson limit. The ratio of the height of the cowl duct of the inlet to the incoming boundary layer was found to affect the starting characteristic. The model with a higher capture ratio showed better total pressure recovery performance.

著者人数: 3名, 資料番号: AC1000035000

著者人数: 5名, 資料番号: AC1000025000

Combustion tests of a scramjet engine were conducted to obtain thrust, lift, and pitching moment essential for evaluation of the engine’s basic performance and for determination of the requirements to control the system of integrated engine and body. The engine was a side-compression type with a swept-back angle of 45 deg, and the inlet was open on the cowl side. Thrust, lift, and pitching moment were measured safely by a force measuring system under Mach 4, 6, and 8 simulated  ight conditions including engine unstart. Measurements changed with transition of combustion conditions.Thrust decreased and lift increased suddenlywhen the engine fell into unstart conditions in theM6 andM8 tests although the changewas gradual in theM4 tests. Effects of engine conŽ gurations were also investigated. Installation of a full-height strut improved combustion performance but increased drag more than thrust.

著者人数: 4名, 資料番号: AC1000023000

著者人数: 5名, 資料番号: AC1000022000

Gas sampling has been used in combustor studies and in scramjet engine testing. Because the gas sampling is based on the assumption that the gas composition is frozen in the sampling process, the critical Damkohler numbers necessary to quench reactions in the gas-sampling probes were evaluated using a reduced kinetic model. The phase plane analysis showed that reactions in probes can be extinguished if the probe Damkohler number is less than about 10. The analytical results were cone rmed by numerical calculations using full kinetics. The shock swallowing into sampling probes was examined using numerical simulations for the low-Reynolds-number e ow. These theoretical results were verie ed by experiments using four kinds of probes with various cone gurations in a Mach 2.5 supersonic combustor. Based on the results, e ne sampling probes with a tip diameter less than 0.3 mm are recommended for scramjet testing. Based on these calibration studies, gas sampling was successfully applied to scramjet engine testing under a e ight Mach number up to 8, to reveal interesting features in the internal e ow in swept-back engines.

著者人数: 6名, 資料番号: AC1000020000

Joint Propulsion Conferences, 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference（July 14-17, 2013.), San Jose, California, United States., Ground firing-tests of rocket-ramjet combined-cycle engine model with 3 m length was conducted under Mach 8 flight condition at Ramjet Engine Test Facility in Kakuda Space Center of JAXA. Numerical analysis was also conducted for the engine under Mach 8 flight condition. The effects of rocket chamber pressure and a secondary fuel injection on the engine performance such as thrust and specific impulse were examined., 資料番号: AC1400049000, レポート番号: AIAA2013-3668

An experiment was conducted to investigate the effect of injector/combustor geometry on combustion-induced peak wall pressure and associated upstream influence, as well as on mixing/combustion characteristics at an entrance Mach number of 2.5. The length of the constant area section downstream of injection orifices had a strong influence on the above-mentioned characteristics. However, the sweep of the rearward-facing steps on both side walls had little effect on these characteristics, nor did reversing them have any effect. The peak wall pressure and the length of the upstream influence agreed qualitatively with predictions of an analytical model and an empirical formula developed at Johns Hopkins University. Fuel jets injected from the model with the longest constant area section and the fuel equivalence ratio of unity, coalesced at a very early stage downstream of the fuel injection orifices. This coalescence led to a decrease in mixing rate downstream, despite the higher degree of mixing near the injection orifices. The combustion efficiencies were higher than those obtained at NASA Langley in the upstream region due to the higher mixing rate near the injection orifices.

A series of experiments was conducted for studying scramjet nozzles under Mach 8 flight conditions using a high temperature gas flow produced by combustion of monomethyl-hydrazine (MMH) and nitrogen tetroxide (NTO). The stagnation temperature was designed to be 3170K, and the stagnation pressure to be 1MPa. Scramjet nozzles with expansion ratios of 3 (EN3) and 5 (EN5) were examined. Delivered specific impulse (Isp) was found to be 17.3s (EN3) and 26.1s (EN5). Performances of nozzles were predicted using numerical codes, and energy release loss, kinetic loss and two-dimensional loss were identified. The two-dimensional loss was predicted to be 16.5% in the EN3 nozzle and to decrease to 7.0% in the EN5 nozzle. The kinetic loss due to chemical freezing was found to be about 4%. A friction loss evaluated from thrusts suggested laminar flow in the scramjet nozzle with the MMH/NTO experiments. Measured heat flux on nozzles also supported the laminar boundary layer. Experiments using a cold gas flow with the nozzles were also conducted to compare with the hot flow.

In order to get experimental data in subsonic Mach number flight region, two small hybrid rockets installed the ejector-jet were launched, and then the flight analysis was performed. Numerical simulations were conducted for clarification of the flowfields and prediction of the ejector-jet performance. The CFD results revealed the compound compressible flow phenomena involving oblique shock waves and the subsonic flows. The effects of the flight Mach numbers and the rocket pressures to the suction performances were also investigated.

A combustion-capable combined cycle engine model which was constructed based on the rocket and ramjet technology was tested in Mach 4 flight condition. At this speed, engine is designed to shift its operation mode from an ejector-jet to a ramjet. Both modes were simulated by changing the rocket combustion pressure. Even with full rocket exhaust, no effect to the air flow could be observed. The injection point of the secondary fuel affected thrust performance. In the ramjet mode, the pressure rise due to the fuel combustion traveled to the entrance of the combustor, but it stayed near the injection point in the ejector-jet mode.

This report presents recent research activities of the Combined Propulsion Research Group of Japan Aerospace Exploration Agency. Aerodynamics and combustion of the scramjet engine and the rocket–ramjet combined-cycle engine, structure and material for the two engines and thermo-aerodynamic of a re-entry vehicle are major subjects. In Mach 6 condition tests, a scramjet engine model produced about 2000 N net thrust, whereas a model produced thrust almost equal to its drag in Mach 12 condition. A flight test of a combustor model was conducted with Hyshot-IV. A rocket–ramjet combined-cycle engine model is under construction with validation of the rocket engine component. Studies of combustor models and aerodynamic component models were conducted for demonstration of the engine operation and improvement of its performances. Light-weight cooling panel by electrochemical etching examined and C / C composite structure was tested. Thermo-aerodynamics of re-entry vehicle was investigated and oxygen molecular density was measured also in high enthalpy flow.

44th AIAA Aerospace Sciences Meeting and Exhibit(9 - 12 January 2006), Reno, Nevada, 資料番号: ARDS05397000, レポート番号: AIAA 2006-1027

Flowfields in a scramjet engine inlet were numerically simulated by the Navier-Stokes equations with a turbulence model, and aerodynamic effects of the strut position in the scramjet engine inlet were also investigated. Because the RJTF experiments reveal that characteristics of scramjet engine performances are quite dependent on the strut positions. To investigate the RJTF experimental results, the computations were conducted for inflow Mach number of 5.3 on NAL’s Numerical Space Engine. The correlations were obtained between the inlet flowfields and the struts position, and the effects of the RJTF nozzle flow boundary layer were clarified.

The combustion performance of scramjet engines was investigated by gas sampling in Mach 4, 6 and 8 flight conditions. Since gas sampling is based on the assumption that the gas composition is preserved in the sampling process, the critical Damkohler numbers nec- essary to quench reactions in the gas samling probes were evaluated using a reduced kinetic model. The analysis using the phase plane showed that reactions in probes can be extin- guished if the probe Damkohler number is less than about 10. The analytical results were duplicated by numerical calculations using full kinetics. The shock swallowing in sampling probes was examined using numerical simulations for the low-Re flow. These theoretical results were verified by experiments using four kinds of probes with various configurations in a M2.5 supersonic combustor. Based on the results, fine sampling probes with a tip dia- meter less than 0.3 mm are recommended for scramjet testing. The experiments also showed that pitot pressure is more sensitive than wall pressure to combustion in supersonic flow. This implies that the wall pressure distribution may underestimate the combustion ef- ficiency in supersonic combustion ducts.

AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 34th, Cleveland, OH, July 13-15, 1998, 資料番号: AC1000105000, レポート番号: AIAA-1998-3134

AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 33rd, Seattle, WA, July 6-9, 1997, 資料番号: AC1000104000, レポート番号: AIAA-1997-3018

ASME, SAE, and ASEE, Joint Propulsion Conference, 25th, Monterey, CA, July 10-13, 1989, 資料番号: AC1000089000, レポート番号: AIAA-1989-2680